aviation electrician's mate 3 & 2 - navy bmr material/navedtra 14009.pdf · aviation...

NONRESIDENT

TRAINING COURSE

Aviation Electrician's Mate 3 & 2

NAVEDTRA 14009

DISTRIBUTION STATEMENT A: Approved for public release; distribution is unlimited.

PREFACE About this course: This is a self-study course. By studying this course, you can improve your professional/military knowledge, as well as prepare for the Navywide advancement-in-rate examination. It contains subject matter about day-to-day occupational knowledge and skill requirements and includes text, tables, and illustrations to help you understand the information. An additional important feature of this course is its references to useful information to be found in other publications. The well-prepared Sailor will take the time to look up the additional information. Any errata for this course can be found at https://www.advancement.cnet.navy.mil under Products. History of the course: • Jan 1991: Original edition released. • Mar 2003: Minor revision released.

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TABLE OF CONTENTS CHAPTER PAGE

1. Basic Physics.......................................................................................................... 1-1

This chapter has been deleted. For information on Basic Physics, refer to Nonresident Training Course (NRTC) Aviation Electricity and Electronics—Maintenance Fundamentals, NAVEDTRA 14318, chapter 3.

2. Electrical Maintenance and Troubleshooting......................................................... 2-1

3. Power Generation and Control Systems................................................................. 3-1

This chapter has been deleted. For information on power generation and distribution, refer to Nonresident Training Course (NRTC) Aviation Electricity and Electronics—Power Generation and Distribution, NAVEDTRA 14323.

4. Aircraft Electrical Systems .................................................................................... 4-1

5. Aircraft Power Plant Electrical Systems ................................................................ 5-1

6. Aircraft Instruments ............................................................................................... 6-1

7. Compass and Inertial Navigation Systems............................................................. 7-1

8. Automatic Flight Control and Stabilization Systems ............................................. 8-1

APPENDIX

I. Glossary ................................................................................................................. AI-1

II. Symbols.................................................................................................................. AII-1

III. References .............................................................................................................. AIII-1

IV. Review Subset Answers ......................................................................................... AIV-1

CREDITS

The illustrations listed below are included through the courtesy of the designated source. Permission to use these illustrations is gratefully acknowledged. Permission to reproduce illustrations and other materials in this publication must be obtained from the source.

SOURCE FIGURES

Huntron Instruments, Inc. 2-27, 2-28, 2-29, Table 2-4, 2-30, Table 2-5, 2-31, and the related text on pages 2-46 through 2-54

CHAPTER 1

BASIC PHYSICS

This chapter has been deleted. For information on basic physics, refer to Nonresident Training Course (NRTC) Aviation Electricity and Electronics— Maintenance Fundamentals , NAVEDTRA 14318, chapter 3.

CHAPTER 2

ELECTRICAL MAINTENANCE ANDTROUBLESHOOTING

Aircraft maintenance falls into two broadcategories—scheduled maintenance and unsched-uled maintenance. Scheduled maintenance is theactions taken to reduce or eliminate failureand prolong the useful life of the equipment.Unscheduled maintenance is the actions takenwhen a system part or component has failed, andthe equipment is out of service.

In maintenance work of any kind, you use twofundamentals—knowledge and skills. As an AE,you must have specific information about theparticular equipment you repair or keep in goodcondition. Also, you must have certain generalskills and knowledge that apply to many kinds ofequipment and types of work assignments.

The specific information required consists ofspecial procedures and processes, and detailedstep-by-step directions approved by properauthority. You can find this information inpublications or checklists authorized by the NavalAir Systems Command, type commanders, andother authorities.

General maintenance skills and procedures arebased on knowledge that is not in equipmentmanuals. These skills come from schools, on-the-job training, and from training manuals such asthis one.

SAFETY

Learning Objective: In addition to generalsafety precautions, identify safetyprecautions regarding aircraft, personnel,material, and tools.

In the performance of your duties, you comeacross many potentially dangerous conditions andsituations. You install, maintain, and repairelectrical and electronic equipment in confinedspaces where high voltages are present. Amongthe hazards of this work are injury caused byelectric shock, electrical fires, and harmfulgases. Also, you must include improper use of

tools among these hazards. Common sense andcarefully following established rules will producean accident-free naval career.

When working, there is one rule to stressstrongly—SAFETY FIRST. Whether you areworking in the shop, on the line, or during aflight, you should follow prescribed safetyprocedures. When working on or near aircraft,there is the danger of jet blast, losing yourbalance, or being struck by propeller or rotorblades. Because of these dangers, you need todevelop safe and intelligent work habits. Youshould become a safety specialist, trained inrecognizing and correcting dangerous conditionsand unsafe acts.

You must know how to treat burns and howto give artificial ventilation (respiration) topersons suffering from electric shock. In somecases, you may have to perform external heartcompression in addition to artificial ventilationto restore the heartbeat. Artificial ventilation andexternal heart compression performed together areknown as cardiopulmonary resuscitation (CPR).

You should study the CPR training section ofBasic Military Requirements, NAVEDTRA14325. This text provides valuable information;however, it does not replace training in CPRprocedures. It is your responsibility to get firstaid training, including CPR. Personal CPRtraining is available at many Navy medicalfacilities. Many local fire stations and Red Crossagencies also offer this type of training. It isimportant for you to be currently certified in thespecial skills of CPR. One final note, you mustupdate your CPR certification annually.

WARNING

Do not perform CPR unless you have hadthe proper training.

The life of a shipmate could easily dependupon your CPR skills. This statement is not to

say that knowledge of other first aid proceduresis less important.

Cooperation of personnel and being vigilantprevents most accidents that occur in noncombatoperations. You can learn about general aviationsafety by studying Aviation Fundamentals,NAVEDTRA 14318. You should study andreview this manual.

GENERAL PRECAUTIONS

Only authorized personnel can repair andmaintain electronic and electrical equipmentbecause of the chance of injury, the danger of fire,and possible material damage. When you workon electrical equipment, open and tag the mainsupply switches or cutout switches. The tag shouldread as follows: “This circuit is open for repairsand shall not be closed except by direct order of

(usually the person directly incharge of the repairs).”

Securely cover fuse boxes and junction boxes,except when you are working on them. Don’t alteror disconnect safety devices, such as interlocks,overload relays, and fuses except when replacingthem. Never change or modify safety or protectivedevices in any way without proper authorization.

Remove and replace fuses only after the circuitis de-energized. If a fuse blows, replace it witha fuse of the same current rating only. Whenpossible, carefully check the circuit before makingthe replacement since the burned-out fuse oftenresults from a circuit fault.

You should move slowly when workingaround electrical equipment. Maintain goodbalance and DO NOT lunge after falling tools.DO NOT work on electrical equipment if you arementally or physically exhausted. DO NOT touchenergized electrical equipment when standing onmetal, damp, or other well grounded surfaces. DONOT handle energized electrical equipment whenyou are wet or perspiring heavily. DO NOTTAKE UNNECESSARY RISKS.

Some general safety precautions that youshould follow are shown below:

REPORT ANY UNSAFE CONDITION,or any equipment or material that youconsider to be unsafe, to the immediatesupervisor.

WARN OTHERS you believe to beendangered because of known hazards orwho fail to follow safety precautions.

WEAR or USE APPROVED PROTEC-TIVE CLOTHING or EQUIPMENT forthe safe performance of work or duty.

REPORT to the supervisors ANYINJURY or evidence of impaired healthoccurring during work or duty.

EXERCISE REASONABLE CAUTIONduring any unforeseen hazardous occur-rence, as is appropriate to the situation.

The safety precautions that apply to the workof Aviation Electrician’s Mates include those youshould follow when working in and aroundaircraft. You will also work in the electrical orbattery shop. In addition to these, you need toknow the authorized fire-fighting methods forelectrical fires, treatment of burns, and artificialrespiration.

Some of the publications you should befamiliar with are listed below:

1. Airman, NAVEDTRA 140142. Navy Occupational Safety and Health

(NAVOSH), OPNAVINST 5100.23 (series)3. Navy Safety Precautions for Forces Afloat,

OPNAVINST 5100.19

These publications contain a variety ofoperations and functions in the Navy; therefore,they are basic and general in nature. Activities usethese instructions as a basis for establishingspecific safety instructions for their particularequipment, weapons system, or locality.

PRECAUTIONS REGARDINGAIRCRAFT

As an AE, you are exposed to flight linehazards. You will be working around movingmachinery, which is dangerous. Therefore, youneed to be alert when working around aircraft.Always follow your activity’s instructions on theapplication of external power. Make sure you geta thorough safety indoctrination from yoursupervisor.

The maintenance instructions manual (MIM)for each type of aircraft has an illustration, suchas that shown in figure 2-1 for the F/A-18 aircraft.Study the illustrations for each aircraft in youroperating area. Most safety instructions requirethe anticollision light to be operating wheneverthe engine or engines are operating. This gives anadditional warning so you will be aware ofpropellers, rotors, or jet intakes and exhausts.

High-Voltage Precautions

Figure 2-1.-Radiation, intake, exhaust, and turbine bladefailure danger areas.

PRECAUTIONS REGARDINGPERSONNEL AND MATERIALS

If possible, you shouldn’t make repairs onenergized circuits. When repairs on operatingequipment are necessary, only experienced person-nel under the supervision of a senior AE shoulddo the work. Follow every known safety precau-tion carefully. Make sure there is enough light forgood illumination, and there is insulation fromground, using a suitable nonconducting material.Station helpers near the main switch or the circuitbreaker to de-energize the equipment immediatelyin case of emergency. While making the repair,someone qualified in first aid should be standingby in case of injury due to electrical shock.

NEVER work alone near high-voltage equip-ment. Don’t use tools and equipment containingmetal parts within 4 feet of high-voltage circuitsor any wiring having exposed surfaces. Thehandles of all metal tools, such as pliers andcutters, should have rubber insulating tape covers.You can’t use plastic or cambric sleeving orfriction tape alone for this purpose when workingwith circuits having a high potential voltage.

Before touching a capacitor, short-circuit theterminals to discharge the capacitor completely.Permanently attach grounded shorting prods toworkbenches where electrical devices receiveregular servicing.

Don’t work on any type of electrical apparatuswith wet hands or while wearing wet clothing.Don’t wear loose or flapping clothing. Don’t wearthin-soled shoes with metal plates or hobnails;wear safety shoes with nonconducting soles whenavailable. Don’t wear flammable articles, such ascelluloid cap visors.

Before you work on an electrical or electronicapparatus, remove all rings, wristwatches,bracelets, and similar metal items. Make sure thatyour clothes do not contain exposed zippers, metalbuttons, or any type of metal fastener.

Make sure warning signs and suitable guardsare posted to prevent personnel from coming intoaccidental contact with high voltages.

Low-Voltage Precautions

Most people never realize the dangers oflow-voltage electric shock. These hazards arepresent and dangerous. You need to be aware oftheir existence; you should be aware of anyvoltage greater than 15 volts.

DEGREE OF SHOCK. —The current thatmay pass through the body without causingdamage depends on the individual, and the type,path, and length of contact time. The resistanceof your body varies. For example, if the skin isdry and unbroken, body resistance will be quitehigh—on the order of 300,000 to 500,000 ohms.However, if the skin becomes moist or broken,body resistance may drop to as low as 300 ohms;a potential as low as 30 volts could cause a fatalcurrent flow. Therefore, any circuit with a highervalue potential is more dangerous. If a 60-hertz

alternating current passes through the chest cavity,it has the following effects:

• At 1 milliampere (0.001 ampere), you willfeel the shock.

• At 10 milliamperes (0.01 ampere), theshock paralyzes muscles, and a person maybe unable to release the conductor.

• At 100 milliamperes (0.1 ampere), theshock is usually fatal if it lasts for1 second or more. IT IS IMPORTANTTO REMEMBER THAT, FUNDA-MENTALLY, CURRENT, RATHERTHAN VOLTAGE, IS THE CRITE-RION OF SHOCK INTENSITY.

When a person is unconscious because ofelectrical shock, you can’t tell how much currentcaused the condition.

BEGIN ARTIFICIAL RESPIRATIONIMMEDIATELY IF BREATHING HASSTOPPED.

FIRST AID FOR ELECTRIC SHOCK. —Electric shock produces a jarring, shakingsensation. The victim usually feels like he/she justreceived a sudden blow. If the voltage andresulting current is high enough, the victim maybecome unconscious. Severe burns may appearon the skin at the place of contact. Muscularspasm may occur, causing the victim to clasp theapparatus or wire causing the shock. If thishappens the victim is unable to release it.

Use the following procedures for rescuing andcaring for shock victims:

1. Remove the victim from electrical contactat once. DO NOT ENDANGER YOURSELF.Remove the victim by throwing the switch if itis nearby, or cut the cable or wires to theapparatus, using an axe with a wooden handle.(Protect your eyes from the flash when the wiresare severed.) Also, you can use a dry stick, rope,belt, coat, blanket, or any other nonconductor ofelectricity to drag or push the victim to safety.

2. Determine whether the victim is breathing.Keep the person lying down in a comfortableposition and loosen the clothing about the neck,chest, and abdomen for easy breathing. Protectfrom exposure to cold, and watch closely.

3. Keep the victim from moving. In thiscondition, the heart is very weak. Any sudden

muscular effort or activity of the patient mayresult in heart failure.

4. Do not give stimulants or opiates. Send fora medical doctor at once, and do not leave thepatient until adequate medical care is available.

5. If the victim is not breathing, applyartificial ventilation without delay, even thoughthe patient may be lifeless. Do not stop artificialrespiration until the victim revives, or properauthority pronounces the victim is beyond help.

ELECTRICAL FIRES

The three general classes of fires are A, B, andC. Class A fires involve wood, paper, cotton andwool fabrics, rubbish, and the like. Class B firesinvolve oil, grease, gasoline and aircraft fuels,paints, and oil-soaked materials. Class C firesinvolve insulation and other combustible materialsin electrical and electronic equipment.

Electrical or electronic equipment fires arecaused by overheating, short circuits, friction(static electricity), or radio-frequency arcs.Also, equipment may ignite from exposure tonearby class A or B fires. Since class C firesinvolve electrical circuits, electrical shock is anadded hazardous condition. Whenever possible,immediately de-energize any electrical equipmentsexposed to class A or class B fires, or ignitedby such a fire. If the equipment cannot bede-energized completely, use protective measuresto guard against electrical shock. Extinguishingagents other than gases contaminate delicateinstruments, contacts, and similar electricaldevices. Carbon dioxide (CO2) is the preferredextinguishing agent for electrical fires. It does notconduct electricity, rapidly evaporates, and leaveslittle or no residue. It reduces the possibility ofelectrical shock to personnel and damage toequipment as a result of contamination.

A dry chemical extinguishing agent, composedchiefly of potassium carbonate (Purple-K), canbe used on electrical fires. It is a nonconductor,which provides protection against electrical shock.However, damage to electrical or electronic partsmay result from the use of this agent.

CAUTION

Never use a solid stream of water toextinguish electrical fires in energizedequipment.

Water usually contains minerals that make itconductive. (The conductivity of seawater is many

times greater than that of fresh water.) Puredistilled water is not a good electrical conductor,and it is suitable for emergency use on smallelectrical fires. If you must use fresh water orseawater, use a fog head or tip hose nozzle inelectrical in electronic equipment spaces. The fogis a fine diffusion or mist of water particles, withvery little conductivity. However, there is adanger of electric shock unless the equipment iscompletely de-energized. Also, condensationfrom the fog frequently damages the electricalcomponents. Be careful when using the fire hose.Pressure at the fireplug may be as much as 100PSI. An unrestrained fire hose may result inwhiplash for the person holding it.

Do not use foam on electrical fires. Foam candamage equipment and possibly pose a shockhazard to personnel. If necessary, you can usefoam on de-energized circuits. When a blanket offoam goes on a burning substance, the foamsmothers the fire, cutting off the air supply to theburning substance. As the supply of oxygendecreases, the fire dies out.

When fighting electrical fires, you should usethe following general procedures:

1. Promptly de-energize the circuit or equip-ment affected.

2. Sound an alarm according to stationregulations or ship’s fire bill. When ashore, notifythe fire department; if afloat, notify the officerof the deck. Give the fire location, and state whatis burning. If possible, report the extent of thefire.

3. Close compartment air vents or windows.4. Control or extinguish the fire using a C02

fire extinguisher.5. Avoid prolonged exposure to high con-

centrations of carbon dioxide in confined spaces.You can suffocate in confined spaces unlessyou’re using special breathing apparatus.

6. Administer artificial ventilation and oxygento a person overcome by carbon dioxide fumes,and keep the victim warm.

Even under normal conditions, fire aboard aNavy vessel at sea can kill and injure more peopleand cause more ship damage than happens inbattle. All personnel need to know the dangersof fire. You need to know the type and locationof all fire-fighting equipment and apparatus inyour immediate working and berthing spaces andthroughout the ship.

Volatile Liquids

Volatile liquids, such as insulating varnish,lacquer, turpentine, and kerosene, are dangerouswhen used near operating electrical equipmentbecause sparks from the equipment can ignite thefumes. When these liquids are used in compart-ments containing nonoperating equipment, makesure there is enough ventilation to avoid anaccumulation of fumes. Also, make sure the spaceis clear of all fumes before energizing theequipment.

Never use alcohol or gasoline for cleaning; useonly an approved cleaning agent.

Today, all aviation fuels are hydrocarbons.Handling hydrocarbon products is hazardousbecause of their flash point. Products such asgasoline, solvents, and most crude oils begin tovaporizes at or below 80°F; their flash point isreached at 80°F. Their flash point makes themthe most hazardous petroleum products to handle.Other petroleum products, such as kerosene andlubricating oils, have a flash point above 80°F,making them less hazardous.

The vapors of petroleum, gasoline, and otherpetroleum products cause drowsiness wheninhaled. Petroleum vapors in concentrations of0.1 percent can cause dizziness to the extent ofinability to walk straight after 4 minutes ofexposure. Longer exposure and/or greaterconcentrations may cause unconsciousness ordeath.

The first symptoms of exposure to toxic(poisonous) vapors are headaches, nausea, anddizziness. When working in an area where thereare possible toxic vapors, stay alert. If you geta headache, become dizzy, or become nauseous,you might be exposed to toxic vapors. You shouldleave the area and report the condition.

You recover from early symptoms quicklywhen you move to an area having fresh air. If youfind people overcome by vapors, get themimmediate medical attention. First aid consists ofthe prevention of chilling, and, if breathing hasstopped, artificial respiration.

If gasoline remains in contact with you skin,it may irritate the skin, particularly under soakedclothing or gloves. Remove clothing or shoessoaked with gasoline at once. Wash gasoline fromthe skin with soap and water. Repeated contactwith gasoline removes protective oils from theskin, causing drying, roughening, chapping, andcracking, and in some cases, infection. Whileremoving your clothes, an arc, caused by staticelectricity, can cause the fuel to ignite. For this

reason, remove fuel soaked clothes in a runningshower.

If a person swallows gasoline, give first aidimmediately. You should give the victim largeamounts of water or milk and 4 tablespoons ofvegetable oil, if available.

DO NOT INDUCE VOMITING. GETMEDICAL ATTENTION IMMEDI-ATELY.

JP-4 fuel has some of the characteristics ofgasoline. Due to different characteristics, such aslower vapor pressure and high aromatic content(compounds added to increase the performancenumber), handle this fuel very carefully.

JP-5 is a kerosene type of fuel. It has a vaporpressure close to 0 PSI. Since its tendency tovaporize is lower than the more volatile grades,the vapor-air mixture above its liquid surface istoo lean to ignite. For ignition to occur, theliquid’s surface must reach 140°F. Handle thisfuel with care.

Take precautions to prevent personnel frombreathing fumes from any fuel. Also prevent fuelfrom coming in contact with the skin, especiallyif the skin has abrasions, or sores.

Compressed Air

Compressed air, when misused, is dangerous.Injuries occur through hose or fitting failures,causing the hose to whip dangerously andpropelling fitting parts through the air. Blowingdust and small particles become an eye hazard.Air under pressure may cause internal injury oreven death by introducing an airstream into bodytissue, usually through an existing cut or scratch.Air can rupture cell tissues and cause severewounds through injection of minute foreignbodies into the skin. Impurities always exist in ashop’s air supply. Falls can result from trippingover air lines thoughtlessly left lying on the floor.

CAUTION

Compressed air is a special tool, do not useas a substitute for a brush to cleanmachines, clothing, or your person. Inthose cases where an air hose is essentialfor blowing out fixtures and jigs, wear eyeprotection and maintain air pressure belowa maximum of 30 PSI. It is also desirableto place screens around work to confinethe blown particles.

The National Safety Council has published thefollowing general safety rules for working withcompressed air:

Use only sound, strong hose with securecouplings and connections.Make sure there aren’t any sharp points onmetal hose parts.Close the control valve in portablepneumatic tools before turning on air.Turn off air at the control valve beforechanging pneumatic tools. Never kink ahose to stop the airflow.Wear suitable goggles, mask, protectiveclothing, or safety devices.Never use air to blow dust chips from workclothing or from workbenches.Never point the hose at anyone. Practicaljokes with compressed air have causedpainful deaths.When using compressed air, see that nearbyworkers are not in line of airflow.

PRECAUTIONS REGARDING TOOLS

The tools you use should conform to Navystandards as to quality and type. You should usethem only for the purposes of their design.Maintain tools in good repair, and turn in alldamaged or nonworking tools.

As an AE, you will use hand tools. By usingthem correctly, you will improve the quality ofmaintenance and reduce the chance of failures.All third and second class AEs should completeTools and Their Uses, NAVEDTRA 14256.

Carelessness is the biggest menace in any shop.A machine doesn’t inflict injury. Operatorinattention is the cause of most accidents inelectrical and electronics shops today. Remember,all moving machinery is dangerous. Its not safeto lean against any machine that is or may startto move. Don’t start a machine until you knowhow to operate it. Treat a machine with respect;there is no need to fear it. Do not start a machineuntil you fully understand its operation.

Always follow the two basic safety precautionsstated below:

1. Use the proper tool for its intendedfunction, and use it correctly.

2. Keep all tools in working order and in asafe condition.

Sharpen or replace dulled cutting tools.Protect tools from damage while in use or in

stowage. If a tool becomes worn, damaged, orbroken, report it to the work center supervisor.Always clean, inspect, and account for all toolsafter you complete a job. Return all tools to theirproper stowage place.

Nonmagnetic Tools

You can get tools made of nonmagneticmaterials through normal supply channels. Youwill use them to maintain equipment that can bedamaged if tools become magnetized. Normally,these tools are made from beryllium-copper orplastic. They aren’t as rugged as steel tools, andcan be damaged easily. Use them properly; theywill last longer and operate properly.

NOTE: Do not etch beryllium-copper tools.

When working near compasses and othercomponents containing permanent magnets, youalways use nonmagnetic tools. Magnetic-susceptible tools could become magnetizedand transfer this magnetic condition to otherequipment.

Insulated Tools

Often, safety considerations require use ofinsulated tools. You can get many types ofinsulated tools directly through supply channels.If you can’t get insulated tools through normalsupply channels, buy these tools locally, or modifyconventional tools. Put insulated sleeving on thehandles of pliers and wrenches and on the shanksof screwdrivers. Only use tools modified in thisway on low-voltage circuits only because of theinsulating material’s limitations. You can getspecial insulating handles for many common toolsyou use when working on equipment containinghigher voltages.

Power Tools

In working as an Aviation Electrician’s Mate,you will use some shop machinery, such as apower grinder or drill press. In addition to thegeneral precautions on the use of tools, there area few other precautions to follow when workingwith machinery. Some of the precautions are asfollows:

1. Never operate a machine with a guard orcover removed.

2. Never operate mechanical or poweredequipment unless you know how to operate them.

When in doubt, consult the appropriate instruc-tion or ask someone who knows.

3. Always make sure that everyone is clearbefore starting or operating mechanical equip-ment.

4. Cut off the source of power before tryingto clear jammed machinery.

5. Always keep everyone clear when hoistingheavy machinery or equipment by a chain fall.Guide the hoist with lines attached to theequipment.

6. Never plug in electric machinery withoutknowing that the source voltage is the same as thatcalled for on the nameplate of the machine.

Carefully inspect all portable power tools tobe sure they are clean, well-oiled, and in workingorder before you use them. The switches shouldoperate normally, and the cords should be cleanand free of defects. Ground the casings of allelectrically driven tools. Do not use sparkingportable electric tools in any place whereflammable vapors, gases, liquids, or exposedexplosives are present.

Check to make sure that power cords do notcome in contact with sharp objects. Don’t letcords kink. Don’t leave them where they mightbe run over. Don’t let cords contact oil, grease,hot surfaces, or chemicals. When damaged,replace power cords. When unplugging powertools from receptacles, you should grasp the plug,not the cord.

Soldering Iron

The soldering iron is a potential fire hazardand a source of burns. Always assume a solderingiron is hot. Never rest the iron anywhere but ona metal surface or rack provided for that purpose.Keep the iron holder in the open to reduce thedanger of fire from accumulated heat. Don’tshake the iron to get rid of excess solder. The hotsolder may strike someone or the equipment,causing a short circuit. Hold small soldering jobswith pliers or clamps.

When you are cleaning the iron, place thecleaning rag on a suitable surface and wipe theiron across it. Don’t hold the rag in your hand.Disconnect the iron when leaving the work area,even for a short time—the delay may be longerthan planned.

Grounding

A poor safety ground, or one with incorrectwiring, is more dangerous than no ground at all,

A poor ground is dangerous because it doesn’toffer full protection; it lulls you into a false senseof security. The incorrectly wired ground is ahazard because one of the live wires and the safetyground are transposed. When this happens theshell of the tool is hot the instant the plug connectsinto the outlet. You will get a shock unless thesafety ground connects to the grounded side ofthe line on a single-phase grounded system, or nogrounds are present on an ungrounded system.Today, a three-wire, standard, color-coded cordwith a polarized plug and a ground pin is used.

In all properly connected new tools, the greenwire is the safety ground. This wire attaches tothe tool’s metal case at one end and to thepolarized grounding pin at the other end. Itnormally carries no current, and is in use onlywhen the tool insulation fails. When this happens,the safety ground short-circuits the electricityaround the user to ground and protects thatperson from shock.

To check the grounding system resistance, youuse a low-reading ohmmeter to be certain thesafety ground is adequate. If the resistance showsgreater than 0.1 ohm, you should use a separateground strap.

Some old installations do not have receptaclesthat will accept the grounding plug. If you areworking on these, use one of the followingadaptations:

Use an adapter fitting.

Use the old type plug and bring the greenground wire out separately.

Connect an independent safety groundline.

When you use an adapter, connect the groundlead extension to a good ground. Don’t use thecenter screw that holds the cover plate on thereceptacle. Where separate safety ground leads areconnected externally, connect the ground first,and then plug in the tool. Likewise, whendisconnecting the tool, first remove the line plug,and then disconnect the safety ground. ALWAYSCONNECT THE SAFETY GROUND FIRSTAND REMOVE IT LAST.

UNSCHEDULED MAINTENANCE

Unscheduled maintenance is the performanceof a repair action without a set interval.Discrepancies found before, during, and afterflight fall into this category. Some of the duties

performed during unscheduled maintenance areshown below:

Visually check equipment for loose leads,improper connections, damaged or brokencomponents, etc., before applying power tothe equipment. This check applies particularlyto new equipment preserved or stored for longperiods and equipment exposed to the weather.

Conduct a close visual inspection onO-rings, gaskets, and other types of seals whenthe equipment under check is a pressurizedcomponent. A visual inspection often revealsdiscrepancies that you can correct at that timewith minimum labor and parts. Such dis-crepancies, if left uncorrected, might result in lossof an aircraft or a major maintenance problem.

REVIEW SUBSET NUMBER 1

What is the ONE rule you should alwaysfollow when working on electric or elec-tronic equipment?

When working on an electrical system, whatshould you do to the power supply?

Name the manual that you should refer tofor an illustration of the danger areas foryour aircraft.

What exterior light is required to be onduring engine operation?

When working on a/ancircuit, you need to have a person who isfully qualified in first aid standing by.

Before touching a capacitor, you should besure it is discharged by

Q7. When removing a victim from electricalcontact, you should not take what action?

Q8. List the three classes of fires.

Q9. What is the preferred fire extinguishingagent for electrical fires?

Q10. What are the two basic rules you shouldfollow when working with hand tools?

TROUBLESHOOTING TECHNIQUES

Learning Objective: Recognize trouble-shooting techniques, including how toanalyze, detect, and correct faults inelectrical equipment.

You will spend most of your time trouble-shooting the equipment in squadron aircraft. Youmaintain many components and systems that arecomplex and difficult to troubleshoot. However,the most difficult troubleshooting job will becomesimple if you break it down into the followingsteps:

1. Analyze the symptom.2. Detect and isolate the trouble.3. Correct the trouble and test the work.

ANALYZE

If you don’t understand a system or com-ponent, you can’t find out what’s wrong with it.The following are tools that will help you analyzewhat’s wrong with a system or component:

. Maintenance instructions manuals (MIMs)

. Schematics

l Records on the equipment

DETECT AND ISOLATE

The first thing to do when troubleshooting isto visually inspect the component. If parts are

obviously not in proper condition, correct thesefaults before going any further. The types ofthings to look for include open circuit breakers,improperly placed switches, burnt equipment,loose mountings, disconnected components, anddented equipment.

If you can’t find anything after visuallyinspecting the equipment, check the condition offuses and circuit breakers. Sometimes, a circuitprotector will open. If so, reset or replace thecircuit protector and apply power. If the protectoropens again, secure the power because there is aprobable circuit malfunction. Don’t apply poweruntil after you correct the malfunction.

At this point, use the MIM for equipment. Byusing the MIM, you will know which pin in aconnector is the power input to the connector.You can use the MIM to close the correct circuitbreakers and locate the switches to use for testing.If there is power to the component and it still doesnot work, then the fault is with the componentitself.

If a short, ground, or overload condition isnot indicated, take power readings at thecheckpoints shown in the MIMs. Common faultsthat interrupt power through a circuit are brokenwiring, loose terminal or plug connections, faultyrelays, and faulty switches. Look for theseconditions when checking points along a circuit.If there is no power reaching the component,assume that the component is good, and startchecking the power supply. Begin the check at thebus that supplies the power. If you find evidenceof a short circuit or overload, secure the power.At this time, you need to make continuity andresistance checks.

Continuity and Resistance Checks

The process of fault detection often leadsbeyond visual inspection and power checks. Youuse a voltmeter to find out if a power circuit isdelivering power to the proper place. However,the voltmeter won’t identify what is wrong. Anohmmeter is a better instrument for identifyingthe problem. With the ohmmeter, you can findopens, shorts, grounds, or incorrect resistancevalues. You use the schematic to trace circuits incomponents, part by part and wire by wire, untilyou isolate the trouble. You can use the MIM forthe equipment to find the proper resistancereadings.

At intermediate maintenance activities,bench-test installations are used for off-equipmenttesting. In this type of installation, there is a

complete and fully operational system with thecomponents grouped closely together. This letsthe troubleshooter remove any unit and replaceit with the one for testing. Bench-testingcomponents is used when the trouble is hard todetermine. It might be used when the unit isfunctioning, but does not measure up to minimumstandards. Bench tests are highly specialized testinstallations, and only specific tests can berun on specific components. Normally, AEstroubleshoot using a meter and a schematic.

Nonelectric Components

So far, you have learned about fault detectionin systems using electric power. However, agroup of systems and components do not useelectricity. This group includes such itemsas mechanical instruments, direct readinggauges, and mechanisms closely associated withelectrical equipment. Such mechanisms are switchactuators, mounting assemblies, mechanicallinkages, and any type of hardware that is partof an electrical system. You can not check thisequipment with a volt-ohmmeter, but the threebasic rules for troubleshooting remain the same—analyze, detect, and correct.

CORRECT AND TEST

Many faults may occur, and for each fault youmust perform a corrective action. There are rulesthat apply to practically all corrective actions. Insome cases, the detection of a fault involves moretime and labor than correcting it.

Study the job and think through each step.Form a plan of attack, and decide which toolsyou need. In addition to hand tools, considerequipment, such as extension lights and coolingblowers. Refer to the MIM. It lists special toolsyou need and also describes how to gain accessto a system component or area. Ask someone whohas performed that particular job before; he/shecan usually offer some very good pointers.

As a troubleshooter, you should plan how todeal with special safety considerations you mightmeet during a job. Some of these are hydraulicoil drainage from disconnected lines, dangling andunprotected power cables, high-pressure air lines,and the handling of heavy components. Drain orbleed off gas or liquid lines, and handle heavyobjects with special care.

NOTE: Only Aviation Structural Mechanics(AMEs) work on oxygen lines.

If exposed power cables are unavoidably leftloose, use some means to avoid their shorting out.You can do this by covering them with atemporary insulating shield. Place a warning signon the main power switch in the cockpit and onthe external power receptacle. Another very goodprecautionary measure, where applicable, is todisconnect the aircraft battery. In this case, placea sign in the cockpit stating the battery isdisconnected. Complete all of the safety measuresmentioned here before starting work.

The nature of the job determines the pro-cedures you will use to actually do the work. Youwill either repair or replace the component orsystem. Total job quality depends on the qualityof work on smaller components/systems, such assoldering, replacing connecting devices, safetywiring, and using tools.

Learn and practice good techniques. Workingunder the supervision of an experienced AE is oneway to learn them. Another place to look forapproved methods is in Installation PracticesAircraft Electric and Electronic Wiring, NAVAIR01-1A-505. It contains the accepted way ofperforming many practical operations, and it iswell illustrated.

Often, parts replacement is the best way tocorrect a malfunction. However, never use partsreplacement as a method of troubleshooting. Eachyear, thousands of dollars worth of instrumentsand black boxes are returned to overhaul activitieswith labels stating that they are faulty. Uponcompletion of test and check by the overhaulactivity, many have no defect. This practice iswasteful and very expensive. The replacement ofan entire unit when only a small part is at faultreflects poor maintenance practices. A typicalexample of such a case is the replacement of anentire compass amplifier when there is only afaulty fuse. When troubleshooting, remember youare working with equipment that is both expensiveand scarce. Make decisions to replace equipmentonly after thoroughly testing it and making sureit is faulty.

For the procedures to follow when replacingand testing equipment, refer to the removal,installation, and testing section of the MIM. Someof the topics covered in the MIMs are discussedin the following paragraphs.

Removal. Remove equipment in such a waythat you don’t damage it more and you don’tdamage nearby structures. Pay attention so youdon’t loose or misplace small parts. They maybeeasy to replace with new ones, but small items lost

in the aircraft are a very real foreign objectdamage (FOD) hazard. FOD kills!

If you can’t finish installing the new partimmediately, remove all tools and parts from theaircraft during the waiting period. You mightforget and leave them in the aircraft. Thisprecaution applies to any unfinished job involvinga waiting period. Before you remove the defectivepart from the aircraft, pull and tag circuitprotective devices and notify maintenance control,This prevents people from turning on fuel pumpswhen fuel lines are open. Also, it prevents powerbeing applied to loose cables.

Installation. Many of the rules for removalapply to installation. However, there are a fewrules that apply almost exclusively to installingnew equipment. Some units have special condi-tions while in shipment. Look for these conditionsbecause they often require you to remove sealingplugs, locking devices, and other specialequipment from the component before you caninstall it. If adjustments are necessary on newequipment, you might need to make the adjust-ments before completing the installation. Forexample, often the matching of a positiontransmitter to the position indicator is easier todo with only the electric wiring connected tothe actuator. After calibration, complete themechanical mounting. Again, the only guide formaking adjustments to any system is the MIM.In the final installation, make sure that allmounting hardware is in place and properlytightened. Also, make sure all cable and lineconnections are secure and connected to theproper points.

Testing. The entire process of detecting andcorrecting a fault means nothing unless the systemoperates properly. Therefore, the final step of thejob is to test the equipment. This test is usuallyan operational test. Where possible, simplyoperate the new component or system to see thatit is doing its job properly. When a final checkshows the job is complete, you must pass thisword quickly to those concerned. Officially, theaircraft is down until the proper people arenotified.

Q1. List the three steps that will make yourtroubleshooting easier.

What is the best test equipment to use whenlooking for shorts, grounds, opens, andwrong resistances?

You should never use aas a troubleshooting method because itis expensive and does not always fix thefault.

After removing apart, you should alwayslook for small items from the work toensure they do not become what type of ahazard?

List the tools that help you analyze asystem.

What is the first action to take whenbeginning to troubleshoot apiece of equip-ment?

What types of common faults interruptpower through a circuit?

List a reference to which you can refer forapproved maintenance procedures.

You should refer to thefor procedures to follow when removingand replacing equipment?

What is the purpose of testing a piece ofequipment after it has been repaired?

USE OF BASIC TEST EQUIPMENT

Learning Objective: Recognize the varioustypes of general-purpose test equipmentsassociated with aircraft electricalmaintenance, and identify tests the AE willmake using these equipment.

Test equipment, like any other tool the AEuses, is susceptible to damage, misuse, anddeterioration. Knowledge of the Navy calibrationprogram is essential to ensure your shop ordivision has the most reliable test equipment. Formore information on the calibration program,refer to the latest edition of Aviation MaintenanceRatings Fundamentals, NAVEDTRA 10342-3,and NEETS, module 16. Also, when making testson equipment, you should adhere to the followingrules:

1. Always connect an ammeter in series—never in parallel.

2. Always connect a voltmeter in parallel—never in series.

3. Never connect an ohmmeter to anenergized circuit.

4. On test meters, select the highest rangefirst, then switch to lower ranges as needed.

5. When you use an ohmmeter, select a scalethat gives a near midscale reading, since midscaleis where the meter is most accurate.

6. Do not leave the multimeter selectorswitch in a resistance position when not using themeter. The leads may short together dischargingthe internal battery. There is less chance ofdamaging the meter if the switch is on a high acvolts setting, or in the OFF position. Meters thathave an OFF position dampen the swing of theneedle by connecting the meter movement as agenerator. This prevents the needle from swingingwildly when moving the meter.

7. View the meter from directly in front toeliminate parallax error.

8. Observe polarity when measuring dcvoltage or current.

9. Do not place meters in the presence ofstrong magnetic fields.

10. Never measure the resistance of a meteror a circuit with a meter in it. The high currentrequired for ohmmeter operation may damagethe meter. This also applies to circuits withlow-filament-current tubes and to some types ofsemiconductors.

11. When you are measuring high resistance,do not touch the test lead tips or the circuit. Doing

so may cause body resistance to shunt the circuit,causing an erroneous reading.

12. Connect the ground lead of the meter firstwhen making voltage measurements. Work withone hand whenever possible.

CONTINUITY TESTS

In open circuits, current flow stops. The causemay be a broken wire, defective switch, etc. Todetect open circuits, you perform a continuity test,which will tell if the circuit is complete orcontinuous.

To make a continuity test, you will use anohmmeter. Normally, you will make this test ina circuit where the resistance is low, such as theresistance of a copper conductor. If the resistanceis very high or infinite, the resistance between thetwo points shows an open circuit.

Look at figure 2-2. It shows a continuity testof a cable connecting two electronic units. Youcan see that both plugs are disconnected, and theohmmeter is connected in series with conductorA, which is under test. The power should be off.When checking conductors A, B, and C, thecurrent from the ohmmeter flows through plug2 (female), conductor A, B, or C, to plug 1(female). From plug 1, current passes through thejumper to the chassis, which is grounded to theaircraft structure. The aircraft structure serves asthe return path to the chassis of unit 2 andcompletes the circuit through the series-connectedohmmeter. The ohmmeter shows a low resistancebecause no break exists in conductors A, B, orC. However, checking conductor D (shown byreading in figure) reveals an open. The ohmmetershows maximum resistance because currentcannot flow in an open circuit. With an opencircuit, the ohmmeter needle is all the way to theleft since it is a series-type ohmmeter (reads rightto left).

If you can’t use the aircraft structure as thereturn path, use one of the other conductors(known to be good.) This technique will alsoreveal the open in the circuit.

GROUNDED CIRCUIT TEST

Grounded circuits happen when some circuitpart makes contact either with the aircraft metallicframework or with a metal object that is a ground.The most common cause of grounds is insulationfraying from a wire, allowing the conductor tocome in contact with aircraft structure.

Figure 2-2.-Continuity test.

When testing for grounds, you use anohmmeter or some other continuity tester. Bymeasuring the resistance to ground at any pointin a circuit, you can determine if the point goesto ground. Look at figure 2-2. After removing thejumper from pin A of plug 1, you can test eachcable conductor for grounds. To do this, connectone meter lead to ground and the other to eachof the pins of one of the plugs. A low resistanceshows that a pin is shorting to ground. For thistest you remove both plugs from their units; ifyou remove only one plug, a false indication ispossible.

SHORT TEST

A short circuit occurs when two conductorsaccidentally touch each other directly or throughanother conducting element. In a short circuit,enough current may flow to blow a fuse or opena circuit breaker. It is possible to have a shortbetween two cables carrying signals and the shortnot blow the fuse.

You use an ohmmeter to check for a short.By measuring the resistance between two con-ductors, you may detect a short between them bya low resistance reading. Refer to figure 2-2. Ifyou remove the jumper and disconnect both plugs,you can make a short test. You can do thisby measuring the resistance between the twosuspected conductors.

Shorts not only happen in cables, they occurin many components, such as transformers, motorwindings, capacitors, etc. The major method for

testing such components is to make a resistancemeasurement and compare the indicated resist-ance with the resistance given on schematics orin maintenance manuals.

VOLTAGE TESTS

Voltage checks are made while the power isapplied; therefore, you must follow the prescribedsafety precautions to prevent injury to person-nel and equipment damage. Voltage tests areimportant. You will use them to isolate mal-functions to major components and to maintainsubassemblies, units, and circuits.

The voltmeter is used to test voltages. Whenyou use the voltmeter, make sure the meter is thecorrect one for the type current (ac or dc) undertest. Also, make sure it has a scale with a suitablerange. Since defective parts in a circuit may causehigher than normal voltages to be present at thepoint of test, use the highest voltmeter rangeavailable first. Once you get a reading, you candetermine if a lower scale is possible. If so, youuse the lower scale as it provides a more accuratereading.

Another consideration in the circuit voltagetest is the resistance and current in the circuit. Alow resistance in a high-current circuit may causea considerable voltage drop. The same resistancein a low-current circuit may cause a minimalvoltage drop. You can check abnormal resistancein part of a circuit with either an ohmmeter ora voltmeter. Where practical, use an ohmmeterbecause the test is then on a dead circuit.

Normally, you will work on low-currentelectronic circuits. Schematics show the voltagesat various test points. If you suspect that a certainstage is defective, you may check the voltage byconnecting a voltmeter from the test point toground. If the suspected stage is good, thevoltmeter readings will match the voltages givenon the schematic.

Some technical manuals contain voltagecharts. These charts usually show the sensitivityof the meter (e.g., 20,000 ohms/volt) you shoulduse to take the voltage readings for the chart. Toget comparable results, you must use a voltmeterof the same sensitivity (or greater) as thatspecified. This is to be sure the voltmeter is notloading the circuit while taking a measurement.If the meter resistance is not considerably higherthan circuit resistance, the reading will bemarkedly lower than true circuit voltage becauseof the voltmeter’s loading effect.

AMMETER

The ammeter connects in series with thecurrent path. Circuits requiring frequent currentreadings or adjustments provide current jacks foruse with a plug-in meter. Some systems have ameter installed as part of the equipment.

OHMMETER

The Navy does not have an instrumentconsisting solely of an ohmmeter. The ohmmeterand the voltmeter form a multimeter. Therefore,you must determine the choice of ohmmeterby the resistance ranges available in thevarious multimeters. Small multi meters havea high range of R x 100; larger multimeters,such as the AN/PSM-4, have a high range ofR x 10,000. VTVM TS-505A/D has a high rangeof R x 1,000,000.

WHEATSTONE BRIDGE

Resistance measurements taken using anohmmeter are not always accurate enough. Thecause of this inaccuracy is an error in metermovement and in the reading of the meter. TheWheatstone bridge (fig. 2-3) is widely used forprecise resistance measurements.

Resistors R1, R2, and R3 are precision,variable resistors. The value of Rx is an unknownvalue of resistance that you must determine. Afterproperly balancing the bridge, you can find theunknown resistance by using a simple formula.

Figure 2-3.-Wheatstone bridge.

The galvanometers (an instrument that measuressmall amounts of current) across terminals b andd shows the condition of balance. When the bridgeis in balance, no difference in potential existsacross terminals b and d; when switch S2 closes,the galvanometers reading is 0.

When the battery switch (S1) closes, electronsflow from the negative terminal of the battery topoint a. Here, the current divides as it would inany parallel circuit. Part of it passes through R1and R2; the remainder passes through R3 and RX.The two currents, I1 and I2, unite at point c andreturn to the positive terminal of the battery. Thevalue of I1 depends on the resistance of R1 andR2, and the value of I2 depends on the resistancesof R3 and RX . The current is inverselyproportional to the resistance.

Adjust R1, R2, and R3 so when S1 closes,and no current flows through G. When thegalvanometers shows no deflection, there is nodifference in potential between points b and d.All of I1 follows the path, and all of I2

follows the path. This means that voltagedrop E1 is the same as voltage drop E3. Similarly,the voltage drops across R2 and RX (E2 and EX)are also equal.

With this information, you can figure thevalue of unknown resistor RX. Divide the voltagedrops across R1 and R3 by their respective voltagedrops across R2 and RX as follows:

Then, multiply both sides of the expression by RX

to separate it:

For example, in figure 2-3, you know that R1 is60 ohms, R2 is 100 ohms, and R3 is 200 ohms.To find the value of RX, use the formula asfollows:

The Wheatstone bridge is widely used to getprecise resistance measurements. You can measurecapacitance, inductance, and resistance for preciseaccuracy by using ac bridges. These bridges consistof capacitors, inductors, and resistors in a widevariety of combinations.

DC VOLTMETER

When selecting a dc voltmeter, you shouldconsider the sensitivity of the meter movementand its effect on the circuit under test. NavyElectricity and Electronics Training Series(NEETS), Module 16, discusses meter sensitivity.

Use a multimeter having low sensitivity forquick, rough readings where approximations areadequate. When desiring a high degree ofaccuracy, use a meter having high sensitivity. Sucha meter has wide application in the maintenance

of medium- and high-impedance electronic circuitsfound in aviation electrical/instrument systems.

A vacuum tube voltmeter, because of its highinput impedance, is the ideal instrument formeasuring low voltage in oscillators, automaticgain control, automatic frequency control, andother electronic circuits sensitive to loading, Whenmeasuring voltages of more than 500 volts, amultimeter having a sensitivity of 20,000 ohms pervolt has an input impedance comparable to mostvacuum tube voltmeters. The input impedance ofmost vacuum tube voltmeters is between 3 and10 megohms. A 20,000-ohm-per-volt meter, whenreading a voltage of 500 volts, has an inputimpedance of 500 x 20,000 or 10 megohms.Therefore, on the 500-volt scale, a multimeter ofthis sensitivity has an input impedance at leastequal to the vacuum tube voltmeter. For voltagereadings over 500 volts, a 20,000-ohms-per-voltmultimeter offers an input impedance higher thanthe vacuum tube voltmeter.

You can use any multirange voltmeter, thoughits sensitivity may not exceed 1,000 ohms per volt,to get fairly reliable readings in a dc circuit. Ifyou do not know the impedance of the circuitunder test, a comparison of two voltage readingswill show if the meter is having a loading effecton the circuit. Take the voltage readings on thelowest usable range and the on the next higherrange. If the two readings are approximately thesame, the meter is not causing appreciable voltagevariations, and you can accept the higher readingas the true voltage. If the two readings differconsiderably, the true voltage may be found bythe following formula:

where E is the true voltage

E1 is the lower of the two voltage readings.

E2 is the higher of the two voltage readings.

R is the ratio of the higher voltage range tothe lower voltage range.

As an example of how the formula works,make the following assumptions:

1. A reading of 22 volts was obtained betweentwo terminals with the meter on the 0-30 volt scale.

2. A reading of 82 volts was obtained fromthe same terminals with the meter on the 0-300volt scale.

The true voltage is found as follows:

E = 82 - 22 + 82 22 × 10 - 1

E = 117.7 volts

MULTIMETER

During troubleshooting, you will often measurevoltage, current, and resistance. Rather than usingthree or more separate meters for thesemeasurements, you can use the multimeter. Themultimeter contains circuitry that allows it to be avoltmeter, an ammeter, or an ohmmeter. Amultimeter is often called a volt-ohm-milliammeter(VOM).

One advantage of a VOM is that no externalpower source is necessary for its operation. Therefore,

no warm-up is necessary. It is a portable, versatilemeter that is free of calibration errors caused by agingtubes or line voltage variations.

The VOM does have two disadvantages:

1. The VOM can load the circuit under test.2. The meter movement is easy to damage

because of improper testing procedures.

The VOM (fig. 2-4) includes all the necessaryswitches, jacks, and additional devices arranged in acompact, portable case. It uses one meter movement.The permanent-magnet, moving-coil metermechanism responds only to direct current. To adaptthis mechanism for measuring alternating currentand voltages, a rectifier or thermal converter must beused.

The use of a rectifier affects the calibration of themeter. Waveforms of various shapes have

Figure 2-4.—Schematic diagram of a typical multimeter.

different average and effective values. Since thismeter mechanism responds only to the averagevalue, you must take into consideration thewaveform shape when calibrating the meter foreffective value. Thus, the meter reading is correctonly on waveforms for which the meter iscalibrated. In this respect, the instrument has adefinite limitation because waveforms may varyconsiderably in conventional circuits. Forexample, the sine-wave voltage response of ahalf-wave rectifier is 32 percent of maximum forthe average value, and 50 percent for the effectivevalue. The full-wave rectifier response is 64 and71 percent, respectively. In the design of themeter, these values are taken into consideration,and the scales are plotted to read in effective unitswhen calibrated against an accurate standard.

DIGITAL MULTIMETER

The digital multimeter is an accurate, reliablemeasuring instrument used at the organizational,intermediate, and depot levels of maintenance.The instrument is easy to use and read. Also, somedigital multimeters have a battery pack for theportable operation.

The Fluke Model 8100A (fig. 2-5) is repre-sentative of digital multimeters presently used in

the fleet. It is a small, compact unit that weighs10 pounds, including the optional battery pack.It is capable of measuring ac and dc voltages toa maximum of 1,000 volts and resistance to 10megohms. Standard features of the Model 8100Ainclude protection against overvoltages, a select-able input filter, autopolarity. push-buttonfunction, and range selection. Also, it has a fullfour-digit readout plus a 1 in a fifth digit to showoverranging. The unit operates from voltages ofeither 115 or 230 volts at frequencies from 50 to500 Hz. An optional rechargeable nickel-cadmiumbattery pack can power the unit for 8 continuoushours. Accessories and options besides therechargeable battery pack are a high-frequencyprobe, a high-voltage probe, and switched ac/dccurrent shunts.

MEGGER

Ordinary ohmmeters cannot be used tomeasure multimillion ohm values of resistances,such as those in conductor insulation. To test forinsulation breakdown, you use a much higherpotential than that supplied by the battery of anohmmeter. This potential is placed between theconductor and the outside of the insulation. Youwill use a megger (megohmmeter) for these tests.

Figure 2-5.-Digital multimeter.

The megger (fig. 2-6) is a portable instrumentconsisting of two main elements.

1. The hand-driven dc generator, whichsupplies the voltage for making themeasurement, and

2. the instrument portion, which shows thevalue of the resistance you are measuring.

The instrument portion is of the opposed-coiltype, as shown in view A. Coils a and b are onmovable member c. A fixed angular relationshipexists between coils; they are free to turn as a unitin a magnetic field. Coil b moves the pointercounterclockwise, and coil a moves it clockwise.

Coil a connects in series with R3 and unknownresistance RX. The combination of coil a, R3, andRX forms a direct series path between the + and– brushes of the dc generator. Coil b connectsin series with R2. This combination also connects

across the generator. Notice that the movablemember (pointer) of the instrument portion of themegger has no restoring springs. Therefore, whenthe generator is not operating, the pointer willfloat freely and may come to rest at any positionon the scale.

The guard ring, shown in figure 2-6, view A,shunts any leakage currents to the negative sideof the generator. This prevents such current fromflowing through coil a and affecting the meterreading.

If the test leads are open, no current will flowin coil a. However, current will flow internallythrough coil b and deflect the pointer to infinity.This reading shows a resistance too large tomeasure. When you connect a resistance, such asRX, between the test leads, current also flows incoil a; this moves the pointer clockwise. At thesame time, coil b moves the pointer counter-clockwise. Therefore, the moving element,

Figure 2-6.-Megger internal circuit and external view.

composed of both coils and the pointer, comesto rest at a position in which the two forcesbalance. This position depends upon the value ofRX, which controls the current in coil a. Becausechanges in voltage affect both coils in the sameproportion, the position of the moving elementis independent of the voltage. If you short the testleads together, the pointer will come to rest at zerobecause the current in coil a is relatively large.Since R3 limits the current, the instrument is notdamaged under these circumstances. Figure 2-6,view B, shows the external appearance of one typeof megger.

Most meggers you will use have a 500-voltrating. (Normally, meggers have friction clutches.)When you crank the generator faster than its ratedspeed, the clutch slips. This prevents the generatorspeed and output voltage from exceeding ratedvalues. A 1,000-volt generator is available forextended ranges. When you want to measure anextremely high resistance (10,000 megohms ormore), a high voltage is needed to cause enoughcurrent flow to operate the meter movement.

When you use a megger, observe the followingprecautions:

• When making a megger test, make sure theequipment is de-energized.

• Observe all rules for safety in preparingequipment for test and in testing, especiallywhen testing installed high-voltage apparatus.

• Use well-insulated test leads, especiallywhen using high-range meggers. After con-necting the leads to the instrument and beforeconnecting them to the component undertest, operate the megger to make sure there isno leakage between leads. The reading shouldbe infinity. Make certain the leads arenot broken and the connections are good bytouching the leads together while turning thecrank slowly. The reading should be zero.

�• When using high-voltage meggers, takeproper precautions against electric shock.There is enough capacitance in most electricalequipment to store up sufficient energy fromthe megger to give a very disagreeable andeven dangerous electric shock. The megger hasa high protective resistance, and its open circuitvoltage is not as dangerous, but always be careful.

• Discharge equipment having a consider-able capacitance before and after makingmegger tests to avoid the danger ofreceiving a shock. You can do this bygrounding or short circuiting the terminalsof the equipment under test.

OSCILLOSCOPE

An oscilloscope is a versatile instrument usedin viewing wave shapes of voltages or currents.It is capable of giving information concerningfrequency values, phase differences, and voltageamplitudes. It can also trace signals throughelectronic circuits to localize sources of distortionand to isolate troubles to particular stages.

The majority of oscilloscopes used for test andother measurements contain a basic presentationunit (screen) and accessory units, such asamplifiers, synchronizing circuits, time-delaycircuits, and sweep oscillators. These circuits areused to display a stationary pattern on thecathode-ray tube (CRT). The sensitivity of anoscilloscope should be adequate for the smallestsignals to have sufficient amplitude for screendisplay.

NEETS discusses the CRT, the oscilloscope,and their operation in detail.

TIME-DOMAIN REFLECTOMETRY

Time-domain reflectometry (TDR) is ameasurement concept widely used in the analysisof wideband systems. The art of determining thecharacteristics of electrical lines by observationof reflected waveforms is not new. For manyyears power-transmission engineers have locateddiscontinuities in power-transmission systems bysending out a pulse and monitoring the reflections.Discontinuity is any abnormal resistance orimpedance that interferes with normal signal flow.

TDR is particularly useful in analyzing coaxialcables, such as those in fuel or oxygen quantityindicating systems. The amplitude of the reflectedsignal corresponds directly to the impedance ofthe discontinuity. You can find the distance to thediscontinuity by measuring the time requiredfor the pulse to travel down the line to thereflecting impedance and back to the monitoringoscilloscope.

TDR analysis consists of inserting an energystep or pulse into a system and the subsequentobservation, at the insertion point, of the energyreflected by the system. Several arrangements arepossible, but the following procedure is used with

the newer, specialized reflectometers (fig. 2-7). The pulsegenerator develops a fast (or incident) step. This step thenpasses through a TEE connector and goes into the systemunder test. The sampling oscilloscope is also attached tothe TEE connector, and the incident step, along with thereflected waveform, shows on the CRT. Analysis of themagnitude, duration, and shape of the reflected waveformshows the type of impedance variation in the systemunder test.

When the fast-rise input pulse meets with adiscontinuity or an impedance mismatch, the resultantreflections appearing at the feedpoint are compared inphase, time, and amplitude with the original pulse. Also,since distance relates to time and the amplitude of thereflected step directly relates to impedance, thecomparison shows the distance to the fault, as well as anindication of its nature. Time-domain reflectometry showsboth the position and the nature (resistive, inductive, orcapacitive) of each line discontinuity. It also reveals thecharacteristic impedance of the line and shows whetherlosses are parallel or series losses.

PHASE ANGLE VOLTMETER

You can determine the overall accuracy of manyelectronic parts by measuring phase angles in computingtransformers, computing amplifiers, and resolver systems.In the past, the most common method used for measuringphase shift or phase angles between signals wasobserving patterns on an oscilloscope. With this method,it was hard to determine small angles and difficult totranslate various points into angles and sines of angles.When one of the signals contained harmonic distortion ornoise, this interference limits the use of oscilloscopepatterns.

In any complex waveform containing a basicfrequency and harmonics, measuring phase shifts

Figure 2-7.-Typical time domain reflectometer.

presents problems. In most applications, interest lies inthe phase relationship of the basic frequencies, regardlessof any harmonics that are present. One requirement of aphase measuring device is measuring the phase differencebetween two discrete frequencies. It must accomplish this,regardless of phase and amplitude of other components ofthe waveform.

The basic block diagram of a phase angle voltmeter isshown in figure 2-8. You should refer to it while readingthis section. There are two inputs—the signal and thereference. Both channels contain filters that pass only thefundamental frequency. Harmonics are highlyattenuated. Each channel has a variable amplitudecontrol and amplifiers to increase the variety of signalsyou can check.

By placing a calibrated phase shifter into thereference channel, that channel signal can be phaseshifted to correspond to the other channel. The phasedetector will detect this action and it will also show on themeter.

The calibrated phase shifter connects to a switch(whose position corresponds to the 0-degree, 90-degree,180-degree, and 270-degree phase shift) and apotentiometer (whose dial is calibrated from 0 degree to90 degrees). The total phase shift is the sum of the tworeadings.

The phase detector is a balanced diode bridge-typedemodulator. Its output is proportional to the signalamplitude times the cosine of the angle of phasedifference between the signal input and the referenceinput.

If the reference input is phase shifted until it is inphase or 180 degrees out of phase with the signal input,the output from the phase detector is proportional to thesignal input amplitude (the cosine of the angle is unity). Ifthe reference input is phase shifted until it is 90 degreesout of phase

Figure 2-8.-Phase angle voltmeter, block diagram.

with the single input, the phase detector outputis zero (the cosine of the angle is zero).

The point where the two signals are in phaseor 180 degrees out of phase is the point ofmaximum deflection on the meter. The differencebetween the in-phase and 180 degrees out-of-phasepoints is in the direction of needle swing, not thedistance it swings. Upon approaching the pointof maximum deflection, the rate of change of themeter reading decreases. This happens because thecosine has a smaller rate of change near 0 degree.This makes it difficult to read the exact point ofmaximum deflection.

Most commercial voltmeters determine thepoint at which the signals are 90 degrees out ofphase as quadrature. The cosine has a maximumrate of change as it approaches 90 degrees, thusit gives a better indication on the meter. Whenthe voltmeter is setup for this point, there mustbe some way of converting the phase shifterreading. You have to convert the reading to showthe correct amount of phase shift, rather than 90degrees more or less than the actual amount.Some confusion exists in this area because

different manufacturers have different methodsof determining the signal quadrant. Manu-facturers also differ on whether the final readingis a leading or a lagging phase shift. This meansthat you should be familiar with the type of phaseangle voltmeter you are using. You can’t assumethat the method used to determine phase angleon one type of meter can be used on another.

DIFFERENTIAL VOLTMETER

The differential voltmeter is a reliable pieceof precision test equipment that providesextremely accurate volt age measurements. Itsgeneral function is to compare an unknownvoltage with a known internal reference voltageand show the difference in their values. The Model893A, differential voltmeter (fig. 2-9) is commonlyused by naval personnel.

You can use the differential voltmeter as aconventional transistorized electronic voltmeter(TVM) or as a differential null voltmeter. You canalso use it to measure variations of a voltage nearsome known value (null detector), high resistance

Figure 2-9.-Differential voltmeter.

values (megohmmeter), and for dBm measure-ments. This meter is a solid-state differentialvoltmeter providing the capability of making dcvoltage measurements from ± 10 microvolt to± 1,100 volts. It also measures ac voltages from0.001 to 1,100 volts over a frequency range from5 hertz to 100 kilohertz. You can make both ofthese measurements without concern of loadingthe circuit. It has four voltage readout dials thatvary the resistance of the divider assembly asdescribed above.

The differential voltmeter uses a built-in nulldetector to measure an unknown voltage. Themeter circuitry compares the unknown voltage toa known, adjustable reference voltage suppliedby the meter. The reference voltage is from ahigh-voltage dc power supply and decade resistordivider assembly strings. You use voltage readoutdials to adjust the decade resistor assembly. Inthis way, you can divide the output from thehigh-voltage power supply precisely into incre-ments as small as 10 microvolt. The readout dialsadjust the meter pointer to 0, and the unknownvoltage is on the voltage dials.

A primary feature of a differential voltmeteris that it does not draw current from the unknownsource for dc measurements when you make themeasurement. Therefore, the determination of theunknown dc potential is independent of its source.

Q1. Describe how you would connect avoltmeter to the circuit under test?

Q2. When using a meter, you should start atwhat range?

Q3. A circuit in which current no longer flowsis known as an .

Q4. What test do you perform to find an opencircuit?

Q5. Blown fuses and open circuit breakers areusually an indication of what type of fault?

Q6. List the two disadvantages of the volt-ohm-milliammeter.

Q7. What type meter should you use to test forinsulation breakdown?

Q8. What test equipment shows you the waveshape of current or voltages?

Q9. What is a discontinuity?

Q10. What meter compares an unknown voltagewith a known internal reference voltage andshows the difference in their values?

AIRCRAFT WIRE AND CABLES

Learning Objective: Recognize aircraftwire and cable characteristics and variousmeans of identifying and splicing wires andcables.

An important part of aircraft electricalmaintenance is determining the correct wire orcable (fig. 2-10) for a given job. For electricalinstallations, a wire is a stranded conductor,covered with an insulating material. The termcable, as used in aircraft electrical installation,includes the following:

Two or more insulated conductors con-tained in the same jacket (multiconductorcable)

Two or more insulated conductors twistedtogether

One or more insulated conductors coveredwith a metallic braided shield (shieldedcable)

A single insulated center conductor witha metallic braided outer conductor (RFcable)

Figure 2-10.-Cables commonly used in aircraft.

WIRE REPLACEMENT

Before you can replace wiring, you need tofind information about the wire. First, consult theaircraft’s maintenance instructions manual (MIM)for wire replacement since it normally lists the wireused in a given aircraft. When you cannot get thisinformation from the manual, you must select thecorrect size and type of wire needed. The factorsin determining correct wire size are in NEETS,module 4, and Installation Practices AircraftElectric and Electronic Wiring, NAVAIR01-1A-505. The procedures specified in NAVAIR01-1A-505 are mandatory for the maintenance ofnaval aircraft. The information in these publi-cations and in the latest Military SpecificationMIL-W-5088 should help you select the correctreplacement wire.

The data necessary to determine the correctwire for a given application is summarized below:

Current drawn by the load (table 2-1)Length of wire required to go from thesource to the loadAllowable voltage drop between the sourceor point of voltage regulation and the loadThe maximum voltage that is applied to thewireThe approximate ambient air temperatureof the wire installation locationWhether the conductor is a stranded wirein free air or one of a group of wires in abundle or in a conduit

Wire Selection and Characteristics

After you have determined the wire size,consider insulation characteristics. The followingare descriptions of the various military specifi-cations of wires and cables. These militaryspecification numbers are on the reel, spool, orshipping container. Always refer to NAVAIR01-1A-505 before selecting wires and cables forweight discrepancies.

MIL-W-22759. MIL-W-22759 is a fluoro-polymer-insulated single conductor electric wire.It is made of tin-coated, silver-coated, or nickel-coated wires of copper or copper alloy. You mayuse this specification wire in combination withother insulating materials. Temperature andvoltage ratings range from 302°F to 500°F and600 to 1,000 volts, respectively, depending on partnumber and application.

MIL-W-25038. MIL-W-25038 is a high-temperature, fire-resistant electrical wire. Thiswire is made of nickel-clad copper with insulationthat will operate efficiently in ambienttemperatures that exceed +500°F. Its insulationdesign assures emergency operation of electricalcircuits subject to fires.

MIL-W-81044. MIL-W-8 1044 is a single con-ductor with tin, silver, and nickel-coated copperor copper alloy wires insulated with various polymaterials. You may use these poly materials aloneor in combinations, depending upon applicationand part number.

MIL-W-81381. MIL-W-81381 is a polyamide-insulated, single wire with silver and nickel-coatedconductors made of copper and copper alloys.You may use polyamide insulation alone or incombinations with other insulating materials,depending on application and part number.

MIL-C-7078. MIL-C-7078 includes threetypes of electrical cable.

Unshield/unjacketed/twisted—two ormore coded wires with no overall jacket orshield

Jacketed/twisted—two or more codedwires with no overall shield enclosed withina single jacketShielded/jacketed/twisted—one or morecoded wires, twisted, with an overall shieldenclosed within a single jacket

MIL-C-27500. MIL-C-27500 is electrical cablemade up of two to seven-wire specifications.Cables of this type are included in the followinggroups:

1. Unshielded/unjacketed—color-codedwires, spirally laid without an overall jacket

2. Jacketed—color-coded wires, spirally laidwith an overall jacket

3. Shielded/jacketed—one to seven color-coded wires, spirally laid with one or twoshields within an overall jacket

Aluminum Wire Properties

The use of aluminum wire in aircraft is wellestablished. Therefore, you need to know someof the unusual properties of aluminum.

Table 2-1.-Current Rating of Wires

Aluminum is unusual because it forms anelectrically resistant oxide film on all of itssurfaces. It has an inherent property known ascreep, which makes proper installation of aterminal extremely critical. Creep is the tendencyof aluminum to flow away from the point ofpressure. Aluminum is softer than copper, andchemically corrosive when in direct contact withcopper.

The electrically resistant aluminum oxide filmis always present. Therefore, you must eitherpenetrate or remove it to guarantee a satisfactoryelectrical connection by using the specified

compound to remove this film. (Refer to AvionicsCleaning and Corrosion Prevention/Control,NAVAIR 16-1-540.) Initially, tin-plated terminalsand splices are supplied so no oxide film is present.Terminals become dirty when stored or excessivelyhandled. You should clean them by wiping witha soft cloth. Never use a wire brush or anyabrasive method to clean a tinned aluminumsurface.

One problem found in aluminum wire iscorrosion by dissimilar metals. This problemoccurs frequently when aluminum and coppercome in contact. As soon as moisture collects

Figure 2-11.-Examples of wire identification coding.

between the two metals, an electrical differentialexists, which produces a battery effect. This effectquickly results in corrosion.

Use the following techniques to reduce theundesirable effects caused by aluminum’s soft-ness, creep, and oxide film:

Select the proper size of wire and terminals/splices. Terminals and splices used withaluminum wire will have an AL or ALUMstamp.

Select the proper hand tool when crimping.

Handle aluminum wire carefully. Use theproper assembly procedure when preparingit for an electrical connection.

Thoroughly clean any terminals.

Be careful not to scrape or nick the wirewhen stripping.

CAUTION

Never cut aluminum wire with tools thathave reciprocating motion, such as ahacksaw. Reciprocating cutting action“work hardens” aluminum. This will leadto broken and torn strands.

WIRE IDENTIFICATION

Equipment contractors assign the wire identi-fication code for wires and cables having functionletters R, S, T, and Y. The block of wire numbersfor each equipment starts with 1 and continuesfor as many numbers needed to identify all wires.For example, wires of the AN/APS-45 areidentified as APS45-1A20-APS45-975C22, thoseof the AN/ARC-52A would be ARC52-1A22-ARC52-9C22, and the MX-94 would be theMX94-1A20, etc. If a type designation (ANnomenclature) is not assigned for a piece ofequipment (such as commercial equipment), geta block of numbers from the procuring activity.You can get the wire identification code by usingthat portion of the military type designation (ANnomenclature) following the slash (/). You mustexclude the hyphen and any suffix letters.

Wire Marking

You may stamp the identification code onwires either horizontally or vertically, as shownin figure 2-11. The preferred method is to stampthe identification marking directly on the wire orcable with a hot foil stamping machine. Use this

method wherever possible. If the wire insulationor outer covering won’t stamp easily, stamplengths of insulating tubing (sleeves) with theidentification marking. Then, install the sleeve onthe wire or cable. The following types of wireusually have sleeve identification markings:

1. Unjacketed shielded wire2. Thermocouple wires3. Multiconductor cable4. High-temperature wire with insulation

difficult to mark, such as, TFE, fiber glass,etc.

CAUTION

Do not use metallic markers or bands foridentification. Do not use any method ofmarking that will damage or deform thewire or cable.

Whatever method you use to mark the wire,make sure the marking is legible and the colorcontrasts with wire insulation or sleeving. Useblack stamping for light-colored backgrounds andwhite on dark-colored backgrounds. Make surethat markings are dry so they don’t smear. Stampwires and cables at intervals of not more than 15inches along their entire lengths (fig. 2-12). Also,stamp wires within 3 inches of each junction(except permanent splices) and at each ending

Figure 2-12.-Spacing of identification stamping on wire andcable.

point. Stamp wires which are 3 to 7 inches longin the center. You don’t need to stamp wires lessthan 3 inches long.

Wiring and Cable Identification Codes

To make maintenance easier, each of theconnecting wires in an aircraft have an identi-fication. The identification is a combination ofletters and numbers. The marking also identifiesthe circuit it belongs to, gauge size, and otherinformation relating the wire to a wiring diagram.This marking is the cable identification code. Theaccompanying discussion explains the code usedin aircraft wiring. You can find complete detailsin the Military Specification: Wiring, AerospaceVehicle, MIL-W-5088.

NOTE: Stranded conductor wire is used forflexibility in installation and service. Wire sizesapproximate American Wire Gage (AWG). How-ever, they vary sufficiently so it is improper torefer to aircraft wire as AWG.

NOTE: For an in-depth study of wiringspecifications, limitations, and repair, you shouldrefer to NA01-1A-505.

The basic wire identification code for circuitsis read from left to right, except circuit functionletters R, S, T, or Y. Look at figure 2-12 as youread this section.

Unit number. Where two or more identicalitems of equipment are in the same aircraft, useprefix unit numbers 1, 2, 3, 4, etc., to differentiatebetween wires and cables. To make it easier tointerchange items, identical wiring is locatedin left and right wings, nacelles, and majorinterchangeable structural assemblies, withoutusing the unit number. Use the unit number forcircuit functions R, S, T, and Y, only wherecomplete duplicate equipment is installed. Unitnumbers don’t apply to duplicate componentswithin a single complete equipment, such asduplicate indicators or control boxes.

Circuit function letter (except R , S , T ,and Y). The circuit function letter identifies thecircuit function (table 2-2). When using a wireor cable for more than one circuit function,the circuit function letter of the predominantcircuit applies. When functional predominance isquestionable, use the circuit function letter for thewire or cable having the lowest wire number.

Wire number. The wire number consists ofone or more digits. It identifies the different wiresin a circuit. A different number is used for wire

Table 2-2.-Wiring Circuit Function Code

not having a common terminal or connection, asshown below:

1. Wires with the same circuit function havinga common terminal connection or junctionhave the same wire number but differentsegment letters.

2. Beginning with the lowest number, assigna number to each wire in numericalsequence, insofar as practical.

Wire segment letter. A wire segment is aconductor between two terminals or connections.The wire segment letter identifies differentconductor segments in a particular circuit. Use a

Such wires and cables connect to the ground net-work of aircraft electrical systems without causingany circuit malfunctions. For electronic systemswith interconnecting ground leads, but only onesegment actually grounded to structure, N identi-fies the segment actually grounded to the structure.

2. Phase letter A, B, or C is a suffix on thewire identification code. It identifies the phase orwires in the three-phase power ac distributionsystems. The phase sequence is

3. Phase letter V is a suffix on the cableidentification code. It identifies the ungroundedwire or cable in a single-phase system.

4. For thermocouple wire, use the followingdifferent letter for wire segments having acommon terminal or connection. Wire segmentletters are in alphabetical sequence. The letter Aidentifies the first segment of each circuit startingat the power source. If a circuit contains only onewire segment, this wire segment is A. Do not usethe letters I and O as segment letters. Use doubleletters AA, AB, AC, etc., when there is more than24 segments. Two permanently spliced wires do

CHROM—Chromel

ALML—Alumel

IRON—Iron

CONS—Constantan

COP—Copper

not require separate segment letters if the spliceis for modification or repair.

Wire size number. The wire size numberidentifies the size of the wire or cable. For coaxialcables and thermocouple wires, the wire sizenumber is not included. For thermocouple wires,a dash (-) is used instead of the wire size number.

Ground, phase, or thermocouple letter(s).

1. Ground cable letter N is a suffix to the wireidentification code. It identifies any wire or cable

Suffix (when required). When using aluminumwire, add ALUMINUM OR ALUM to theidentification code.

CABLE STRIPPING AND HEATSHRINKABLE TUBING

Nearly all wire and cable electrical conductorshave some type of insulation. When makingelectrical connections with wire, you must removea part of this insulation, leaving the end of thewire bare. To help you remove insulation, use awire and cable stripping tool similar to the one

that completes the circuit to the ground network. shown in figure 2-13.

Figure 2-13.-Wire stripping method.

suffixes:

The operation of this basic tool is efficient andeffective and extremely simple. To operate it,insert the wire end in the proper direction to thedepth to be stripped. Now position the wire soit rests in the proper groove for that size wire, andsqueeze.

Heat-shrinkable tubing is a plastic-like tubing(similar to insulation sleeving) that will shrink toa smaller diameter with proper heating. Place thetubing over the joint, terminal, or part needinginsulation. Now apply heat with a heat gun, oven,or other appropriate heat source. When the tubingreaches a specific temperature (shrink temperaturedepends upon the type of tubing), it quicklyshrinks around the object, forming a snugjacket. In addition to being an insulator, theshrinkable tubing helps relieve strain and addswaterproofing. Figure 2-14 gives some of thetypical uses of heat-shrinkable tubing.

SOLDERING

Wires are soldered to form a continuous andpermanent metallic connection, having a constantelectrical value. When soldering, strive forsuperior workmanship. If you are sloppy, youcreate problems and compound difficulties insystem troubleshooting techniques. Study andrefer to NAVAIR 01-1A-505 when you aresoldering.

The most important part of soldering is theselection of the correct iron for the job. Solderingirons are available in wattage ranges from 20 to500 watts. Use irons with wattage ratings of 60,100, and 200 watts for general work in aircraftelectrical wiring. Use pencil irons with a ratingof 20 to 60 watts for soldering small parts, Thesoldering iron for printed circuit soldering is alightweight, 55-watt iron with a 600°F (316°C)Curie point tip control. This iron has a three-wirecord, which prevents leakage currents that coulddamage the printed circuits.

When soldering, you should select a solderingiron with a thermal capacity high enough so theheat transfer is fast and effective. (Refer totable 2-3.) An iron with excessive heat capacitywill burn or melt wire insulation. One with toolittle heat capacity will make a cold joint in whichthe solder does not alloy with the work.

A soldering iron should also suit theproduction rate. Do not select a small pencil ironwhere you need a high steady heat flow.

Table 2-3.—Approximate Soldering Iron Size for Tinning

What type of wire should you use to carry600 to 1,000 volts with a temperature ratingbetween 302°F to 500°F?

What type of metal used in wire forms aresistant oxide film on all its surfaces?

When stamping wire identificationnumbers, at what interval should you stampthe wire?

In wire identification numbers, the suffixN means the wire completes the circuit to

What is the most important item to considerwhen soldering?

AIRCRAFT ELECTRICALHARDWARE

Learning Objective: Recognize uses forand characteristics of aircraft electrical andmechanical hardware.

You shouldn’t always use the same mountingparts that you removed from the installation.Before reusing the parts, inspect them to makesure they aren’t defective or damaged. Also check

Figure 2-14.-Typical heat shrinkable tubing.

the instruct ions; some parts can't be reused. Then,and only then, may you use the removed parts.

The hardware you should use when installingelectrical equipment in aircraft is specified in theapplicable MIM. You should always use properparts. If you need to substitute, make sure thesubstitute item is satisfactory.

General information about mounting partsand consumable components can be found invarious Navy training manuals. Aircraft StructuralHardware for Aircraft Repair, NAVAIR 01-1A-8,and Installation Practices Aircraft Electric andElectronic Wiring, NAVAIR 01-1A-505, aresources of detailed information.

If you can’t get the mounting parts specifiedby the applicable illustrated parts breakdown(IPB), you may make a temporary installationusing suitable substitute parts. Replace these partswith the proper items as soon as you receive them.Always check with the work center supervisorbefore you make any substitution. When makingpart substitutions, you should give special con-sideration to the following factors:

Corrosion. Pay attention to the chemical ormetallic composition of the part. Choose a partthat doesn’t contribute appreciably to the dangerof corrosion.

Strength. The strength of the substitute partmust be the same, or greater than the oneprescribed. (When determining the strength,consider the tensile, compression, and shearstrength, as applicable to the specific use.)

Size. Substitute nuts, bolts, and screws shouldbe the same size as the prescribed item. In allcases, washers must have the same inner diameteras the prescribed item. A different outer diameteror thickness is acceptable.

Length. The length of substitute screws orbolts must be enough for the particular installa-tion. However, length can’t be long enough to in-terfere with any moving part. Hardware shouldn’tcome in contact with other aircraft items, suchas electrical wiring, hydraulic lines, etc.

Magnetic properties. Equipment installed inspecific areas of the aircraft shouldn’t causedistortion of the magnetic fields of the area.Examples of such items include the magneticcompass, magnetic anomaly detection equipment,radio direction finder, or gyros. In areascontaining these types of equipment, any

substitute part must have the same mag-netic properties and characteristics as the oneprescribed.

Style. Most items of mounting hardware areavawe in various styles. It is usually easy to findscrews and bolts that are the same in all respectsexcept the type head. Use these parts as sub-stitutes, provided they have all required specialfeatures.

Special features. If a bolt requires torque toa given value, a suitable torque wrench for thattype part must be available. If the MIM calls forlockwire, the part must have suitable provisions.

Lubrication or coating. If specific instructionscall for lubrication or coating of the parts, followthose instructions for the substitute part as wellas for the prescribed part.

SHOCK MOUNTS

Electrical and electronic equipment must beprotected from the effects of vibration in aircraft,as vibration is a major problem. Most amplifiers,instrument panels, and other fragile parts haveshock mounting protection. Shock mounts arealso known as vibration insulators. You can tracethe failure of many systems to faulty shockmounts; therefore, you should periodically checkshock mounts. If you find that they are defective,replace them before equipment is damaged.

Figure 2-15, views A and B, shows two typesof shock mounts used in naval aircraft. View Ashows mounts that are individually replaceable.Each mount has a rod that extends into thevibration eliminating material. You can replacethis type of mount by drilling out the mountingbase rivets and riveting the replacement inposition. The replacement must be of the samesize and type as the mount that it is replacing. Theweight of the unit being protected determines thetype of mount you use. If you use a mountdesigned for a heavier unit, it won’t give to protectthe unit. [f you use a mount designed for a lighterunit, it can easily pull away from the base anddamage the unit.

The shock mount unit in figure 2-15, view B,is used with a particular piece of equipment. Youreplace this type of mount as a unit. The vibrationinsulators are made of hollow rubber and lockedinto place when manufactured. Inspect theseinsulators for cracks or splits. If you find damage,replace the complete shock mount unit.

Figure 2-15.-Typical shock mounts.

Shock absorbing materials commonly used inshock mounts are usually electrical insulators.These types of shock mounts are mounted so theyhave an electrical bond to the aircraft structure,preventing static electricity buildup. (See fig. 2-15,view A.) You need to inspect the bonding strapwhen inspecting the shock mounts. Replace orrepair defective or ineffective bonding straps.

ELECTRICAL CONNECTORS

In this section, the word connector is used ina general sense. It applies to connectors with ANnumbers and those with MS numbers. ANnumbers were formerly used for all supply itemscataloged jointly by the Army and Navy. Manyitems, especially those of older design, continueto carry the AN designator; although the supplysystem is shifting over to military specification(MS) numbers.

Electrical connectors provide detachablecoupling between major components of electricaland electronic equipment. These connectorsare built to withstand the extreme operatingconditions imposed by airborne service. Theymust make and hold electrical contact withoutexcessive voltage drop despite extreme vibration,rapid shifts in temperature, and changes inaltitude.

Connectors consist of two portions—the fixedportion, called the receptacle, and the movableportion, called the plug. Plug assemblies may bestraight or angled (usually 90 degrees). Receptacleassemblies may be of the wall-mounted, box-mounted, or integral-mounted types. MS numbersand letters identify the type, style, andarrangement of a connector.

Connectors vary widely in design and appli-cation. A coupling nut or ring holds the twoassemblies firmly together. The assembly consistsof an aluminum shell containing an insulating

insert, which holds the current-carrying contacts.The plug usually attaches to the cable end and isthe part of the connector on which the couplingnut mounts. The receptacle is the half of theconnector to which the plug connects. Thereceptacle is usually mounted on a part of theequipment.

In naval aircraft, connectors with crimp-typecontacts are widely used. Maintenance is easierbecause you can remove the contact from theconnector. If the connector is damaged, you canremove the contacts and replace the connectorshell. If just a connector pin is damaged, you canremove and replace the pin. This is a considerableadvantage over the solder-type connector, bothin convenience and time savings. A discussion ofthe special tools you need to remove and insertcrimped contacts is contained in InstallationPractices Aircraft Electric and Electronic Wiring,NAVAIR 01-1A-505.

Some common types of subminiature con-nectors are shown in figure 2-16. They areused on instruments, switches, transformers,amplifiers, relays, etc.

FABRICATION OF CABLES

Occasionally, you will have to make a cableusing connectors. The type of connector you willuse is specified in the MIM for the particularaircraft. The following steps outline the procedureyou should use to make a cable:

Disassemble the connector to allow accessto the terminals. Devise a way to hold theconnector so both hands are free.Cut the cables to the correct length.Strip the wire end with a wire stripper orknife. If you use a knife, avoid cutting ornicking the wire strands. Tin the bare wireend.Run the wires through the connectorassembly and coupling nuts.Make sure all surfaces are clean.Flow rosin-core solder into the connectorterminals.Hold the tip of the soldering iron againstthe terminal. As the solder melts, push thewire into the cavity. Hold the wire steadywhile the solder cools.

When you solder, be careful not to injure theconnector insulation with the soldering iron.Follow a prearranged sequence (fig. 2-17). Therecommended sequence is to start from the

Figure 2-16.-Subminiature connectors.

bottom connection and work left to right,moving up a row at a time. After soldering theconnections, the shields, if used, are soldered toa common terminal or ferrule. Then lace the cableand reassemble and moistureproof the connector,if necessary. Fabricating instructions are con-tained in NAVAIR 01-1A-505.

MOISTUREPROOFING

Present Navy practice is to use pottedconnectors (moistureproof or environmentproofconnectors). All jet- and carrier-type aircrafthave potted connectors. On other aircraft, usemoistureproofing sealant on electrical connectorsin areas where a chance of failure exists. All

Figure 2-17.-Connector soldering sequence.Figure 2-18.-Installation of cable terminals on terminal

block.connectors in wheel wells, wing fold areas, engineareas, engine nacelles, or cockpit decks have ahigh chance of failure and are sealed. In addition,moistureproof all connectors that interconnectflight/basic navigation equipment.

Vibration and lateral pressure fatigue wiresat the solder cup. Moistureproofing reduceselectrical connector failures by reinforcing thewires against vibration and lateral pressure.The sealing compound also protects electricalconnectors from corrosion and contamination byexcluding metallic particles, water moisture, andaircraft liquids. One result of the connector’sbetter dielectric characteristics is the reducedchance of arc-over between pins.

TERMINAL BLOCKS

Terminal blocks, made from an insulatingmaterial, support and insulate a series of terminalsfrom each other, as well as from ground. Theyprovide a way to install terminals within junctionboxes and distribution panels.

The two methods of attaching cable terminalsto terminal blocks are shown in figure 2-18. ViewA uses a standard nonlocking nut. In thisinstallation method, the use of a lockwasher isnecessary. View B shows the preferred method.When using an anchor nut, or self-locking nut,you omit the lockwasher. The use of anchor nuts

is especially desirable in areas of high vibration.In both installations, a requirement exists for theuse of a flat washer, as shown in the drawing.

Each terminal board in the aircraft electricalsystem is identified by the letters TB followed bythe number of the individual board. Each studon the terminal board is identified by a number.The lowest number in the series starts at the endnearest the terminal board identification number.The identification number is on the structure towhich the terminal board attaches. It mounts onany identification strip cemented to the structureunder the terminal board. When replacing aterminal board, don’t remove the identifica-tion marking. If the identification marking isdamaged, replace it with one that is the same asthe original.

JUNCTION BOXES

Junction boxes accommodate electricalterminals or other equipment. Individual junctionboxes are named according to their function,location, or equipment with which they areassociated. Junction boxes have a drain hole(except boxes labeled “vaportight”) at the lowestpoint to drain water, oil, condensate, or other

liquids. Figure 2-19 shows a representativejunction box for housing and protecting severalterminal blocks.

When you install a junction box, make surethe screw or bolt heads are inside the box. Don’tinstall attaching hardware so the threaded part ofthe screw or bolt protrudes inside the junctionbox. The sharp thread edges of protrudinghardware may damage wire insulation.

SUPPORT CLAMPS

Clamps provide support for conduit and openwiring and serve as lacing on open wiring. Clampsusually have a rubber cushion, or they are ofall-plastic construction. When used with shieldedconduit, the clamps are of the bonded type(fig. 2-20, view A); that is, there is a provisionfor electrical contact between the clamp andconduit. Use unbended clips for the support ofopen wiring.

A strap-type clamp (fig. 2-20, view B) or anAN 742 (fig. 2-20, view C) is used for long cableruns between panels. The preferred method forsupporting cable runs of all types is using AN 742clamps. MS 25281D plastic clamps are for usewhere the maximum temperature does not exceed250°F. When using the strap-type clamp, youmust make sure the clamps hold the cable firmly

Figure 2-20.-Cable clamps.

away from lines, surface control cables, pulleys,and all movable parts of the aircraft. Use theseclamps only as a temporary measure. Replace witha permanent installation as soon as possible.

When cables pass through lightening holes, theinstallation should conform to the examplesshown in figure 2-21. In each case, the AN 742cable clamp holds the cable firmly. Route thecable well in the clear of the edges of the lighteninghole to avoid any chance of chafing the insulation.If any wire is closer than one-fourth inch to theedge of the lightening hole, use a grommet(a rubber cushion) to protect the wires.

Protect wire bundles from the following:

High temperature

Battery acid fumes, spray, or spillage

Solvents or fluids

Abrasion in wheel wells where exposed torocks, ice, or mud

Damage due to personnel using thebundle as handholds or footsteps

Damage due to shifting cargo

Figure 2-19.-Aircraft junction box. Figure 2-21.-Routing cables through lightening holes.

Never support any wire or wire bundle froma plumbing line carrying flammable fluids oroxygen. Use clamps on these lines only to ensureseparation of the wire bundle from the plumbingline. Whenever possible, route wires and bundlesparallel with or at right angles to the stringers orribs of the area involved. See figure 2-22.

Don’t install single wires or wire bundles withexcessive slack. Slack between support points,such as cable clamps, should not normally exceedone-half inch. (This is the maximum you shouldbe able to deflect the wire with moderate handforce.) You may exceed this slack if the wirebundle is thin and the clamps are far apart. Theslack must never be so large that the wire bundlecan touch any surface. Allow a sufficient amountof slack near each end for the following reasons:

� To permit ease of maintenance

To allow replacement of terminals at leasttwice

To prevent mechanical strain on the wires,cables, junctions, and supports

To permit free movement of shock andvibration mounted equipment

To permit shifting of installed equipmentfor purposes of maintenance

CONDUIT AND FITTINGS

In many aircraft, the use of conduit is limited.This is a good practice because it saves weight andensures wide separation of cables. The separationof the electrical system makes it less vulnerableto gunfire. However, some current aircraft,especially those with limited space for wirerouting, use conduit.

Figure 2-22.-Routing cables.

Conduit comes in two basic types—flexibleand rigid. Its chief functions are to act as radioshielding and as a support and protection forwires.

Conduit fittings attach either flexible or rigidconduit to junction boxes and other equipment,and usually include ferrules and coupling nuts.Various forms of both are in use along withspecial designs of locknuts, box connectors, andcoupling adapters.

Couplings are straight or angular in design sothey fit all needs. Ferrules are bushings or flangesapplied to the ends of the conduit to give greaterstrength and support to the coupling nuts. Theyare either crimped or swagged on by the use ofcrimping or swaging tools.

SAFETY WIRE

There are three types of safety wire:

1. Lockwire is a heavy wire used to secureparts against accidental opening. Use lockwire inall areas of high vibration such as an aircraft’scompartment. Electric connectors are lockwiredin high-vibration areas that are inaccessible forperiodic maintenance and inspection.

2. Shear wire is a lighter wire used to secureparts subject to periodic disconnection, mainte-nance and inspection, or for parts you need toremove quickly.

3. Seal wire is a thin, easily breakable wireused as a seal on fire-extinguishing systems,oxygen regulators, and other emergency devicesthat need a quick release for use. You use seal wireto show whether these devices have been used ortampered with.

BONDING AND BONDING DEVICES

A bond is a union that exists between twometallic objects and results in electrical con-ductivity between them. Aircraft electricalbonding is the process of obtaining electricalconductivity between aircraft metallic parts orbetween the aircraft structure and installedequipments. An isolated conducting part or objectis one physically separate from the aircraftstructure and other conductors bonded to thestructure. A bonding connector provides theelectrical conductivity between metallic aircraftparts not in electrical contact. Bonding jumpersand bonding clamps are examples of bondingconnectors.

Purpose

An aircraft can become highly charged withstatic electricity while in flight. If the aircraftdoesn’t have a proper bond, all metal parts won’thave the same charge. A difference of potentialwill then exist between various metal surfaces.Neutralization of the charges flowing in paths ofvariable resistance produces electrical disturbances(noise) in the radio receiver. If the resistancebetween isolated metal surfaces is large enough,charges can collect, causing a spark, which is afire hazard. If lightning strikes the aircraft, a goodconducting path lessens severe arcs and sparks thatcould damage the aircraft and possibly injure itsoccupants.

The aircraft structure is also the ground forthe radio. For the radio to function properly, aproper balance is necessary between the aircraftstructure and the radio antenna. This means thesurface area of the ground must be constant.Control surfaces, for example, may becomepartially insulated from the remaining structure.This is caused by a film of lubricant building upon the hinges. This would affect radio operationif bonding did not take care of the condition.Bonding also provides the necessary low-resistancereturn path for single-wire electrical systems. Thereasons for bonding are summed up as follows:

• To reduce radio and radar interferences byequalizing static charges that accumulate

• To reduce the fire hazard by preventingstatic charges from accumulating betweentwo isolated members and creating a spark

• To minimize lightning damage to theaircraft and its occupants

• To provide the proper ground for properfunctioning of the aircraft radio

• To provide a low-resistance return path forsingle-wire electrical systems

• To provide a means of bringing the entireaircraft to the earth’s potential and keepingit that way while it is grounded to the earth

Parts Requiring Bonding

The design in current naval aircraft is to keepthe number of bonding jumpers to a minimum.As a result, jumpers are very important and need

replacing whenever necessary to keep them ingood condition. Some of the aircraft parts thatrequire bonding are identified below.

Control surfaces. Each control surface shouldhave at least two bonding jumpers, This doesn’tapply to trim tabs.

Engine mounts. Use at least four bondingjumpers across each engine mount support. Thisprovides a current path between the engine mountand aircraft structure.

Engine cowling. Use at least four sym-metrically placed bonding jumpers to bond theengine ring cowling to the engine across the rubbermounts at the front end of the cowling.

Equipment mounts. Place bonding jumpersacross shock mounts supporting electrical andradio equipment and the instrument panel.

Methods and Materials

You should install bonding connections sothey won’t break or loosen. This prevents avariation in the resistance during movement. Oneprimary objective for bonding is to provide anelectrical path of low dc resistance and low RFimpedance; therefore, it is important that thejumper be a good conductor of ample size. Makethe bonding wire as short as possible. Bond partsdirectly to the basic aircraft structure rather thanthrough other bonded parts, insofar as practical.Install bonding jumpers so they do not interferewith aircraft movable components. (See fig. 2-23.)

Contact of dissimilar metals in the presenceof an electrolyte, such as salt water, produces anelectric action (battery action), which causescorrosion in the connection. The intensity of thiselectric action varies with the kinds of metals.Bonding frequently causes the direct contact ofdissimilar metals. In such cases, the metals usedare of the kind that produce minimum corrosion.You should make connections so that if there iscorrosion, its in replaceable elements such asjumpers, washers, or separators.

Don’t use self-tapping screws for bondingpurposes. Jumpers shouldn’t be compression-fastened through plywood or other nonmetallicmaterial. When you are performing a bondingoperation, clean the contact surfaces of insulatingfinishes or surface films before assembling. Afterinstallation, refinished the assembly with asuitable protective finish. You should refer toNAVAIR 01-1A-505 for detailed informationabout bonding procedures.

Figure 2-23.-Bonding methods.

CABLE LACING AND TYING

Wire groups and bundles are laced or tied toprovide ease of installation, maintenance, andinspection. Lacing or tying keeps cables neatlysecured in groups and avoids possible damagefrom chafing against equipment or interferencewith equipment operation.

You tie a group or bundle of wires by usingindividual pieces of cord tied around the groupor bundle at regular intervals. You lace a bundleor group of wires by securing them together insideenclosures by a continuous piece of cord. You usethis cord to form loops (half hitches) at regularintervals around the group or bundle.

Cotton, nylon, or fiber glass cord are used totie or lace. The cotton cord has a wax coating tomake it resist moisture and fungus.

SLEEVING

Sleeving helps protect connections fromaccidental shorting and moisture and to lengthenthe arc-over path between contacts. Don’t useinsulating sleeves on connections or connectorsthat require moistureproofing. Don’t use sleevingon connectors that have a sealing grommet thatcovers the soldered connection.

Sleeving, commonly called spaghetti, hasmany applications in naval aviation. Some of the

applications are for covering soldered connec-tions, open bus bars, and permanent splices.

For general-purpose wiring, use either clearor opaque flexible vinyl sleeving. For high-temperature applications (160°F to 400°F), youshould use silicone rubber or fiber glass sleeving.Where resistance to synthetic hydraulic fluids orother solvents is necessary, use nylon sleeving,either clear or opaque.

Bus bars normally have panels or junctionboxes to protect them against shorting. If the busbars aren’t enclosed, use of protective coating isdesirable. You can use sleeving by slitting a pieceof vinyl tubing and wrapping it around the barafter making all connections. Be careful to choosea tubing that has a large enough diameter topermit a generous overlap when tied in place(fig. 2-24, view A).

You can temporarily protect a wire withdamaged insulation by using a sleeve of flexibletubing (fig, 2-24, view B) 1 1/2 times the outsidediameter of the wire. The length should be at least2 inches longer than the damaged portion of theinsulation. Split the sleeve lengthwise and wrapit around the wire at the damaged section. Then,tie the sleeve with nylon braid at each end andat 1-inch intervals over the entire length. Since youmust replace the damaged wire, use this procedureas an emergency measure.

To cover a permanent splice (fig. 2-24, viewC), slip the sleeving on the wire before the splicingoperation. After completing the splice, move thesleeving to cover the finished splice, and then tieit at each end with nylon braid.

You will use a slightly different arrangementfor repairing shielded wire (fig. 2-24, view D). Inthis application you use two pieces of sleeving,one for the braid and one for the wire itself.

What publication contains information onmounting hardware for aircraft parts?

When substituting hardware, what shouldyou consider before making the substi-tution?

What is the ONLY reason for clamping awire bundle to a plumbing line?

Q4. List the three types of safety wire.

Q5. What is the purpose of bonding?

MAINTENANCE OF MOTORSAND GENERATORS

Learning Objective: Recognize electricalmotor and generator maintenance proce-dures.

Since most motors and generators are notaccessible during flight, the first sign of motortrouble in many instances is complete failure.Generators and motors are mechanically similar,so the preventive maintenance you perform isbasically the same. You should be sure that themotor and/or generator mount is secure. Inspectthe mechanical linkage for proper alignment andsecurity. Check that the electrical connections aresecure and clean. Check for signs of corrosion.Make sure of proper voltage.

You should inspect the exposed portion of thewindings for evidence of overheating. If theinsulation on the windings or leads is cracked andbrittle, replace it. If, you smell burning insulation,or notice smoke or excessive noise, replace themotor.

When dismantling a motor, you shouldcarefully remove the bearings and wipe themclean. Be sure to wrap the bearings in clean oilpaper until needed during reassembly. Replacebearings showing pronounced stickiness orbumpy operation. During inspection of bearingassemblies, check for the presence of cracks andpitted surfaces. Also, check for any physicaldamage present in the bearing elements.

Most ac motors are prelubricated at the timeof manufacture and have sealed bearings. Thebearings require a visual inspection only. If thebearings are damaged, turn the motor in foroverhaul. Before reassembling the motor, wipethe rotor and stator clean with a lint-free cloth.During all cleaning operations, use the correctcleaning solvents. As a final check, conduct anoperational inspection of the motor, using theapplicable MIM.

Figure 2-24.-Typical use of sleeving.

Corrosion control is a particularly important • Detailed inspection for corrosion andfactor when dealing with rotating machinery. failure of protective systemsCorrosion inspection includes the following:

• Prompt treatment of corrosion and

• An adequate cleaning program touchup of damaged paint areas

Effective protection begins with a compre-• Thorough periodic lubrication hensive cleaning schedule. There is no single

cleaning agent or process that will clean all parts.A cleaning agent that will clean one set of partsmay not clean another. The cleaning agent maynot be usable because it attacks the alloys or metalmaking up the part. Therefore, different cleaningagents are necessary. The selection of these agentswill vary for different surfaces and equipment.You should refer to the following manuals forinformation on cleaning agents and their use:Aircraft Weapons Systems Cleaning and Corro-sion Control, NAVAIR 01-1A-509, and AvionicCleaning and Corrosion Prevention/Control,NAVAIR 16-1-540.

In addition to approved dry-cleaning solvents,you can use trichlorethylene and inhibited methylchloroform to clean electrical and electronicequipment. Inhibited methyl chloroform (tri-chlorethane) does present hazards to personneland insulation. For information on such hazards,refer to NAVSHIPs’ Technical Manual, chap-ter 9600.

Insulation can be affected by methyl chloro-form. Insulation exposed to methyl chloroformdeteriorates proportionally with the time ofexposure to the solvent. After a 5-minuteexposure, methyl chloroform noticeably softensvarnishes and reduces dielectric strength andabrasion resistance. A continuous immersion of1 hour completely destroys the properties ofmost insulating materials and varnishes used inarmatures, field coils, and similarly constructedelectrical components. Glass melamine andlaminated phenolics, however, are only slightlyaffected by such immersion.

Methyl chloroform itself is nonflammable, butafter 90 percent evaporation, the residue containsa high percentage of the flammable inhibitor.When the solvent evaporates from containers, youmust recognize the flammability of the residue.

Some of the special precautions you shouldobserve when using methyl chloroform include thefollowing:

Avoid prolonged or repeated breathing ofvapor or contact with the skin. Do not takeinternally.

Prevent contact with open flame, as thismay form highly toxic phosgene.

Immerse electrical equipment less than 5minutes, as prolonged immersion destroysmost insulating materials.

WARNING

When using cleaning solvents, avoidburning the skin or inhaling the fumes. Usesolvents in a well-ventilated room only.

Test the solvent on a portion of thecomponent. This ensures that it will notdestroy, blister, or otherwise damage thematerial.

DO NOT use on oxygen equipment—theinhibiting agent is flammable.

DO NOT use on hot equipment—accelerated evaporation will increase thetoxic hazard.

There are a variety of reasons for generatorfailures or apparent failures. Before removing agenerator that isn’t delivering its rated voltage orcurrent, determine that the trouble is not aproblem in the control, feeder, or regulatingcircuits. You should refer to the applicableoverhaul instruction manuals as a guide indetermining exact voltage and frequency toleranceduring these tests.

Periodic inspection of generators includes thefollowing:

• Check for defects in the mounting flange.If you find defects, remove and replace thegenerator. Inspect mounting studs, nuts, andsafety wire for security.

• Check removable units such as ventilationtube, air blast cover, brush inspection band,terminal box, end shield, and conduit for tightnessand general mechanical condition. Repair anyminor dents. Replace any parts with cracks.

• Check the generator and cable connectionsfor overheating. Discoloration and a burnt odorshow overheating. Clean or replace the metalparts. Repair or replace damaged cables.

• If you find defects you cannot repair withthe generator installed, remove it for re-pair or overhaul.

The maintenance and testing techniques usedto maintain ac generators apply to dc generators.For detailed instructions on a specific generator,refer to the applicable manuals that are provided

for the aircraft and component. Corrosion controlof generators is the same as for motors /inserters.

Generators are not likely to be overloaded. Adefective load (one that is highly inductive, andnot the equivalent of resistance load) may resultin low voltage, excessive field current, oversaturation, and continued low voltage. Anycondition resulting in abnormal field current mayresult in overheating. This condition is dangerousto the affected generator and the entire system,dc as well as ac.

PRINTED CIRCUITS

Learning Objective: Identify character-istics of printed circuits, including modulesand potted components, and recognizecircuit construction features.

The move toward replaceable units has leadto several new methods of electronic equipmentconstruction. The printed circuit is an example ofthis construction. This design or type of circuitgives speed and economy to maintenance, as wellas saving space and weight.

CIRCUIT CONSTRUCTION

Photoetching is one method used to manu-facture printed circuits. In this method, a plasticor phenolic sheet with a thin layer of coppercoating is used. The copper coating is covered witha light-sensitive enamel and a template of thecircuit that appears when the sheet is exposed tolight. The entire sheet is then exposed to light. Theexposed area of copper reacts to the light. Anetching process then removes this area. Theexposure of the printed circuit is similar to aphotographic exposure. The enamel on theunexposed circuit protects the copper from theetching bath that removes the exposed copper.After the etching bath, the enamel is removedfrom the printed circuit. This leaves the surfaceready for soldering parts and connections.

Some manufacturers use machinery to mountstandard parts like capacitors, resistors, andtransistors—further speeding manufacture. Cir-cuits produced by this method operate as well asconventional circuits and are easy to repair.

Look at figure 2-25. From the troubleshooters’standpoint, it shows an improved type ofconstruction that consists of a removable modularsubassembly. The modules have numerous internaland external test points to make troubleshooting

easier. Most test racks have plug-in extensions thatlet you raise the module so all parts are accessiblefor test and repair. The module is not expendable,but it is easy to repair, since all parts are ofconventional design. Miniature and subminiatureparts are so common in today’s electronic equip-ment that they are considered to be conventional.

PRINTED CIRCUIT MAINTENANCE

Although troubleshooting procedures forprinted circuits are similar to those for con-ventional circuits, repair of printed circuitsrequires considerably more skill and patience.According to OPNAVINST 4790.2, only AIMDsbeing certified as having the capability mayperform microminiature circuit repair. Activitycertification means that the AIMD has in-dividually certified technicians assigned andNAVAIR approved, operable repair equipmentavailable, and does not include any specificauthorization to repair.

Initial repair technician certification is grantedonly upon successful completion of the appro-priate NAMTRAGRUDET course or equivalentcontractor training course, approved and providedby NAVAIR. Each individual must be recertifiedperiodically to assure continued quality ofperformance.

Q1. List the manuals to which you should referfor the correct cleaning agents and proce-dures.

Q2. For a technician to repair a printed circuitboard, what certification must he/she meet?

TEST EQUIPMENT

Learning Objective: Identify functions,capabilities, and operating characteristicsof various types of aircraft test equipment.

You can classify test equipment as eithercommon or peculiar. In some cases, test equip-ment may be peculiar to a specific system, but thatsystem may be used on several different aircraft.

207.233

Figure 2-25.—Printed circuit module.

Test equipment can also be classified by whereit is used. You use line test equipment on theaircraft to determine a system’s operationalreadiness and to isolate malfunctions to aparticular part. Normally, this type equipment ispainted yellow to distinguish it from the aircraftequipment.

Bench test equipment is used in the shop toisolate malfunctions to a particular module,subassembly, or to the exact resistor, capacitor,etc. Occasionally, you will use line test equipmentwith bench test equipment to perform completechecks of a part. Whenever you use test equip-ment, follow the instructions in the MIM or otheroperating instructions for the specific testequipment, aircraft, or system.

NEVER TRY TO TROUBLESHOOTEQUIPMENT FROM MEMORY!

By using test equipment instructions and thereferences for equipment under test, you preventfalse findings. You also prevent damage to testequipment and systems and the loss of time.

MA-2 AIRCRAFT ELECTRICALPOWER TEST SET

The MA-2 aircraft electrical power test set(fig. 2-26) lets you completely test and check allaircraft ac and dc power generation system parts.The test set is of modular construction. consistingof four major assemblies:

1. Variable-speed drive2. Control console

Figure 2-26. -MA-2 aircraft electrical power test set assembly.

3. Load bank4. Oil supply system

Oil Supply System

Variable-Speed Drive Assembly

The test set variable-speed drive assembly canproduce 55 horsepower loads at speeds between2,000 and 11,000 RPM. Speed regulation is betterthan ± 1 percent in this speed range. Speed iscontrollable between 55 and 11,000 RPM. Theupper frame of the variable-speed drive housesthe drive motor reversing switch and main inputpower circuit breaker. The line radio noise filters,unit-under-test cooling air blower, and the trolleyand hoist assembly are also on the upper frame.The speed control panel on the control consoleprovides control of the gearbox output shaftspeed.

Control Console

The control console may be wall mounted,floor mounted, or mounted on the test setassembly. The control console provides the meansfor you to program and monitor the test functionremotely from the other subassemblies of the testset. A stowaway type of shelf, extending the fulllength of the control console, gives you workingspace. A card file drawer, convenient poweroutlets, and covered terminals are located on theleft-hand end of the control console. The panelscontain dead front switches and controls; and allpower exceeding 125 volts ac is within enclosuresfor operator safety. The control console consistsof a variable-speed drive panel, ac metering panel,dc metering panel, a load bank control panel, anda terminal panel.

Load Bank

The load bank provides for the loading of acgenerators, inverters, dc generators, transformer-rectifiers, and frequency converters. The loadbank houses the resistive elements (both acand dc) along with the load switching relays.The temperature sensors and a self-contained,forced-air cooling system are also located withinthe load bank housing. The load bank is limitedto selecting and applying adjustable resistive acbalanced loads from 1 to 30 kW (0 to 10 kW perphase). The dc limitation is 0- to 500-ampere dcloads.

The oil supply system is electrically driven,which provides oil cooling and pressures undertest. The system contains all controls, valves, andinstruments needed to maintain and monitor oilflow. It also maintains and monitors temperaturesand pressures of CSD/generators under test. Thetest set allows the four major assemblies to operateas one unit, or they can be separated to permitoperation in confined areas.

Q1. What are the two types of test equipment?

Q2. What test equipment is used to testgenerators, voltage regulators, and reverse-current relays in as many operating con-ditions as possible?

SEMICONDUCTOR TESTING

Since semiconductors have replaced vacuumtubes, the testing of semiconductors is vital. Inthis section, three basic types of equipment arediscussed—the Huntron Tracker 1000, HuntronTracker 2000, and the Automatic TransistorAnalyzer Model 900 in-circuit transistor tester.

Huntron Tracker 1000

You will test components with HuntronTracker 1000 using a two-terminal system, wheretwo test leads attach to the leads of the componentunder test. The 1000 tests components in-circuit,even when there are several components inparallel. The following types of devices are testedusing the Huntron Tracker 1000:

Semiconductor diodes

Bipolar and field effect transistors

Bipolar and MOS integrated circuits (bothanalog and digital)

Resistors, capacitors, and inductors

The 1000 is used on boards and systems withALL voltage sources in a power-off condition. A0.25 ampere signal fuse (F1) connects in series withthe channel A and B test terminals. Accidentallycontacting test leads to active voltage sources(e.g., line voltage, powered-up boards or systems,charged high voltage capacitors, etc.) may causethis fuse to open, making replacement necessary.When the signal fuse blows, the display showsopen circuit signatures, even with the test leadsshorted together.

CAUTION

The device to be tested must have all powerturned off and have all high-voltagecapacitors discharged before connectingthe 1000 to the device.

The line fuse (F2) should only open when therean internal failure inside the instrument.

Therefore, you should always locate and correctthe problem before replacing F2.

The front panel of the 1000 makes functionselection easy. The 1000 uses interlocking push-button switches for range selection. A toggleswitch is used for channel selection, and integralLED indicators show the active functions.

The CRT displays the signatures of the partsunder test. The display has a graticule consistingof a horizontal axis that represents voltage, anda vertical axis that represents current. Thehorizontal axis is divided into eight divisions,which lets you estimate the voltage at whichsignature changes occur. This is mainly useful indetermining semiconductor junction voltagesunder either forward or reverse bias.

Push in the power on/off switch. The 1000should come on line with the power LEDilluminated.

Before you can analyze signatures on theCRT, you must focus the 1000. To do this, turnthe intensity control to a comfortable level. Now,adjust the focus control (back panel) for thenarrowest possible trace. Aligning the trace isimportant in determining the voltages at whichchanges in the signature occur. With a short circuiton channel A, adjust the horizontal control untilthe vertical trace is even with the vertical axis.Open channel A, and adjust the vertical controluntil the horizontal trace is even with the

horizontal axis. Once set, you should not have toadjust these controls during normal operation.

Turn the power off by pushing the powerswitch in. When you turn the power on again, thesame intensity setting will be present.

The 1000 has three impedance ranges—low,medium, and high. To select these ranges, pressthe appropriate button on the front panel. Alwaysstart with the medium range; then you can adjustfor other ranges. If the signature on the CRT isclose to an open (horizontal trace), try the nexthigher range for a more descriptive signature. Ifthe signature is close to a short (vertical trace),try the next lower range.

There are two channels (channel A andchannel B) that you can select by moving thetoggle switch to the desired position. When usinga single channel, plug the red probe into thecorresponding channel test terminal. Then plugthe black probe into the common test terminal.When testing, connect the red probe to thepositive terminal of the device (i.e., anode, +V,etc.). Connect the black probe to the negativeterminal of the device (i.e., cathode, ground, etc.).By following this procedure, the signature willappear in the correct position on the CRT display.

The alternate mode of the 1000 providesautomatic switching back and forth betweenchannel A and channel B. This allows easycomparison between two devices or the same pointon two circuit boards. You select the alternatemode by moving the toggle switch to the ALTposition. The alternate mode is useful whencomparing a known good device with the samedevice whose quality is unknown.

The signal section applies the test signal acrosstwo terminals of the device under test. The testsignal causes current to flow through the deviceand a voltage drop across its terminals. Thecurrent flow causes a vertical deflection of thesignature on the CRT display. The voltage acrossthe device causes a horizontal deflection of thesignature on the CRT display. The combinedeffect produces the current-voltage signature ofthe device on the CRT display.

An open circuit has zero current flowingthrough the terminals and a maximum voltageacross the terminals. In the LOW range, adiagonal signature from the upper right to the

207XFigure 2-27.-Circuit signatures: View A-Low-range open circuit; view B—medium and high range open circuit; and view

C—all ranges short circuit.

lower left of the CRT (fig. 2-27, view A)represents an open circuit. In the HIGH andMEDIUM ranges, an open circuit shows as ahorizontal trace from the left to the right (fig.2-27, view B). When you short the terminalstogether, the maximum current flows through theterminals, and the voltage at the terminals is zero.A vertical trace from the top to the bottom of theCRT graticule in all ranges shows this short (fig.2-27, view C).

The CRT deflection drivers boost the low-leveloutputs from the signal section to the highervoltage levels needed by the deflection plates inthe CRT. The HORIZONTAL and VERTICALcontrols on the front panel adjust the position ofthe trace on the CRT display.

You use three other CRT controls to adjustthe brightness and clarity of the trace—lNTEN-SITY, FOCUS, and ASTIGMATISM. The frontpanel intensity control is the primary means ofadjusting the visual characteristics of the trace.The focus control is on the back panel and isoperator adjustable. The astigmatism trim pot isinside the 1000 on the main printed circuit board.The pot is factory adjusted to the correct setting.

Huntron Tracker 2000

The Huntron Tracker 2000 (fig. 2-28) is aversatile troubleshooting tool having the followingfeatures:

• Multiple test signal frequencies (50/60 Hz,400 Hz, 2000 Hz)

207XFigure 2-28.-Huntron Tracker 2000.

Four impedance ranges (low, medium 1,medium 2, high)

Automatic range scanning

Range control: High Lockout

Adjustable rate of channel alterationand/or range scanning

Dual polarity pulse generator for dynamictesting of three terminal devices

LED indicators for all functions

Dual channel capability for easy compari-son

Large CRT display with easy to operatecontrols

GENERAL OPERATION. —You will testcomponents using the 2000 two terminal system.It also has a three terminal system when using thebuilt-in pulse generator. When using this system,you place two test leads on the leads of thecomponent under test. The 2000 tests componentsin-circuit, even when there are several parts inparallel.

Use the 2000 only on boards and systems withall voltage sources in a power-off condition. A

0.25 ampere signal fuse connects in series with thechannel A and B test terminals. Accidental contactof the test leads to active voltage sources, suchas line voltage, powered-up boards or systems,and charged high voltage capacitors, may causethis fuse to open, making replacement necessary.When the signal fuse blows, the 2000 displaysshort circuit signatures even with the test leadsopen.

CAUTION

The device under test must have all powerturned off and all high-voltage capacitorsdischarged before connecting the 2000 tothe device.

The line fuse should only open when there isan internal failure inside the instrument. Alwayslocate the problem and correct it before replacingthis fuse.

Front Panel. —The front panel of the 2000makes function selection easy. All push buttonsare the momentary action type. Integral LEDindicators show which functions are active. Lookat figure 2-29 and table 2-4 for details about eachitem on the front panel.

207XFigure 2-29.-Front panel.

Table 2-4.-Front Panel Controls and Connectors

Table 2-5.-Back Panel Controls and Connectors

Back Panel. —Secondary controls and con-nectors are located on the back panel (fig. 2-30and table 2-5).

CRT Display. —The signature of the partunder test is displayed on the CRT. The displayhas a graticule consisting of a horizontal axis thatrepresents voltage, and a vertical axis thatrepresents current. The axes divide the display into

Quadrant 2 displays negative voltage (-V)and positive current (+I)

Quadrant 3 displays negative voltage (-V)and negative current (–I)

Quadrant 4 displays positive voltage (+V)and negative current (–I)

four quadrants. Each quadrant displays differentThe horizontal axis divides in eight divisions,

portions of the signatures.which allows the operator to estimate the voltageat which changes in the signature occur. This is

• Quadrant 1 displays positive voltage (+V) useful in determining semiconductor junctionand positive current (+I)

voltages under either forward or reverse bias.

Figure 2-30.-Back panel.207X

OPERATION OF PANEL FEATURES. —The following section explains how to use thefront and back panel features.

Turn the Power/Intensity knob clockwise. The2000 comes on with the LEDs for power,channel A, 50/60 Hz, low range, and pulse/DCilluminated.

Focusing the 2000 display is an important partof analyzing the test signatures. First you adjustthe intensity control to a comfortable level. Then,adjust the focus control (back panel) for thenarrowest possible trace.

Aligning the trace is important in determiningwhich quadrants the portions of a signature arein. With a short circuit on channel A, adjust thetrace rotation control until the trace is parallel tothe vertical axis. Adjust the horizontal controluntil the vertical trace is even with the vertical axis.Open channel A, and adjust the vertical controluntil the horizontal trace is even with thehorizontal axis. Once set, you should not have toreadjust these settings during normal operation.

Range Selection. —The 2000 has four im-pedance ranges—low, medium 1, medium 2, andhigh. You select these ranges by pressing theappropriate button on the front panel. Start withone of the medium ranges, i.e., medium 1 ormedium 2. If the signature on the CRT displayis close to an open (horizontal trace), select thenext higher range for a more descriptive signature.If the signature is close to a short (vertical trace),select the next lower range.

The High Lockout feature, when activated,prevents the instrument from entering the highrange. This feature works in either the manual orauto mode.

The auto feature scans through the fourranges—three with the HIGH LOCKOUT acti-vated at a speed set by the RATE control. Thisfeature allows you to see the signature of apart in different ranges while freeing your handsto hold the test leads.

Channel Selection. —There are two channelson the 2000-channel A and channel B. You selecta channel by pressing the appropriate front panelbutton. When using a single channel, plug thered probe into the corresponding channel testterminal. Plug the black probe into the commontest terminal. When testing, connect the red probeto the positive terminal of the device; i.e., anode,+V, etc. Connect the black probe to the negativeterminal of the device; i.e., cathode, ground, etc.Following this procedure should assure that the

signature appears in the correct quadrants of theCRT display.

The ALT mode is a useful feature of the 2000.It lets you compare a known good device with adevice of unknown quality. In this test mode, youuse common test leads to connect two equivalentpoints on the boards to the common test terminal.The ALT mode of the 2000 allows you toautomatically switch back and forth betweenchannel A and channel B, so you can easilycompare two devices. You may also compare thesame points on two circuit boards. Select the ALTmode by pressing the ALT button on the frontpanel. You may vary the alternation frequencyby using the RATE control.

NOTE: The black probe plugs into thechannel B test terminal.

When using the alternate and auto featuressimultaneously, each channel is displayed beforethe range changes. Figure 2-31 shows the sequenceof these changes.

Frequency Selection. —The 2000 has three testsignal frequencies—50/60 Hz, 400 Hz and 2000Hz. You can select these by pressing theappropriate button on the front panel. In mostcases, you should start with the 50/60 Hz testsignal. Use the other two frequencies to view smallamounts of capacitance or large amounts ofinductance.

Pulse Generator. —The built-in pulsegenerator of the 2000 allows dynamic, in-circuittesting of certain devices in their active mode. Inaddition to using the red and black probes, youuse the pulse generator. The output of the pulsegenerator connects to the control input of the

207XFigure 2-31.-Auto/alternate sequence.

device under test with one of the blue micro clipsprovided. The pulse generator has two outputs,G1 and G2, so you can test three terminal devicesin the alternate mode.

A variety of output waveforms is availableusing the pulse generator selector buttons. Firstselect the pulse mode or the dc mode using thePULSE/DC button.

In the pulse mode, the LED flashes at aslow rate.

In dc mode, the LED is continuously on.

Then select the polarity of output desired usingthe positive (+) and negative (–) buttons. Allthree buttons function in a push-on/push-offmode, and only interact with each other to avoidthe NOT ALLOWED state.

After selecting the specific output type, set theexact output using the LEVEL and WIDTHcontrols. The LEVEL control varies the magni-tude of output amplitude from zero to 5 volts(peak or dc). During pulse mode, the WIDTHcontrol adjusts the duty cycle of the pulse outputfrom a low duty cycle to 50 percent maximum(square wave). The start of a pulse is triggeredby the appropriate zero crossing of the test signal.This results in the pulse frequency being equal tothe selected test signal frequency. The WIDTHcontrol setting that selects the duty cycledetermines the end of a pulse. The WIDTHcontrol has no effect when in the dc mode.

Troubleshooting Tips

You will use the Huntron Tracker 1000 andthe Huntron Tracker 2000 to test various typesof devices and circuits. Some troubleshooting tipsare given in this section.

• Perform most tests using the medium orlow range.

• Use the high range only for testing at ahigh impedance point, or if higher testvoltages are required (i.e., to test theZener region of a 40-volt device).

�• Sometimes, component defects are moreobvios in one range than another.If a suspect device appears normal for onerange, try the other ranges.

• Use the low range when testing a singlebipolar junction, such as a diode, abase-emitter junction, or a base-collectorjunction. It offers the best signature.

• Use a higher range to check for reverse biasleakage.

• When performing in-circuit testing, do adirect comparison to a known good circuit.

• The 1000 test leads are not insulated at thetips. Be sure to make good contact to thedevice(s) under test. (NOTE: This tip per-tains to the 1000 only.)

When you troubleshoot, try relating the failuremode of the circuit under test to the type of defectthe 1000 shows. For example, expect a cata-strophic printed circuit board failure to have adramatic signature difference from that of a nor-mal device of the same type. A marginally operatingor intermittent board may have a failed part thatshows only a small pattern difference from normal.

If you cannot relate a system failure to aspecific area of the printed circuit board, beginby examining the signatures at the connector pins.This method of troubleshooting shows all theinputs and outputs. It will often lead directly tothe failing area of the board.

Devices made by different manufacturers,especially digital integrated circuits, are likely toproduce slightly different signatures. This isnormal and may not show a failed device.

Remember, leakage current doubles with every10-degree Celsius rise in temperature. Leakagecurrent shows up as a rounded transition (wherethe signatures show the change from zero currentflow to current flow) or by causing curvature atother points in the signatures. Leakage currentcauses curvatures due to its nonlinearity.

Never begin the testing of an integrated circuitusing the low range. If you initially use the lowrange, confusion can result from the inability ofthis range to display the various junctions. Alwaysbegin testing using the medium range. If thesignature is a vertical line, switch to the low range.Here you can check for a short or low impedance(less than 500 ohms). Switch to the low range ifthe device is suspect and appears normal in themedium range. This will reveal a defective inputprotection diode not evident when using themedium range.

NOTE: The 2000 test leads are conductiveonly at the tips. Be sure to make good contact withthe device(s) under test.

When testing analog devices or circuits, usethe low range. Analog circuits contain many moresingle junctions. Defects in these junctions showmore easily when using the low range. Also, the54-ohm internal impedance in the low rangemakes it less likely that parts in parallel with thedevice under test will sufficiently load the testerto alter the signature.

When testing an op amp in-circuit, compareit directly to a known good circuit. This is becausethe many different feedback paths associated withop amps can cause an almost infinite number ofsignatures.

Often when checking a Zener diode in-circuit,it will not be possible to examine the Zener regiondue to circuit leakage. If you must see the Zenerregion under this condition, unsolder one side ofthe diode to eliminate the loading effects of thecircuit.

HUNTRON TRACKER 1000. —Bipolarintegrated circuits containing internal shortsproduce a resistive signature (a straight line). Thisline begins in the 10 o’clock to 11 o’clock position.It ends in the 4 o’clock to 5 o’clock position onthe display when using the low range. This typeof signature is always characteristic of a shortedintegrated circuit. It results from a resistive valueof 4 to 10 ohms, typical of a shorted integratedcircuit. A shorted diode, capacitor, or transistorjunction always produces a vertical (12 o’clock)straight line using the low range.

HUNTRON TRACKER 2000. —Bipolarintegrated circuits containing internal shortsproduce a resistive signature (a straight line)beginning in the 1 o’clock to 2 o’clock position.This signature ends in the 7 o’clock to 8 o’clockposition when using the low range. This type ofsignature is characteristic of a shorted integratedcircuit. This results from a resistive value of 4 to10 ohms. A shorted diode, capacitor, transistorjunction, etc., always produces a vertical (12o’clock) straight line when using the low range.

Automatic Transistor Analyzer Model 900

You can use this instrument to test bipolartransistors and diodes in any one of three differentmodes. Two modes, the VIS and SND, can beused either in-circuit or out-of-circuit.

• In the VIS mode, red and green lightsflashing in or out of phase with the amberlight show the condition of the device under test.

• In the SND mode, the Sonalert™ alsoindicates good devices by beeping out ofphase with the amber light. The intent of theSND mode is to permit the operator to performin-circuit tests on transistors or diodeswithout having to look at the light display.

The third mode is the METER mode. You canonly use this for out-of-circuit testing. In theMETER mode, you may measure Beta and material identity. Also, you can measureemitter base voltage, base current (Ib), andcollector current (Ic). There are four ranges forthe Beta mode—one for small signal transistors,two ranges for medium power transistors, and onefor large power transistors.

In the VIS mode and the SND mode, themaximum voltage, current and signal levelsapplied to the device under test are within safelimits. Therefore, the device under test will receiveno damage nor will any adjacent circuitry.

This instrument will test transistors and diodesin-circuit in the VIS or SND mode if the totaldynamic shunt impedances across the junctionsare not less than 270 ohms. Also, the totaldynamic shunt for the emitter to collector mustnot be less than 25 ohms. If such should occur,the test set will give the indication for a SHORT.

The 8-inch meter, which reads from left toright, has two scales marked 0-10 and 0-50. TheO-10 range is used in the leakage collector currentand Vbe (IDENT) modes. The 0-50 range is usedin the BETA modes. Notice a mark on the meterjust short of half-scale with the nomenclatureGERMANIUM and SILICON. This mark is thereference in the IDENT mode. As the metermarkings show, those readings below the markshow the device material is germanium. Thereadings above the mark show the device is silicon.On the slanting horizontal panel immediately infront of the meter face are the appropriate testsockets and two push-button switches. One switchis ZERO and the other BETA. On the verticalfront panel immediately below the push-buttonswitches are knobs marked ADJ and CAL. At thetop center is the POLARITY switch marked PNPand NPN. In the center of the vertical front panelis the RANGE switch, the FUNCTION switchand the Sonalert™. Near the bottom of thevertical front panel are the probe jacks. The slideswitch for turning the instrument on and off isalso in this location.

VIS MODE: TRANSISTOR. —To testtransistors with the visual indication only, turn

the FUNCTION switch to the XSTR-VIS mode.The amber light should flash at about a l-secondrate. Insert the transistor under test in the propersocket.

In this mode, you perform two tests on thetransistor. The amber light shows the performanceof each test. When the amber light is out, this isthe EB-BC test mode. When the amber light ison, this is the emitter-collector test mode. The testshows good transistors by one pair of similarlycolored lights (green for NPN and red for PNP)when the amber light is off. When the amber lightis on, no lights show good transistors. The left-hand lights show the condition of the emitter-base.The right-hand lights show the condition of thebase-collector. The absence of one or all lights inthe EB-BC test mode shows an open or opens.The occurrence of both a red and a green lighton either side in the EB-BC test mode shows ashort.

For more information about the Model 900tracker, refer to Maintenance Manual, All Levelsfor Automatic Transistor Analyzer Model 900,ST810-AD-0PI-010, for patterns other than thosejust discussed. There are 96 possible patternslisted.

VIS MODE: DIODE. —You can not properlytest diodes in the XSTER mode. To test a diode,insert the diode in the proper socket and turn theFUNCTION switch to the DIODE/VIS mode. Ifthe diode is good, a pair of green lights will flashout of phase with the amber. If a pair of red lightsflash out of phase with the amber light, thediode is either installed improperly or markedimproperly. If the diode has a short, additionallights will flash out of phase with the amber. Nolights will flash in phase with the amber.

You cannot properly test transistors in theDIODE mode. When testing transistors, only onetransistor should be in the test socket at one time.Do not leave any diodes in the diode socket whiletesting transistors. When testing diodes, do notleave transistors in the transistor sockets. If youdo not observe these precautions and the devicesleft in the socket are defective, incorrect lightindications will occur. These indications maymislead the operator into believing the deviceunder test is defective.

WARNING

Unit being tested must be disconnectedfrom ac outlet, and all capacitors capableof storing electricity y should be discharged.

IN-CIRCUIT TESTING. —When testingdiodes in-circuit, attach the emitter lead to theanode of the diode. Attach the collector lead tothe cathode. When testing transistors, attach theleads to the right terminals as shown by theschematic. If the operator happens to fasten theleads to the transistor in the wrong order, anerroneous display will result. However, if thetransistor is good, the instrument will give a goodindication. The indication will be for the transistorof the opposite type. A good NPN improperlyconnected will give good PNP indications and viceversa. If the device is bad, the instrument will givea bad indication. You cannot make a qualitativeanalysis of the kind of failure unless you attachthe proper leads to the correct terminals.

To ensure the instrument will show the correcttype of transistor (PNP or NPN), you mustidentify the base lead. Use the following procedureto identify the base lead:

Disconnect the lead to the emitter terminalon the instrument. Only the light repre-senting the emitter junction should go out.Reconnect the emitter lead.Disconnect the lead to the collectorterminal of the instrument. Only the lightrepresenting the collector junction shouldgo out.Should both lights go out during the tests,the connections are incorrect.Rearrange the leads on the transistor andperform-the tests again. You should nowsee the proper results.

There are six possible combinations for theconnection of these leads. Four of thesecombinations are incorrect. These will cause theinstrument to give an incorrect indication as totransistor type (PNP or NPN). The other twocombinations will give proper indications, but youstill may not know which leads are the emitter andcollector terminals. You will know whether thetransistor is good, and whether it is an NPN orPNP transistor. If you must know which leadsare the emitter and collector terminals, it ispossible to find out after identifying the base leadusing the meter mode for Beta.

SND Mode. —In either the XSTR/SND modeor DIODE/SND mode, light patterns showinggood devices will have an accompanying beeping

™. The beeping will besound from the Sonalertout of phase with the amber light.

METER Mode. —Before testing a transistorin any of the METER modes, you should test thetransistor in one of the visual modes. This willtell you whether the transistor is an NPN or aPNP. After determining this, put the POLARITYswitch in the proper position to agree with theindication in the visual mode.

Beta. —To test the Beta of the transistor, setthe FUNCTION switch to the BETA position.Next, set the RANGE switch to the appropriateposition according to the power capability of thetransistor under test. After the RANGE switchis in the proper position, operate the push-buttonswitch marked ZERO. Now adjust the ADJ knobfor a zero reading on the meter. Next, actuate thepush-button switch marked BETA and adjust theCAL knob for full-scale deflection. Release theBETA push-button switch; now the Beta of thedevice will show on the meter. Take care inselecting the Beta range to test the transistor. Itis possible to damage small signal transistorsshould you try to test them in the 2 mA Ib (LG.PWR. XSTR) mode.

Leakage: —To test a transistor for set the FUNCTION switch on theproper position. Next, set the RANGE switch tothe 100 mA position. Then push the switchmarked ZERO and adjust the ADJ knob for azero reading on the meter. Now release the ZERObutton. Set the RANGE switch on the lowestleakage range, which will still permit less than full-scale deflection on the meter. You may now readthe leakage directly off the meter. Read the first and then . Use this order because themeter will read down scale when switching from

Also, you can increase the metersensitivity. However, if you read first andthen switch to the meterill read up scale.It is now possible to peg the meter. Although themeter has protection, avoid undue abuse.

Material Identity: Transistor. —To use thisinstrument in the IDENTITY mode, set theFUNCTION switch to IDENT. Check the ZEROADJUST on the meter as mentioned before. Aftersetting the ZERO, release the ZERO pushbutton.Now note whether the needle reads above orbelow the mark on the meter face just short ofhalf scale. If the meter reads below the mark, thedevice is a germanium transistor. If it reads abovethe mark, it is a silicon transistor. Thisinformation can be extremely useful when tryingto substitute transistors.

Leakage: Diode. —To test the reverse leakageof diodes, install the diode in the diode socket.You now determine whether the diode is good bytesting the device in the visual mode. Once youdetermine that the diode is good, place thePOLARITY switch to NPN. Turn the FUNC-TION switch to the mode, and set theRANGE switch to 100 mA. Now check to see thatthe meter is at zero, as mentioned before. Afterzeroing the meter, set the RANGE switch on thelowest range possible that still permits less thanfull-scale deflection on the meter. Read theleakage on this range.

Material Identity: Diode. —To test the materialidentity of a diode with the diode properlyinstalled in the socket. Place the POLARITYswitch in the PNP position (zero the meter) andthe FUNCTION switch in the IDENT position.Using the leads, short the base and collectorterminals together. The meter will show eithergermanium or silicon, as described before in theIDENT mode for transistors.

CAUTION

Do not identity test transistor material withthe base and collector leads shortedtogether. This may create an erroneousreading.

Model 109 Probe

The Model 109 probe, used with the Model900 tester, is easy to use, having one-handoperation. It automatically adjusts to any spacingbetween one-thirty second inch to five-eighthsinch. You can rotate each probe point in a full360-degree circle. The points are individuallyspring loaded for proper contact. You canconnect the probe to three printed circuit boardterminations. The probe has the extremely lowcontact resistance of less than .005 ohm. The useof the probe eliminates unsoldering while makingin-circuit tests of transistors, diodes, ICS, andother components. Finally, the retractable cordstretches to a full 12 feet.

DESCRIPTION. —The Model 109 three-pointprobe speeds servicing of printed circuitassemblies that have transistors, diodes, and mostother board mounted components. You can makeinstant connections to three points on a printedcircuit board. You will make rapid evaluation of

transistors using the Model 109 probe with the Model900 Automatic Transistor Analyzer in-circuit. You canaccomplish a complete test of all stages in a piece ofelectronic equipment in a matter of minutes. You canalso use the Model 109 to make temporary componentsubstitutions on the printed circuit board.

OPERATION. —Connect the leads of the Model109 probe to an appropriate piece of test equipment.Determine the connection points on the printed circuitboard to connect to the test equipment. Apply theModel 109 probe points to the circuit board. Press theprobe towards the board to ensure a good connection.The Model 109 probe green point is slightly shorterthan the yellow and blue. This allows connection ofthe collector and emitter before the base to providemaximum ease of use.

The Model 109 probe is a valuable aid whenmaking resistance and voltage measurements using aconventional VOM or VTVM. Use the yellow and blueprobe points as the negative and positive meter feeds.You can make rapid evaluations of entire circuitsfaster than with any other method because each pointpierces through conventional resist coatings andsolder residues.

Q1. The Huntron Tracker 1000 and 2000 are for use oncircuit boards and systems with all voltage sourcesin what condition?

Q2. What mode on the Automatic Transistor Analyzer

Model 900 has the Sonalert™? Q3. What type signal display does the Huntron

Tracker 1000 and 2000 show when the signal fuseis open and the test leads shorted together?

Q4. When using the Huntron Tracker 2000, why must

you make good contact with the test leads?

Q5. What is the minimum total shunt impedanceacross the junction of the diode or transistor undertest using the Automatic Transistor AnalyzerModel 900 to ensure a good test reading?

JET IGNITION SYSTEM TESTER

The jet ignition system tester detects and isolatesfaults in the jet engine ignition system. You can usethe tester to make the following checks:

• Operational check of the ignition system. Youare able to check the engine through theignition unit to the spark plug.

• Operational check of the spark plugs.

• Check of the ignition unit output in sparks

per second.

• Power input to the ignition unit.

The panel of the ignition system tester is shownin figure 2-32. If you want the procedures for usingthe tester, refer to the tester’s operation and serviceinstruction manual and the engine ignition systemmaintenance manual.

Figure 2-32.-Jet ignition system test panel.

Figure 2-33.-Tachometer Indicator-Generator Test Set TTU-27/E.

TACHOMETER INDICATOR-GENERATOR TEST SET TTU-27/E

The TTU-27/E has complete facilities for testingthe percent and RPM types of tachometer indicators.You can also use it to test four-pole reciprocatingengine tachometer generators and two-pole jet enginetachometer generators. You can use the tachometerindicator to make the following tests:

• Test tachometer indicators for startingvoltage and calibration accuracy

• Test tachometer generators for starting

voltage, calibration accuracy, speed, andoutput voltage under load conditions

• Test tachometer generators and indicators

either on or off the aircraft

Figure 2-33 shows a TTU-27/E tester. All theoperating controls, switches, and indicators required

for the tester are on the panel assembly. The two-speed test pad accommodates either two-pole or four-pole tachometer generators for testing. The RPMcounter indicator provides a precision readout oftachometer generator speed with scales of RPM x 2,RPM x 1, and PERCENT RPM. The tester uses 115-volt, 400-hertz ac power.

NOTE: Don’t try to operate the tachometertester until after you make the ground connection ofthe power lead.

The drive assembly consists of a gearbox and amaster tachometer generator. The gearbox drives thetwo-speed test pad, which mounts the tachometergenerator under test. A variable dc drive motorcontrols the drive speed of the gearbox. Therefore, thedrive motor controls the drive speed of the mastertachometer and the tachometer generator under test.

The RPM indicator assembly is part of theprecision speed indicator. The assembly consists of atwo-phase servomotor, which drives an

in-line digital display. There are three counterranges—0 to 10,000, 0 to 5,000, and 1 to119.4—for the respective scales. The componentsof the RPM indicator are inside a solder-sealedmetal container to prevent dirt from entering.

You can check indicator calibration by themaster tachometer generator. Compare indicatorRPM or percent RPM indication to the readingon the RPM indicator of the tachometer tester.

For additional information on the operationand service instruction of the TTU-27/E tester,refer to NAVWEPS 17-15CM-2.

JET CALIBRATION (JETCAL)ANALYZER

Of the many factors affecting jet engine life,efficiency, and safe operation, the most important

are exhaust gas temperature and engine speed. Ifthe exhaust gas temperature is just a few degreeshigher than it should be, the turbine blade life isreduced as much as 50 percent. Abnormally highexhaust gas temperature resulting from excessengine speed can cause premature engine failure.If the exhaust gas temperature is too low, jetengine efficiency and thrust is reduced. Either ofthese two conditions makes engine operationdangerous. By using the JETCAL analyzer, youcan detect or prevent these conditions.

You can more accurately check signs of fuelsystem trouble, tailpipe temperature, and RPMwith the JETCAL analyzer than with aircraftgauges. With proper use of the JETCAL analyzer,you can detect malfunctions in these systems.

The JETCAL analyzer (fig. 2-34) is a rugged,portable instrument made of aluminum, stainless

Figure 2-34.-JETCAL analyzer.

steel, and plastic. The major components of theanalyzer are the thermocouples, RPM and EGTindicators, resistance and insulation check circuits,and the overheat detection test circuits.

On the EGT system functional test and thethermocouple and harness checks, the JETCALanalyzer has an accuracy of ± 4°C at the testtemperature. The test temperature is usually themaximum operating temperature of the jetengine. To find the maximum engine operatingtemperatures, you should check the applicableMIMs.

The first test you perform is a functional testof the EGT system. In this test, you heat theengine thermocouples in the tail cone of the engineto test temperature. This heat comes from theJETCAL heater probes through the necessarycable connections. To measure the temperatureof the heater probes, thermocouples are embeddedin the probes. You may read the heater probetemperature on the JETCAL potentiometer. Atthe same time, you can read the aircraftthermocouple temperature on the aircraft EGTindicator. You can then compare the readings foraccuracy.

On engines that have a balancing type ofthermocouple system, you must remove thebalancing thermocouple from the circuit. Then,you can check the remaining thermocouplesindividually or together. Check the balancingthermocouple using a single probe. Read theoutput of the balancing thermocouple on theJETCAL potentiometer, and compare it to theheater probe thermocouple reading.

Gas temperature is critical to engine operation;therefore, use the JETCAL for nozzle scheduling.During temperature adjustments, make alltemperature readings on the JETCAL potenti-ometer. This is necessary because you must readengine temperature accurately to ensure the engineis operating at optimum engine conditions. Whenchecking and adjusting the engine exhaust gastemperature, install a switch box in the EGTcircuit before starting the test. You will use theswitch box to switch the cockpit indicator into thecircuit, or to switch the temperature indication ofthe engine thermocouple and harness to theJETCAL potentiometer. You can also readtemperature readings from the engine thermo-couple and harness on the JETCAL analyzer bymaking the necessary connections.

Incorporated in the analyzer is the TAKCALunit (RPM indicator) check circuit. This circuitreads engine speed with an accuracy of ±0.1percent. The TAKCAL unit check can also be

used to troubleshoot the aircraft’s tachometersystem. After testing the exhaust gas temperatureand engine speed systems, you may use selectedportions of the JETCAL analyzer circuits toestablish the proper relationship between exhaustgas temperature and engine speed.

The JETCAL analyzer requires a powersupply of 95 to 135 volts, 50 to 400 hertz. It willoperate in temperatures from –55°C to +70°C.You must use a 95- to 135-volt ac powersupply for the TAKCAL unit check and thethermocouple check. You may perform all otheroperations using emergency batteries when anexternal power supply is not available. Youactuate the batteries by a push-button switch.When the switch is released, the batteries are outof the circuit. However, you can depress theswitch when using ac without damaging thebatteries. To preserve the life of the emergencybatteries, use an external ac power supplywhenever possible.

The JETCAL analyzer has the followingprimary and separate functions:

• To check the entire jet aircraft exhaust gastemperature system for error without runningthe engine or disconnecting the wiring.

• To check individual thermocouples beforeplacing them in the aircraft.

• To check each engine thermocouple forcontinuity.

• To check the thermocouples and harnessfor accuracy of output.

• To check the resistance of the EGT circuit,without the EGT indicator, to assure al-lowable limits.

• To check the insulation of the EGT circuitfor shorts or grounds.

• To check the EGT indicators.

• To check engine thermocouples andharness on the engine with the engineremoved from the aircraft.

• To read engine RPM to an accuracy of±0.1-percent engine runup.

• To use the RPM check (TAKCAL) andpotentiometer to establish the proper rela-

tionship between exhaust gas temperature andengine speed during tabbing. (Tabbingis the procedure to adjust fixed or variableexhaust gas tail cone areas during normalaircraft checks every 30 to 50 hours.)

• To check aircraft fire detector, overheatdetector, and wing anti-icing systemsby using tempcal probes.

Refer to the MIM of a particular aircraft forthe proper procedure when you check and adjustthe variable nozzle systems. You can get detailedinstructions for adjusting the EGT with theJETCAL analyzer from the applicable MIM.

Which piece of test equipment would youuse to test an engine ignition unit outputin sparks per second?

Before using the TTU-27/E test set, whatconnection must you make to ensure properoperation?

What is the minimum voltage for using theTAKCAL unit of the JETCAL Analyzer?

SYNCHROPHASER TEST SET

The synchrophaser test set (fig. 2-35) testspropeller synchrophaser electronic units. The testset generates all pulses (dc and ac) required to testthe synchrophaser electronic unit independent ofits associated components.

The test set is completely solid state. Eachswitch position programs a particular test. It doesthis by connecting the appropriate inputs to thesynchrophaser and the meter required to verifysynchrophaser performance.

The test set generates a master pulse and aslave pulse. The pulse circuits closely simulate thesynchrophaser pulse input under dynamic flightconditions. You program different phase andspeed relationships of master and slave pulses,under test specifications, by setting the selectorswitches for a particular test.

Figure 2-35.-Synchrophaser test set.

The output of synchrophaser-associated com-ponents is simulated by the test set. It providesa simulated tachometer signal and the resistanceof synchrophaser controls for certain tests. Whenprogrammed, these signals help measure thesynchrophaser dynamic response to off-speed,off-phase, speed reset, resync, and throttleanticipation signals.

Control adjustments are minimal. Only a fewof the automatically programmed tests requireadjustment of either of the feedback gainpotentiometers. The accuracy of control adjust-ment is optimized by designing test set circuits thatrequire few adjustments. This adjustments will beto null or zero settings on center-scale zero meters.

A high degree of accuracy and repeatabilityis a feature of the gain measuring circuits withinthe tester. Conventional gain circuits requireapplication and measurement of incrementalvoltages to the synchrophaser for comparison withresultant output voltages. The gain measurementcircuit has a galvanometers, demodulator, dcpower supply, and calibrated potentiometer in abridge circuit. When the gain test is programmed,rotate the calibrated potentiometer to null thegalvanometers. With the galvanometers nulled, thepotentiometer calibration gives a direct readoutof amplifier gain. This circuit avoids inaccuracies

of input settings and error amplification ofincremental gain comparison. It also avoidsthe chance of error in calculating the gainfactor.

For complete instructions on this tester, youshould refer to the current Operation and ServiceInstruction Manual, NA 17-15CFA-2.

PRESSURE-TEMPERATURE TESTSET TTU-205C/E

The pressure-temperature test set (fig. 2-36)provides regulated pitot and static pressure forchecking performance characteristics of aircraftpneumatic instruments, air data systems, andother auxiliary equipment. You can also conductdynamic tests, quantitative calibration tests,pneumatic-system leak tests, and total tempera-ture probe tests using this tester.

The TTU-205C/E is a compact assembly. Itconsists of a control panel, components assembly,and combination carrying case with removablecover. All the controls, indicators, switches,electrical, and pneumatic connectors are on the

front panel of the test set. All the components areon the underside of the control panel. Thismakes the subassemblies and components easilyaccessible when the panel is not in it’s carryingcase. The test set simulates airspeed and altitudeinformation, which is displayed on the frontpanel. See table 2-6 for test set particulars.

The test set has the following five capabilities:

1. It performs overall system checks on thefollowing equipment, pneumatic systems, air datasystems, flight instruments, and pneumaticancillary equipment. This equipment may beeither installed in or removed from the air-craft.

2. It provides simulated total temperature forthe USAF MA-1 Probe, a platinum resistancesensing element whose resistance is 50.0 ohms at0°C.

3. It connects directly into the aircraft’spneumatic system to measure and control staticpressure (Ps) and pitot pressure (PT).

4. It accepts electrical power from an externalsource.

Figure 2-36.-Test Set, TTU-205C/E.

Table 2-6.-Pressure-Temperature Test Set TTU 205C/E

5. It incorporates safety features that prevent CAPACITIVE-TYPE LIQUIDdamage to the test set and the unit under test. QUANTITY TEST SET TF-20

CAUTIONMany types of capacitive fuel quantity testers

are in use in naval aviation. All operate onThe power requirements for this tester are the same basic principle—that of a variable115 volts, single phase, 400 Hz. For oncapacitor. Since it is impractical to describebench or hangar use, make sure that the all of the various testers, only the TF-20 iscorrect frequency of 400 Hz is available. discussed. You can use this set to test most of theThe tester will not work on 115-volt, capacitive fuel quantity systems currently used on60-hertz power. naval aircraft.

The Model TF-20 liquid quantity test set(fig. 2-37) contains circuitry that makes the testercapable of measuring tank unit probe capacitanceand measuring tank unit insulation resistance.It also simulates total probe capacitance forchecking aircraft fuel quantity indicators andcalibrating the test set capacitance indicator dialand megohmmeter scales.

You may test the fuel-gauging system eitherinstalled in the aircraft or removed from the air-craft. The test set includes a transit case, a shock-mounted instrument case, and a set of adaptercables. The tester has circuitry for measuringcapacitance and dc resistance. It also simulatesthe capacitance of compensated and uncompen-sated-type fuel tank probes. The instrumentsoperate with 115-volt, 400-hertz power.

All operating controls are on the front panel.Three BNC-type receptacles (COMP, COAX, andUNSH) are located on the upper middle part of

the panel. They provide connections to the tankunit probe under test. A pair of binding posts,(EXT RES) connect the tester to unknown highexternal resistances. Use these bonding posts whenyou can’t connect the tester to the TANK UNITreceptacles with the accessory cables. Another setof three BNC-type receptacles (COAX, UNSH, andCOMP) are located on the bottom right of the panel.These three receptacles let you connect the testset to the aircraft fuel-quantity-gauge indicator.

All adjustment controls for tester self-calibration are accessible on the front panel. TheZERO ADJ and HIGH ADJ controls provideadjustment for zero and the high end of the scaleon the CAPACITANCE indicator. The ZEROADJ and MIDSCALE ADJ controls adjust theMEGOHMS indicator so its pointer will positionat zero and at midscale. These four adjustmentshave covers so the calibrating controls will not beaccidently disturbed when using the test set.

Figure 2-37.-TF-20 liquid quantity test set.

When using this test set, you should followthe instructions in the applicable MIM for thesystem under test.

ANGLE-OF-ATTACK TEST SETAN/PSM-17A

The angle-of-attack test set lets you test,adjust, calibrate, simulate, and monitor the angle-of-attack indicating system. The test set alsoprovides a means for you to test the aircraftapproach lights, cockpit index lights, and the stallwarning system.

The test set (fig. 2-38) consists of a controlpanel enclosed in a case. It also includes the cablesand components to interconnect and test theangle-of-attack system and associated componentswithout their removal from the aircraft. Thecontrol panel contains a microammeter, adifferential pressure gauge, a potentiometercontrol, a bellows assembly, indicator lamps, andelectrical connectors. Also, the panel has varioustoggle and rotary switches to select and controlcircuits that are within the test set.

A radiometer system makes up the largestportion of the test set. This system consists of thenullmeter, the SIM OR NULL potentiometercontrol, and the NULL switch. It lets you test theangle-of-attack transmitter by placing switches invarious positions and using the potentiometer dialto simulate known inputs to the indicators.

Figure 2-38.-Angle-of-Attack Test Set AN/PSM-17A.

The pressure system consists of a differentialpressure gauge, test pressure connector, bellowsassembly, pressure control, and surge chamber(test set case). It dynamically tests the angle-of-attack transmitters. An air pressure or vacuumtransmits through a hose to parts of the AOAtransmitter probe by positioning the control onthe bellows assembly. This slight pressure causesthe probe to rotate. Thus, the pressure system ofthe test set simulates conditions corresponding tovarious aircraft angles of attack.

A series of indicator lamps (three indexer,three approach, and a stall warning) are on thetest set control panel. They simulate the actionof the aircraft indexer and approach lights andthe stall warning vibrator. Two additional lampsshow when the test set has ac and dc power.

The power requirements for the test set are28-volt dc and 110-to 120-volt, 400 hertz, single-phase ac.

AIR-CONDITIONING TEST SETAN/PSM-21A

The PSM-21A (fig. 2-39) is for flight linecheckout and troubleshooting of electrical

Figure 2-39.-Air-conditioning Test Set AN/PSM-21A.

components in the cabin, pressure suit, andequipment air-conditioning systems. To accom-plish checkout of these systems, apply externalelectrical power to the aircraft.

Checkout of the system under test involvessimulation of sensor and limiter inputs by the testset. You can connect external test equipment totest points on the test set to measure resistance,voltage, or waveforms to determine systemoperation. You can also determine system opera-tion by visual monitoring. For example, with aknown electrical input into the air-conditioningsystem, air-conditioning valves should move toa known position.

What test set is used to test pitot-staticsystems?

Name the two measurements you can takewith the TF-20 test set.

The AN/PSM-17A test set is used to testwhat system?

VERSATILE AVIONICS SHOP TEST(VAST) EQUIPMENT

The VAST system now in use aboard Navyaircraft carriers can cope with the continuallychanging character of avionics testing. VAST willsignificantly reduce the space required byequivalent special and manual support equipment.

In its basic form, a VAST station is assembledfrom an inventory of functional building blocks.These blocks furnish all the necessary stimuli andmeasurement capability to check many navalavionics equipments. As new avionics come intouse, test station configurations are modified byadding new building blocks to furnish newparameters or greater precision to existingcapabilities.

A typical VAST station consists of a computersubsystem, a data transfer unit (DTU), and astimulus and measurement section. The stimulusand measurement section contains functionalbuilding blocks configured to meet the intendedtest application. Figure 2-40 shows a typicalcarrier-based VAST station.

The carrier-based test station is controlled bya computer subsystem that executes test programsto assure accurate and satisfactory testing. Thecomputer subsystem includes a general-purposedigital computer, which not only executes testroutines but also provides diagnostic and com-putational capabilities. It also processes data and

Figure 2-40.-VAST station.

Figure 2-41.-Data transfer unit (DTU).

furnishes a permanent record of test results. Twomagnetic tape transports provide rapid access toavionics test programs and immediate availabilityof VAST self-check programs.

The data transfer unit (DTU) (fig. 2-41) servesas the operator-machine interface. It synchronizesinstructions and data flow between the computerand the functional building blocks. The DTU alsocontains the display and control panels.

The DTU control panel lets the operatorcommunicate with the computer. You will alsocommunicate with the stimulus and measurementsection of VAST by a keyboard and mode selectkey. You can operate the test station in manual,semiautomatic, or fully automatic modes.

The DTU contains a maintenance panelthat monitors station auto-check results andshows building-block faults. Transmission ofinstructions from the control computer is on arequest/acknowledge basis. Essentially, theresponse rate is controlled by the stimulus andmeasurement systems. This lets you transmitinstructions at an asynchronous rate correspond-ing to the maximum frequency at which a givenbuilding block or avionics unit can respond. Thereis no requirement for immediate program storagein the DTU.

VAST STATION

A VAST station may have as many as 14 racksof stimulus and measurement building blocks.(See fig. 2-42.) Large station configurations maycontain as many as 17 core building blocks.

Figure 2-42.-VAST station with building blocks.

VAST was designed with ease of maintenanceas a primary objective. In addition to themodularized design of VAST building blocks,there are three levels of fault detection that ensuresrapid confidence tests and easy fault location.The three levels of detection are auto-check,self-check, and self-test.

You may initially uncover a fault throughauto-check. The auto-check is inherent in the logicand control design of the VAST station. Itincludes verification of instructions and faultmonitoring. Auto-check works on a continuousbasis during station operation and, when a faultoccurs, interrupts testing.

The second level of VAST fault detection isself-check. This is a programmed sequence thatthe VAST operator starts through the DTUkeyboard. Self-check may be either internal or ata system level. Internal self-check measures theability of a building block to perform against itsown internal standards. System self-check requiresthe use of two or more building blocks in a testconfiguration selected to isolate faults within thetest set up.

The self-check philosophy you use to verifyVAST operation is based upon confirmation ofkey system elements first. It then uses theseelements to check the remaining building blocks.Checks of basic core building blocks are byinternal standards. Once satisfactory performanceis assured, the system uses their capabilities tocheck the remaining building blocks. The systemaccomplishes check-out of noncore buildingblocks by using any combination of core measure-ment or stimulus building blocks.

The final level of VAST fault detection isself-test. This is a series of test programs used tolocate faults within a building block. If a buildingblock contains a malfunction as a result of theself-check routine, remove the faulty buildingblock and connect to VAST as a unit under test.The system will then conduct self-test programs.

INTERCONNECTING DEVICE

In its simplest form, the interconnecting deviceconsists of an adapter cable, which connects theunit under test (UUT) to the VAST interface. Insome cases, passive and active circuits areintroduced to change impedance levels or toamplify low signals. These circuits are introducedas part of the electrical interface in the

interconnecting device. Ordinarily, this is notrequired if the avionics equipment has beendesigned with the requirements for VAST. Veryoften, you may obtain passive circuit functionsthrough the use of standard plug-in modules.

The last element of the test program is theinstruction booklet or microfilm strip. Theseinstructions give the details for all the steps tofollow in testing any given unit. The steps includeinitial procedures, such as hook up and clearingoperations, and proceed to the final steps ofdisconnect and UUT closeout.

OPERATION OF A VAST STATION

In a typical VAST procedure, ease in opera-tion of the actual testing becomes apparent. Youmay make the initial setup of the weaponreplaceable assembly (WRA), including removalof dust covers, cooling provisions, and con-nections to the interface device, off station tominimize disruptions of station operators. Finalconnections between the VAST stations interfacepanel and the UUT take only a few momentsdirectly at the station.

The operator begins testing by selecting thecode that starts the test program. Before applyingpower or stimulus to the UUT, you shouldconduct continuity tests to ensure proper testprogram selection. This will also ensure noconditions exist that will damage the VASTstation or the UUT once active tests start. Ifeverything checks out OK, the testing proceedsautomatically. The operator only has to respondto instructions that appear on the CRT display.The program will not stop until it encounters afault or it reaches a program halt.

The purpose of programmed halts is to allowmanual intervention during testing to makeadjustments and observations. When the identi-fication of faults and the operator’s instructionsare required—such as interpreting a complexwaveform—the operator may refer to the testprogram instructions. Upon completion of the testprogram, the CRT display shows any closeoutprocedures.

A VAST station is completely autonomousand normally operated under computer controlin a fully automatic mode. Of course, the operatorcan select any one of three semiautomatic modesor a manual mode.

The semiautomatic modes include a one-group, a one-test, and a one-step mode. Theseauxiliary modes permit detailed observation ofvarious test sequences. They are also useful in

performing work-around procedures in recon-ciling differences in equipment and program modestatus, and in the verification of repairs.

In the manual mode, the test station iscompletely off-line with respect to the computer.Instructions come from the operator through thekeyboard on a one-word-at-a-time basis. (Seefig. 2-43.) Although you won’t use the manualmode for avionics testing, it is useful fordebugging new programs. You can also integratenew building blocks into the station, and performself-check operations on some of the buildingblocks using this mode.

Figure 2-43.-VAST control panel.

ELECTROSTATIC DISCHARGE

Learning Objective: Recognize the hazardsto ESD-sensitive devices, to include properhandling and packaging techniques.

The sensitivity of electronic devices andcomponents to electrostatic discharge (ESD) hasrecently become clear through use, testing andfailure analysis. The construction and designfeatures of current microtechnology have resultedin devices being destroyed or damaged by ESDvoltages as low as 20 volts. The trend in thistechnology is toward greater complexity, increasedpackaging density, and thinner dielectrics betweenactive elements. This trend will result in deviceseven more sensitive to ESD.

Various devices and components are suscepti-ble to damage by electrostatic voltage levelscommonly generated in production, test, andoperation, and by maintenance personnel. Thesedevices and components include the following:

All microelectronic and most semicon-ductor devices, except for various powerdiodes and transistors

Thick and thin film resistors, chips andhybrid devices, and crystals

All subassemblies, assemblies, and equipmentcontaining these components/devices withoutadequate protective circuitry are ESD sensitive.

You can protect ESDS items by implementingsimple, low-cost ESD controls. Lack of imple-mentation has resulted in high repair costs,excessive equipment downtime, and reducedequipment effectiveness.

The operational characteristics of a systemmay not normally show these failures. However,under internal built-in-test monitoring in adigital application, they become pronounced, Forexample, the system functions normally on theground; but, when placed in an operationalenvironment, a damaged PN junction mightfurther degrade, causing its failure. Normalexamination of these parts will not detect thedamage unless you use a curve tracer to measurethe signal rise and fall times, or check the partsfor reverse leakage current.

STATIC ELECTRICITY

Static electricity is electrical energy at rest.Some substances readily give up electrons while

others accumulate excessive electrons. When twosubstances are rubbed together, separated or flowrelative to one another (such as a gas or liquidover a solid), one substance becomes negativelycharged and the other positively charged. Anelectrostatic field or lines of force emanatebetween a charged object to an object at adifferent electrostatic potential (such as more orless electrons) or ground. Objects entering thisfield will receive a charge by induction.

The capacitance of the charged object relativeto another object or ground also has an effect onthe field. If the capacitance is reduced, there isan inverse linear increase in voltage, since thecharge must be conserved. As the capacitancedecreases, the voltage increases until a dischargeoccurs via an arc.

CAUSES OF STATIC ELECTRICITY

Generation of static electricity on an objectby rubbing is known as the triboelectric effect.Table 2-7 lists substances in the triboelectric series.

The size of an electrostatic charge on twodifferent materials is proportional to the separa-tion of the two materials. Typical prime chargegenerators commonly encountered in a manu-facturing facility are shown in table 2-8.

Electrostatic voltage levels generated bynonconductors can be extremely high. However,air will slowly dissipate the charge to a nearbyconductor or ground. The more moisture in theair the faster a charge will dissipate. Table 2-9shows typical measured charges generated bypersonnel in a manufacturing facility. Note thedecrease in generated voltage with the increase inhumidity levels of the surrounding air.

NOTE: The triboelectric series is arranged inan order that when any two substances in the listcontact one another and are separated, thesubstance higher on the list assumes a positivecharge.

EFFECTS OF STATIC ELECTRICITY

The effects of ESD are not recognized.Failures due to ESD are often misanalyzed asbeing caused by electrical overstress due totransients other than static. Many failures, oftenclassified as other, random, unknown, infantmortality, manufacturing defect, etc., are actuallycaused by ESD. Misclassification of the defect isoften caused by not performing failure analysisto the proper depth.

Table 2-7.-Triboelectric Series

COMPONENT SUSCEPTIBILITY

All solid-state devices (all microcircuits andmost semiconductors), except for various powertransistors and diodes, are susceptible to damageby discharging electrostatic voltages. Thedischarge may occur across their terminals orthrough subjecting these devices to electrostaticfields.

Table 2-8.-Typical Charge Generators

Table 2-9.-Typical Measured Electrostatic Voltages

LATENT FAILURE MECHANISMS

The ESD overstress can produce a dielectricbreakdown of a self-healing nature when thecurrent is unlimited. When this occurs, the devicemay retest good but contain a hole in the gateoxide. With use, metal will eventually migratethrough the puncture, resulting in a shorting ofthis oxide layer.

Another structure mechanism involves highlylimited current dielectric breakdown from whichno apparent damage is done. However, thisreduces the voltage at which subsequent break-down occurs to as low as one-third of the originalbreakdown value. ESD damage can result in alowered damage threshold at which a subsequentlower voltage ESD will cause further degradationor a functional failure.

ESD ELIMINATION

The heart of an ESD control program is theESD-protected work area and ESD groundedwork station. When you handle an ESD-sensitive(ESDS) device outside of its ESD protectivepackaging, provide a means to reduce generatedelectrostatic voltages below the levels at which theitem is sensitive. The greater the margin betweenthe level at which the generated voltages arelimited and the ESDS item sensitivity level, thegreater the probability of protecting that item.

PRIME GENERATORS

Look at table 2-8. It lists ESD primegenerators. All common plastics and other primegenerators of static electricity should be prohibitedin the ESD protected work area. Carpeting shouldalso be prohibited. If you must use carpet, itshould be of a permanently anti-static type.Perform weekly static voltage monitoring wherecarpeting is in use.

CAUTION

Anti-static cushioning material isacceptable; however, the items cited shallbe of conductive material to preventdamage or destruction of ESDS devices.

PERSONAL APPAREL ANDGROUNDING

An essential part of the ESD program isgrounding personnel and their apparel when

handling ESDS material. Means of doing this aredescribed in this section.

Smocks

Personnel handling ESDS items should wearlong sleeve ESD protective smocks, short sleeveshirts or blouses, and ESD protective gauntletsbanded to the bare wrist and extending towardthe elbow. If these items are not available, useother anti-static material (such as cotton) that willcover sections of the body that could contact anESDS item during handling.

Personnel Ground Straps

Personnel ground straps should have aminimum resistance of 250,000 ohms. Based uponlimiting leakage currents to personnel to 5milliamperes, this resistance will protect personnelfrom shock from voltages up to 125 volts RMS.The wrist, leg, or ankle bracelet end of the groundstrap should have some metal contact with theskin. Bracelets made completely of carbon-impregnated plastic may burnish around the areain contact with the skin, resulting in too high animpedance to ground.

ESD PROTECTIVE MATERIALS

There are two basic types of ESD protectivematerials—conductive and anti-static. Conductivematerials protect ESD devices from staticdischarges and electromagnetic fields. Anti-staticmaterial is a nonstatic generating material. Otherthan not generating static, anti-static materialoffers no other protection to an ESD device.

CONDUCTIVE ESD PROTECTIVEMATERIALS

Conductive ESD protective materials consistof metal, metal-coated, and metal-impregnatedmaterials (such as carbon particle impregnated,conductive mesh or wire encased in plastic). Themost common conductive materials used for ESDprotection are steel, aluminum, and carbon-impregnated polyethylene and nylon. The lattertwo are opaque, black, flexible, heat sealable,electrically conductive plastics. These plastics arecomposed of carbon particles, impregnated in theplastic, which provides volume conductivitythroughout the material.

ANTI-STATIC ESD PROTECTIVEMATERIALS

Anti-static materials are normally plastic typematerials (such as polyethylene, polyolefin,polyurethane, nylon), which are impregnated withan anti-static substance. This anti-static migratesto the surface, and combines with the humidityin the air to form a conductive sweat layer on thesurface. This layer is invisible and, althoughhighly resistive, is amply conductive to preventthe buildup of electrostatic charges by triboelectric(or rubbing) methods in normal handling. Simplystated, the primary asset of an anti-static materialis that it will not generate a charge on its surface.However, this material won’t protect an enclosedESD device if it comes into contact with a chargedsurface.

This material is of a pink tint, a symbol of itsbeing anti-static. Anti-static materials are forinner-wrap packaging. However, anti-static trays,vials, carriers, boxes, etc., are not used unlesscomponents and/or assemblies are wrapped inconductive packaging.

HYBRID ESD PROTECTIVE BAGS

Lamination of different ESD protectivematerial (that is, conductive and anti-static) isavailable that provides the advantage of both ina single bag.

ESDS DEVICE HANDLING

The following are general guidelines applicableto the handling of ESDS devices:

• Make sure that all containers, tools, testequipment, and fixtures used in ESDprotective areas are grounded beforeand/or during use, either directly or bycontact with a grounded surface.

• Personnel handling ESDS items mustavoid physical activities in the vicinity ofESDS items that are friction-producing, forexample, removing or putting on smocks, wipingfeet, sliding objects over surfaces, etc.

• Personnel handling ESDS items must wearcotton smocks and/or other anti-staticallytreated clothing.

• Avoid the use or presence of plastics,synthetic textiles, rubber, finished wood,

vinyls, and other static-generating materials(table 2-9) where ESDS items are handledout of their ESD protective packaging.

• Place the ESD protective materialcontaining the ESD item on a groundedwork bench surface to remove any chargebefore opening the packaging material.

• Personnel must attach personnel groundingstraps to ground themselves before removingESDS items from their protective packaging.

• Remove ESDS items from ESD protectivepackaging with fingers or metal grasping toolonly after grounding and place on the ESDgrounded work bench surface.

• Make periodic electrostatic measurementsat all ESD protected areas. This assures the ESDprotective properties of the work station and allequipment contained there have not degraded.

• Perform periodic continuity checks ofpersonnel ground straps (between skin contactand ground connection), ESD grounded workstation surfaces, conductive floor mats,and other connections to ground. Performthis check with a megohmmeter to make suregrounding resistivity requirements are met.

ESDS DEVICE PACKAGING

Before an ESDS item leaves an ESD protectedarea, package the item in one of the followingESD protective materials:

• Ensure shorting bars, clips, or non-corrective conductive materials are correctlyinserted in or on all terminals or connectors.

• Package ESD items in an inner wrap, oftype II material conforming to MIL-B-81705,and an outer wrap of type I material con-forming to MIL-B-81705. You may use alaminated bag in instead of the above pro-vided such meets the requirements of M-B-81705.Cushion-wrap the item with electrostatic-free material conforming to PPP-C-1842,type III, style A. Place the cushioned item intoa barrier bag fabricated from MIL-C-131 andheat-seal closed, method 1A-8. Place the wrapped,cushioned, or pouched ESDS item in bags con-forming to MIL-B-117, type I, class F, style 1.

Mark the packaged unit with the ESD symboland caution (fig. 2-44).

TESTING/REPAIR

Before you work on ESDS items, make surethe following precautions/procedures are met.

• Be sure that work area, equipment, andwrist strap assembly have a proper ground.

• Attach wrist strap and place metal tools,card extractors, test fixtures, etc., on groundedbench surface.

• Place conductive container on the benchtop. Remove component/assembly from packag-ing. Remove shorting devices, if present. Handlecomponents by their bodies and lay them onconductive work surface or test fixtures.

• Test through the connector or tabs only.

• Do not probe assemblies with testequipment.

• After testing, replace shorting devices andprotective packaging.

• Use of Simpson Model 260 or equivalentto test parts or assemblies is prohibited. Youmust use a high input impedance meter suchas a Fluke 8000A Multimeter.

• Dielectric strength tests are not permitted.

Figure 2-44.-ESDS markings.

CHAPTER 3

POWER GENERATION ANDCONTROL SYSTEMS

This chapter has been deleted. For information on power generation and distribution, refer to Nonresident Training Course (NRTC) Aviation Electricity and Electronics—Power Generation and Distribution, NAVEDTRA 14323.

CHAPTER 4

AIRCRAFT ELECTRICAL SYSTEMS

As an Aviation Electrician (AE), you will workwith many electrical systems. You must be awell-rounded aviation technician. You may workdirectly or indirectly with all other aviationmaintenance ratings. You may work with Avia-tion Machinist Mates (ADs) on power plantdiscrepancies, or with Aviation StructuralMechanics (AMs) on electrohydraulic malfunc-tions.

In this chapter, you will learn about varioussystems that AEs regularly maintain. Maintainingnaval aircraft systems is the number one priority;teamwork and maximum effort will pay dividendsin flight safety and mission accomplishment.

AIRCRAFT LIGHTING SYSTEMS

Learning Objective: Recognize operatingprinciples and construction features ofaircraft lighting, both internal andexternal, and identify types of lamps usedin these circuits.

The lighting system in an aircraft serves twoprimary purposes—it provides specialized lightsources outside the aircraft and illuminates theinterior. Exterior lights provide illumination atnight for navigation and formation flying. Otherexterior light operations include, signaling,landing, and anticollision. The interior lightingprovides illumination for instruments, equipment,cockpits, and cabins. In addition, the electrical,electronic, and mechanical systems use lightsto indicate normal operation or a possiblemalfunction. Lights also show the position of thelanding gear, bomb bay doors, sonobuoy exitdoors, etc.

Various types and sizes of light assemblies areused on present-day naval aircraft. The lightingrequirement governs the selection of a particularlight assembly. Most light assemblies consists ofa housing (fixture), a lamp, and a lens.

DESCRIPTION AND TYPESOF LAMPS

Aircraft lamps are devices that provide sourcesof artificial light. The incandescent light is themost common type. It uses an electrical sourceto heat a filament until it is white hot. Normally,the source voltage is 26 to 28 volts ac or dc.However, some lighting systems use lower voltagelamps and step-down transformers to supply 3 to6 volts to the lamps. The lower operating voltageallows the lamp filament to be larger, and it helpsto reduce lamp failure due to vibrations.

When a lamp fails, the replacement must bethe same as the original lamp or an approvedalternate. Spare lamps in the aircraft are shock-mounted. If they weren’t, they would probablyfail earlier than the original lamps because coldfilaments are subject to fatigue failure sooner thanhot ones. You need to be sure that the glass bulbof the lamp is clear and free from grease and dirt.To help keep bulbs clean, don’t touch the glassbulb with your bare hands, if possible.

The parts of a lamp are the bulb, filament,and base. Incandescent lamps vary chiefly inelectrical rating, base type, bulb shape, and bulbfinish.

The electrical rating of lamps is expressed asa combination of volts, watts, amperes, andcandlepower. (Candlepower is the luminousintensity expressed in candles and used to specifythe strength of a light source.) The lamp ratingis found on either the base or the bulb of the lamp.On small lamps, an identifying number representsthe electrical rating. This number is the same asthe military specification (MS) dash number. Youcan find it on the lamp base.

The base types vary as to size, number ofelectrical contacts, and the method of securing ina socket. The most common types of bases arethe single- or double-contact bayonet (push in andturn). These bases are used on aircraft since theylock in the socket and do not loosen because ofvibration. Single contact bases are used in

single-wire systems. In this system, one side of thelamp filament connects to the base and the otherto the contact. Double contact bases are used ifthe single-wire system is not practicable. In thissystem, the filament connects to contacts and doesnot make an electrical connection to the baseunless the lamp has dual filaments. Dual filamentsuse a common contact to the base. The single- anddouble-contact lamps are not interchangeable.

Some bases are of the screw type. They aren’tused often because they loosen easily. Other lampshave an indexing-type base that has offset indexpins. (See figure 4-1.) The index-type base is usedto make sure that the lamp sits in the socket andthe light shines in the proper direction.

Some lamps do not have a base; they solderdirectly onto a circuit board or permanentlymount in control box panels. These lamps gothrough careful testing and selection so they lastthe life of the aircraft. Figure 4-2 shows one suchlamp, the grain of wheat lamp.

The bulb shape is shown by a combination ofletters and numerals. The letter shows the bulbshape, and the number shows the approximatemaximum diameter of the bulb, in eighths of aninch. The codes and the shapes they indicate forthe more common glass bulbs are as follows:

G Globular

GG Grimes globular

S Straight

T Tubular

PAR Parabolic aluminum reflector

R Reflector

By looking at the letterknow that a bulb designated

designations, youas T-6 is a tubular

Figure 4-2.-Grain of wheat lamp.

bulb with a 6/8-inch diameter. A variety of sizesand shapes of bulbs are listed in the DefenseLogistics Agency (DLA) Identification List. Thislist is available through your local supply supportcenter. Figure 4-3 shows some common bulbshapes.

Most aviation lamps are either clear glass orfrosted on the inside. For a particular application,however, a bulb may be partially frosted; for

Figure 4-3.-Common bulb shapes.

Figure 4-1.-Lamp bases.

example, to cut down emission of light in aparticular direction. Another bulb may bepartially silvered to prevent emission in a specifieddirection and/or to concentrate the light in otherdirections. Some applications call for coloredbulbs; for example, in instrument illumination andsafety lights. The letters just before the militaryspecification (MS) dash number shows bulbfinish-R for red, SB for silvered bowl. If noletter is present, the lamp is clear glass or frosted.You can also provide colored lighting by usingclear lamps with a colored lens cover.

There are many special-purpose lamps in useon naval aircraft. Three of the most commontypes are listed below:

1. The parabolic, sealed-beam landing andtaxi light. Signal lights also are used insealed-beam lamps.

The midget-flange type of light is for usein instrument panels and control boxes.Fuselage and signal lights use lamps havingtwo filaments in parallel to provide fastsignaling (smaller filaments heat and coolfaster).

EXTERIOR LIGHTING

Many types of lights are used to meet theexterior lighting requirements of naval aircraft.The principal types of exterior lights are thenavigation or position lights, anticollision lights,landing lights, and formation lights.

Figure 4-4 shows the components used in theexterior lighting system of a carrier-type aircraft.This figure shows the lights common to navalaircraft, but does not show every type oflight in use on different aircraft. The lighting

Figure 4-4.-Exterior lighting.

requirements vary from one aircraft to another,depending upon the aircraft type.

Navigation Lights (Position Lights)

Navigation lights on the aircraft attract visualattention to its position and heading at night. Astandard minimum set of navigation lights,meeting Federal Aviation Administration (FAA)requirements for light distribution and intensity,is on all military heavier-than-air aircraft.

The standard minimum set of navigation lightsfor night operations consists of the following:

1. One red light on the tip of the left wing2. One green light on the tip of the right wing

3. One white light on the tail, so it is visibleover a wide angle from the rear

NOTE: The position for these navigationlights is on a stationary surface to the extreme left,right, and aft on helicopters (fig. 4-5).

In some aircraft configurations, navigationlights burn steadily; in other configurations theyburn steadily or flash about 80 flashes per minute.

Fuselage Lights

Fuselage signal lights are part of the aircraftnavigation lighting system. They provide amethod of visual signaling. When installed, twoor more fuselage lights are necessary, one on thetop and one on the bottom of the aircraft. If itis not practicable to install the light on the bottomof the aircraft, such as a seaplane or a radomeobstruction, you install two lights. Position oneon each side as near the bottom of the aircraftas possible. Fuselage lights may burn steadily orflash at a constant rate. On some aircraft, manualkeying of lights for signaling is available.

Anticollision Lights

Anticollision beacon lights are an FAArequirement for all aircraft. Their primarypurpose is flight safety, during daylight hours aswell as night.

One type of anticollision light consists of two40-watt reflector-type lights and a red lensassembly. An ac or dc motor rotates the bulbassembly, causing it to flash 80 to 90 flashes perminute. Slip rings provide electrical power to thebulbs.

On another type of anticollision light, thebulb is stationary. The flashing is caused bymotor-driven reflectors in the light assembly.

Figure 4-5.-Typical helicopter lighting.

Most aircraft are now using strobe lights toprovide anticollision warning to other aircraft.This system has a 3,000 to 3,500 candlepowerwhite light for day and 150 to 200 candlepower redlight for night. Both flash at 60 flashes per minute.

Aircraft equipped with in-flight fueling tankercapabilities can turn off the lower anticollisionlight during delivery of fuel to another aircraft.Tanker aircraft use a bluish-green lens over theiranticollision lights to identify this capability toother aircraft in need of fuel.

Landing Lights

Modern naval aircraft have high candlepowerlanding lights to illuminate the landing strip.Multiengined aircraft usually have a landing lightmounted on each wing. Single-engine aircraft useonly one landing light, which normally mountson the port wing.

Landing lights are usually retractable (fig. 4-6).Sealed-beam lights with a rating of 28 volts, 600watts are used. They use split-field series dc

Figure 4-6.-Retractable landing lights.

motors or ac induction motors to provide driveoperation. The 28-volt ac is provided by anautotransformer located within each lightassembly. In the retracted position, movable lightsare flush with the undersurface of the wing. Insome aircraft, you can install a landing light inthe nosewheel fairing door. Other aircraft havea fixed landing light, mounted in the leading edgeof the wing.

As you read this section, refer to figure 4-6.You are going to learn about the operatingprinciples of a typical retractable landing light.

The landing light switch on the pilot’s lightingpanel controls the landing light motor. When theswitch is in either the EXTEND or RETRACTposition, power goes simultaneously to themagnetic brake (releasing the brake) and the drivemotor. Limit switches open the motor circuitwhen the light has reached its limit of travel ineither direction (extended or retracted position).This light can be put in any position between itstravel limits by turning the switch off. When thepower to the motor is off, the gear train holdsthe light in that position.

The lamp illuminates by a sliding contact afterthe light assembly travels downward about 10degrees. Illumination continues until the lampagain reaches the 10-degree position whentraveling in the upward direction. The lamp willremain lighted in any extended position past the10-degree position, regardless of power appli-cation to the control motor. The pilot can adjustthe angle of the light beam to fit the operation.A switch lets you turn the lamp off while the lightis in any position, as shown in figure 4-6, view B.

The landing light assembly design lets youadjust the maximum extended position for eachparticular aircraft installation. The light opens toan extended position of 73 degrees, ±3 degrees,from the retracted position. The light assemblyis capable of being extended to positions rangingfrom 50 degrees to 85 degrees from the retractedposition.

When you ground-test landing lights, don’t letthem overheat or be damaged by extending themagainst ground support equipment. You shouldNEVER look directly at an illuminated landinglight because it can cause permanent eye damage.

Approach Lights

The carrier aircraft approach light systems givethe pilot and landing safety officer (LSO) positiveindication of safe or unsafe landing configura-tions. All shipboard naval aircraft have approach

lights mounted so they are clearly visible to theLSO. Installations of the approach lights varywith aircraft. Older aircraft have them in the portwing leading edge. Modern aircraft, such as theF-14 (fig. 4-7) and F/A-18, have the approachlights on the nose landing gear door. Most jetaircraft have a three-lamp approach light assemblyconnected to the angle-of-attack indicator. Theapproach light control is by cam-actuated switchesin the angle-of-attack indicator. The lights operatewith landing gear completely down, with aircraftoff the gear, and arresting gear extended.

The three lights in the approach light assemblyare red, amber, and green. When the red light ison, it shows that the angle of attack of the aircraftis low. When the green light is on, it shows thatthe angle of attack of the aircraft is high. Whenthe amber light is on, the angle of attack is prefectfor the landing approach.

The approach lights work in conjunction withthe cockpit angle-of-attack indexer lights (fig.4-7), which give the pilot angle-of-attackinformation. These lights mount in the cockpiton either side of the windshield. They are clearlyvisible to the pilot, but don’t obstruct vision.

The indexer lights operate by cam-operatedswitches in the angle-of-attack indicator. At verylow angles of attack, a red-colored inverted Villuminates. This warns the pilot to increase theaircraft angle of attack. At slightly low angles ofattack, both the inverted V and circular symbol(doughnut) illuminate. At desired angles of attack,the amber-colored doughnut illuminates. Atslightly high angles of attack, both the V anddoughnut illuminate. At very high angles ofattack, a green-colored V symbol illuminates,warning the pilot to decrease the aircraft angleof attack. The indexer information correspondsto the approach lights assembly informationdisplayed to the LSO.

Approach lights show the LSO three things—(1) aircraft angle of attack, (2) landing gear isdown and locked, and (3) arresting gear extension.A limit switch in the arresting gear circuit operateswhen an unsafe condition exists, causing the lightsto flash.

An arresting gear (tailhook) override switchallows bypass of the tailhook circuit for carrierlanding practice at airfields. The override switchis of the momentary ON type. When in the ONposition, the arresting gear switch bypass relayenergizes. The approach lights operate only inconjunction with landing gear, since the arrestinggear is up. When the circuit de-energizes, itautomatically reverts to its normal condition

Figure 4-7.-Angle-of-attack components.

(arresting gear switch not bypassed). In most (port) light covers are red. The lamp assembly isaircraft, the circuit returns to normal by loweringthe arresting gear, turning off the battery orgenerator switch, or removing external power.

Formation Lights

Naval aircraft have formation lights for nightformation flying. In a typical formation lightinstallation, wingtip formation lights mountwithin the upper and lower surfaces of the aftsection of each wingtip. Translucent diffusingwindows mount flush with the wingtip surfaceabove and below each light. The right-hand(starboard) light covers are green and the left-hand

accessible through the lower wingtip cover plate.

Fuselage formation lights installations (fig.4-4) are in box assemblies on each side of thefuselage. Each has a plastic window for lightemission and access to the lamp is by removingthe box cover. These lights connect in parallel withthe wingtip formation lights; therefore, theyilluminate simultaneously.

NOTE: Manufacturers sometimes identifylights having the same purpose as formation lightsby another name such as join-up lights or striplights.

Taxi Lights

Taxi lights help the pilot maneuver the aircraftbefore and after flight. On aircraft having anosewheel, the taxi light assembly mounts on themovable strut so the light turns with the wheel.Light installation provides maximum visibility forthe pilot and co-pilot. (See figure 4-4 for a typicaltaxi light installation.)

In-Flight Refueling Probe Light

Most modern carrier aircraft have limitedrange due to their small fuel capacity. To increaserange and flight time, day and night in-flightrefueling is necessary. Most naval shipboardaircraft have provisions for in-flight refueling. Fornight refueling, in-flight refueling (IFR) probelights mount on the fuselage forward of the probe.The light lens is usually red in color andilluminates the refueling probe and drogue fromthe refueling aircraft at night. Figure 4-4 showsa typical IFR probe light installation.

Hover Lights and Spotlights

In helicopter installations, hover lightsilluminate the area directly beneath the aircraft(fig. 4-5, callout 3). These lights serve severaldifferent purposes, including landing and searchand rescue. A spotlight can be mounted in thenose of a helicopter (fig. 4-5, callout 7), and itcan extend or retract. An ON/OFF/RETRACTswitch on the pilot’s control stick and a spring-loaded, four-way thumb switch markedEXTEND/RETRACT/LEFT/RIGHT controlthe spotlight. It can rotate through 360 degreesof azimuth.

INTERIOR LIGHTING

Various types of lights and lighting systemsare used for interior illumination of naval aircraft.Almost all lights and lighting systems fall underone of the following types:

• Instrument lighting

• Cockpit lighting

• Cabin and passageway lighting

• Indicator lights

An important consideration in interior lightingof aircraft is the prevention of undue eyestrain.

Your eyes adjust slowly to changing lightintensities, which can cause fatigue or eyestrain.Aircraft lighting designs produce as littlediscomfort as possible. As an AE, you maintainthe lights. You should follow the specificationswhen replacing fixtures and lamps. You shouldobserve the following general considerations whenworking with the interior lighting of aircraft:

Use lenses to diffuse light that lies withinthe pilot’s field of vision.

Eliminate all bright spots of light, directsources of light, and reflections.

Use sparingly any surface that reflectslight, such as chromium or nickel.

Use quick-change lighting fixtures so thatlamps may be changed rapidly.

Instrument Lights

The first use of artificial light in aircraftwas for the instrument illumination. Operatingmodern aircraft depends on instruments; there-fore, instrument lighting is important. There aredifferent methods of instrument lighting, andwhich one is best makes a decision difficult. Nomatter what system of lighting is in the aircraft,the light must not be visible outside the aircraft.

Indirect (mask) lighting is desirable becauseit doesn’t produce objectionable reflections. Thelamps mount in the instrument panel. The panelhas a reflector cover (mask), which has openingsin it for observing the instruments. The lightreflects until it becomes diffused and floods theentire panel. Even though this system producessatisfactory lighting, it is not in common use todaybecause of space and weight requirements.

Figure 4-8.-Rheostat switch.

Figure 4-9.-Interior lighting.

Another method of instrument lighting(commonly called post lighting) uses installation ofspecially adapted shields in front of the instruments.Small lamps in red-filtered sockets mount in the frontsurface of the cover shields, which cast light down andonto the instruments. These shields mount around theinstrument for vision and are flanged to direct thelights properly for dial illumination.

Rheostats control the intensity of light in mostaircraft (fig. 4-8). The rheostat is a variable resistorused to limit current through the circuit. When thepivoting arm reaches the high-resistance end of therheostat, it slips off the contact surface and breaks thecircuit. This arrangement is a rheostat switch.

A typical aircraft instrument lighting system isshown in figure 4-9. You should refer to this

figure as you read this section. The lightingequipment consists of edge-lighted control panels,individual lights for instruments, and red flood-lights for overall lighting. The edge-lighted panelsprovide diffused lighting for the control andindicator panels.

Some instruments have a light source as anintegral part of the instrument, and all instrumentdesigns will soon have integral lighting. Sinceindividual lights for instruments are compatiblewith the integrally lighted instruments, the twotypes of light sources may be intermixed.

Edge-lighted control and indicator panels (fig.4-9) have a nongloss, black background with whitelettering for maximum ease of reading. Smalllamp assemblies are mounted so light is diffusedthrough the plastic panels. These lamps have redfilters to cut out glare. Bulb replacement isrelatively simple, Just unscrew the top cap fromthe assembly, pull the bulb from the cap, replacethe bulb, and reassemble the unit.

Most instrument lights consist of panel lightsand instrument integral lights. The essential buspowers the instrument lights through circuitbreakers or autotransformers through fuses.

The CONSOLE panel lights control is avariable intensity control for the console paneledge lights. As the control rotates clockwise, theintensity of the console edge lights increases. Also,as the control rotates to the OFF position, a switchwithin the control actuates, removing power fromthe console flood switch dim contact. Therefore,the console red floodlights operate in dim onlywhen the console panel lights control is in aposition other than OFF.

The CONSOLE FLOODS switch is a three-position switch (BRT/DIM/MED) for selectingthe intensity of console floodlights. Selection ofthe BRT or MED position energizes the lightsregardless of the position of the console panellights control. The WHITE FLOOD switch hastwo positions—OFF and ON. Placing the switchto ON energizes the white floodlights. TheNORMAL WARNING LIGHTS switch has twopositions—NORMAL and TEST—and it is springloaded to the NORMAL position. Placing theswitch in the momentary contact TEST positionilluminates all of the warning lights in the pilot’scockpit.

The INSTR PANEL EMERG FLOOD switch(fig. 4-9) has three positions–OFF, DIM, andBRT. The switch turns on the red floodlightsabove the main instrument panel in an emergencywhen normal instrument lighting malfunctions.

In some aircraft, tiny grain-of-wheat lampshave permanent mounts in the control-box panels.Having been aged and carefully selected foroutput and reliability, the lamps should last forthe lifetime of the aircraft.

Cavities and shallow trenches in the panel backaccommodate the panel connector, lamps, androute lamp leads to the panel connector. Theconnector, lamps, and leads are then potted intothe back of the panel with a clear plastic pottingmaterial. Normally, the lamps are in parallel, andfailure of one lamp will not appreciably degradethe panel lighting. The main advantages ofembedded lamps are long life, ruggedness,resistance to aircraft vibrations, and betterillumination.

Cockpit Lights

The term cockpit lighting is rather broad. Itsmeaning varies, depending on the type aircraftbeing described. In fighter-type aircraft, it maybe interior lighting that consists of individuallylighted instruments and switches, lighted controlpanels, and necessary floodlights. It includes theinterior lighting just mentioned as well as manyother special lighting assemblies.

A much used lighting device is the smallincandescent spotlight known as a utility lightassembly. These assemblies, installed at crewstations, are in a position where they provideillumination of equipment that the crew memberuses. The light from the lamp assembly can focusin either a small spot or in a wide beam. It alsohas a red filter for night use. Figure 4-10 showsone type of light in which an ON-OFF switch andan intensity control rheostat control lightoperation.

Cockpit extension light assemblies providecrew members with an extension light for readingmaps or illuminating small areas. These assembliesconsist of a connecting cord, switch, and lamphousing assembly. By adjusting the assembly, youcan change the size of the light beam.

Indicator Lights

Cockpit personnel get aircraft operating statusfrom various indicator (warning, caution, andadvisory) lights in the cockpit. These lights havemany purposes, such as showing the position ofthe landing gear, arresting hook, wings, and bombbay doors. They can also show low oil pressure,equipment overtemperature, generator failure,and other data.

Figure 4-10.-Cockpit utility light.

The construction of indicator lights varies,depending upon the particular job they perform.They mount in the aircraft where they are easilynoticeable when glowing. It is important that bulbreplacement be quick and easy since some lights

relay vital information concerning safety of flight.Whenever practical, assembly design allows bulbreplacement in flight without the use of tools.Most warning light designs have a push-to-testfeature or a test switch that lets you determineif the bulb is good. The test switch energizesvarious test relays, which, in turn, either providepower or a ground circuit for energizing the lights.This allows you to check the bulb withoutoperating the equipment.

Legend-type lights (fig. 4-11) show specificfunctions on the lens surface. They are prominenton late model aircraft.

• The warning lights are red. They warn thecrew of an emergency or unsafe operat-ing condition, which requires immediatecorrective action.

• The caution lights are yellow. They alertthe crew to a minor malfunction or im-pending dangerous condition requiring attention,but not necessarily immediate corrective action.

• The advisory lights are green. They showthe crew a safe or normal configuration,or a performance condition.

Indicator lights must be bright enough to seeduring daylight operation but not bright enoughto cause eyestrain at night. You obtain brilliancecontrol by connecting resistors in the lightingcircuit, and by placing dimmer caps on the lights.Other ways of control are special adaption of edgelighting, and special types of lenses that dim bytwisting the lens.

Figure 4-11.-Legend-type lights.

An example of the use of indicator lights isthe system used to show high oil temperature. Athermoswitch in the oil return line closes when thetemperature reaches above normal. This usuallyprovides a path to ground illuminating the OILHOT light. The light goes out only when the oilcools to normal temperature.

Another application of the use of an indicatorlight is the landing gear unlocked warning light.In some aircraft, a red light glows in thetranslucent landing gear control lever when thegear

is not in either the up or down position.

List the two primary purposes of aircraftlighting.

Why should spare light bulbs be shockmounted?

What are the two most common types ofbulb bases?

What type of light is used to attract visualattention to the aircraft’s position andheading at night?

On which wing tip will the green navigationlight be found?

A bluish-green anticollision light identifieswhat type of aircraft?

When the angle of attack is perfect for thelanding approach, what color AOA lightilluminates?

AIRCRAFT ELECTROHYDRAULICAND PNEUMATIC SYSTEMS

Learning Objective: Recognize operatingprinciples and characteristics of aircraftelectrohydraulic and pneumatic systems.

The word hydraulics is from the Greek wordfor water. Hydraulics originally meant the studyof physical behavior of water at rest and inmotion. Today the meaning includes the physicalbehavior of all liquids, including hydraulicfluid.

Hydraulics is the science of liquid pressure andflow. In its application to aircraft, hydraulics isthe action of liquids, forced under pressurethrough tubing and orifices, to operate variousmechanisms.

The primary concern in working withhydraulics is to control the flow of fluid. You dothis by using solenoids that simply turn on, shutoff a flow of fluid, or change flow direction. Insome cases, electrical devices may schedule aprecise amount of fluid flow. In any case, youmust understand the characteristics of liquids.

You should recall from chapter 1 that a liquidhas no definite shape and conforms to the shapeof its container. Also, a liquid can be only slightlycompressed. Another fact you should recall is theability of a liquid to transmit pressure.

The physics of fluids and basic hydraulicprinciples are contained in Fluid Power,NAVEDTRA 14105. Study the first two chaptersof this publication with this chapter if you needto review basic hydraulic principles.

Although some aircraft manufacturers makegreater use of hydraulics than others, thehydraulic system of the average modern aircraftperforms many functions. Among the unitscommonly operated by hydraulics are the landinggear, wing flaps, speed brakes, wing foldingmechanisms, flight control surfaces, canopy,bomb bay doors, wheel brakes, and arresting gear.

BASIC HYDRAULIC SYSTEM

All hydraulic systems are essentially the same,regardless of their function. Hydraulics are on the

The actuating unit converts fluid pressure intouseful work. The actuating unit may be an

Figure 4-12.-Basic hydraulic system, hand-operated pump.

farm, in industry, aboard ship, and many otherplaces as well as in aircraft. Regardless of itsapplication, each hydraulic system has a minimumof four components—reservoir, pump, selectorvalve, and actuating unit—plus lines throughwhich the fluid flows. Figure 4-12 shows a basichydraulic system, with the four essentialcomponents and their relationship within thesystem.

The reservoir provides storage for a supply offluid for operation of the system. It replenishesthe fluid of the system when needed, providesroom for thermal expansion, and normallyprovides a means for bleeding air from the system.

The pump creates a flow of fluid. The pumpin this system is hand operated; however, aircraftsystems have engine-driven or electric-motor-driven pumps. There are two types used in navalaircraft—the piston type and the gear type. Oneaircraft may incorporate both types due to therequirements of the various hydraulic systems.

The selector valve directs the flow of fluid.These valves actuate either manually or bysolenoid; they may operate directly, or they mayoperate indirectly with the use of mechanicallinkage.

actuating cylinder or a hydraulic motor. Anactuating cylinder converts fluid pressure intouseful work by linear/reciprocating-mechanicalmotion. A hydraulic motor converts fluid pressureinto useful work by rotary-mechanical motion.

Look at figure 4-12, You can trace the flowof hydraulic fluid from the reservoir through thepump to the selector valve. With the selector valvein the #1 position, fluid flow created by the pumpis through the valve to the right-hand end of theactuating cylinder. Fluid pressure then forces thepiston to the left. As the piston moves left, fluidon that side of the piston flows out, up throughthe selector valve, and back to the reservoir. Withthe selector valve in the #2 position, fluid fromthe pump flows to the left side of the actuatingcylinder, reversing the process. You can stoppiston movement at any time by moving theselector valve to neutral. In this position, all fourports close, and pressure is equal in both workinglines.

You can see how hydraulic systems work bylooking at this basic system. You can makeadditions to it to provide additional sourcesof power, operate additional cylinders, makeoperation more automatic, or increase reliability.These additions are all made on the frameworkof the basic hydraulic system diagram shown infigure 4-12.

ELECTROHYDRAULIC SYSTEMS

Basically, there are two types of electricallycontrolled, hydraulically operated components.Selector valves that start, stop, or change thedirection of a fluid flow (similar to a switch inan electrical circuit), and control valves thatschedule fluid flow (much like a potentiometercontrols current flow). Each of these componentsis manufactured in varying degrees of complexity,depending on its use. Each manufacturer uses itsown identification for its components, such astransfer valve, engagement valve, servo valve, etc.

Aircraft hydraulic systems normally operateat a pressure of 3,000 PSI; therefore, eachcomponent must be capable of operating at thispressure. You should observe all applicable safetyprecautions when working with high-pressurehydraulic or pneumatic systems.

Hydraulic Surface Control Booster System AFCS ENGAGEMENT VALVE. —TheAFCS engagement valve (fig. 4-13) is in thedisengage or relaxed position. No voltage is at thecoil, and 3,000 PSI hydraulic pressure pushes thesolenoid piston to the left. This action allow thehydraulic fluid to flow to the right side ofthe spring-loaded piston. With the same pressureon both sides, the spring-loaded piston movesto the left, preventing any further flow in the

The hydraulic surface control booster systemcontains (figs. 4-13 and 4-14) both selector andcontrol components. The automatic flight controlsystem (AFCS) engagement valve is a selectorvalve. It either applies pressure to or removespressure from the AFCS portion of the booster.The hydraulic transfer valve controls the directionand amount of fluid to the booster. system.

Figure 4-13.-Hydraurlic surface control booster with no mechanical input and AFCS disengaged.

Figure 4-14.-Hydraulic surface control booster with no mechanical input and AFCS engaged.

When AFCS engages (fig. 4-14), a 28-volt dc Hydraulic pressure then flows to the other AFCSsignal goes to the solenoid coil to drive its piston components of the boost system.to the right. This action relieves hydraulic pressurefrom the right side of the spring-loaded piston. HYDRAULIC TRANSFER VALVE. —WithThis allows pressure on the left side to move the AFCS engagement, a dc control signal goes frompiston to the right, compressing the spring. the AFCS computer to the transfer valve. In the

static state (zero signal), dc flow through thetransfer valve coil is equal, and the plungermechanism remains in the centered position (withhelp from the centering springs).

Hydraulic pressure from the engagement valveflows to the top land on the hydraulic transfervalve piston. Fluid also flows through a smallorifice to the bottom land of the same piston.Pressure to the bottom of the piston is regulatedby controlling the amount of fluid leaving themodulated pressure chamber. If chamber pressureis 1,500 PSI and one drop of fluid leaves thechamber for every drop that enters, the pressureremains constant. To increase pressure, anunbalanced dc voltage on the coil causes theplunger to block the return orifice for an instantso more fluid enters the chamber than leaves.Conversely, an unbalanced dc signal from theAFCS computer that causes the plunger to risefor an instant lowers chamber pressure.

With zero input signal from the AFCScomputer, the transfer valve piston is in thecentered position (fig. 4-13). The pressure actingon the smaller surface area of the piston’s upperland and the modulated pressure acting on thelarger area of the piston’s lower surface are inbalance, keeping the piston centered.

With an electrical signal input, as shown infigure 4-14, a higher pressure in the modulatedchamber causes the piston to raise. This actionallows fluid to flow to another component of theboost system. The larger the electrical signal input,the more pressure difference on the piston;therefore, more fluid flow.

BOOSTER SYSTEM OPERATION. —Thebooster operates in two separate modes—manualand AFCS. In the manual mode, control surfacedeflection desired by the pilot starts by a lateralmovement of the control wheel (fig. 4-13). Thissmall movement of the control wheel transmitsthrough the feel lever and feel rod to the controllever. The control lever is free to move about itspivot, and a centering spring holds the modulatingpiston in place. Movement of the control leverdisplaces the main control valve, porting fluidfrom two hydraulic systems to the hydraulicactuator piston. (The main control valve andhydraulic actuator work identically, whether inmanual or AFCS modes.)

When the hydraulic actuator piston moves,both the feel lever and the power arm positionthe control surface and the control wheel to theposition the pilot selects. As the pilot applies morepressure to the control wheel, the control surface

moves faster. The direction of the pressuredetermines the direction of control surfacemovement.

If there is no pressure at the control wheel,the main control valve centers (fig. 4-13). Trappedfluid holds the hydraulic actuator piston inposition.

The manual operation of the dual shutoff andbypass valves is from the flight station. If thereis a loss of hydraulic pressure or a malfunctionin the boost system, the flight crew pullsthe handle. Operation of the valves shuts offhydraulic pressure to the hydraulic actuatingpiston and opens both working lines to each other.This prevents hydraulic lock in the actuator pistonand allows control surface movement without theaid of the boost system.

Trapped hydraulic pressure in the modulatorpiston aids the mod piston centering spring inholding the piston centered. This allows positivecontrol of the main control valve by the controllever.

Upon AFCS engagement, the engagementvalve ports hydraulic fluid to the hydraulictransfer valve and to the engagement piston.Movement of the engagement piston locks thecontrol lever in its centered position and preventsmechanical inputs from the control wheels. Thispiston can, however, be overpowered by about15 pounds of force on the control wheel, ifnecessary, in an emergency.

Electrical inputs from the AFCS computerchange to hydraulic fluid pressures in thehydraulic transfer valve, and the pressure controlsthe modulator piston. The modulator pistonchanges the hydraulic pressure into linearmechanical motion, which repositions the maincontrol valve. Movement of the main controlvalve ports hydraulic fluid to the actuator pistonto position the control surface and control wheelas previously described.

Electrical devices on the booster provideinformation inputs to the AFCS. The modulatorpiston linear transducer tells the AFCS the rateof control surface movement, and the surfaceposition synchro transmitter provides positioninformation. Read chapter 8 for more informationabout automatic flight control system operation.

Landing Gear System

Most naval aircraft have hydraulicallyactuated, electrically controlled, retractablelanding gear. Normally, locking the landing gearin the retracted or extended position is automatic.Figure 4-15 is an electrical schematic of the

Figure 4-15.-Landing gear control circuit.

landing gear control circuit of a patrol typeaircraft.

CIRCUIT OPERATION. —The landing gearcircuit includes an electrical, solenoid-operatedselector valve to control hydraulic actuation ofthe landing gear. Current goes to the landing gearselector valve through two single-pole, double-throw (SPDT) switches. These switches operateby a cam on the landing gear control lever,permitting current to flow to either the up or downcoil of the valve. Coil selection depends on controllever position. The landing gear control circuitalso includes electrical control for emergencyextension of the nose gear with emergencyhydraulic system power. A center-off switch(SPDT type) provides control of a doublesolenoid-actuated hydraulic selector valve. Thecenter-off position of the switch is the normalposition during operation of the landing gear withthe main hydraulic system. The down positionselects emergency hydraulic system power for nosegear extension. The bypass position is for use onlyduring retraction of the nose gear with mainhydraulic power after extension by emergencyhydraulic power. The bypass also allows forrelease of emergency system hydraulic pressureany time you desire.

ADDITIONAL CIRCUITS. —In some air-craft, the landing gear control circuit also controlsthe supply of power to the propeller-reversingcircuit (through left and right main torque-linkswitches) and to the stores release circuit (throughleft main torque-link switch). The left main geartorque-link switch supplies power to energize thelanding gear lever locking solenoid. The weightof the aircraft must be off the landing gear shockstrut before you move the landing gear controllever from the wheels down position. The solenoidwill be de-energized by the up sense switch of theleft main gear after the gear retracts. This switchenergizes a relay, whose normally closed contactsare in series with the solenoid circuit. The left andright main gear torque-link switches (parallelconnected) are in series with the power to thethrottle lever-locking solenoid to prevent thethrottles from being placed in reverse propellerrange. When one of the torque-link switchesactuates (by aircraft weight compressing one orboth main landing gear strut oleos), the throttlelevers come into the reverse pitch range. The leftmain landing gear torque-link switches (seriesconnected) also open the stores release circuit.

This prevents accidental release of wing stationexternal stores when the weight of the aircraft hascompressed the shock strut oleo.

A solenoid mechanically prevents movementof the landing gear control lever from thewheels down position when the aircraft weight ison the gear. When the weight of the aircraft ison the landing gear, this solenoid de-energizes,allowing the solenoid armature pin to protrudeoutboard. This solenoid armature pin positionmechanically prevents movement of the landinggear control lever from the wheels down position.Depressing the solenoid armature pin allowscontrol lever movement for emergency or testprocedures.

Arresting Gear System

The arresting gear system stops the aircraftduring carrier landings and emergency fieldarrestments. The primary component of thesystem is the arresting hook on the underside ofthe aft fuselage. The hook is pivoted at theforward end allowing up, down, and sidewaysmotion. The hook engages a runway or cross deckpendant when in the down position. A lever inthe cockpit controls raising and lowering of thehook (fig. 4-16). The lever electrically energizesa hydraulically actuated vertical damper cylinderfor hook retraction and trips an uplatchmechanism through mechanical linkage forlowering the hook. Two horizontal dampers andone vertical damper dampen hook motion fromdeck impact forces. Two centering springassemblies maintain the hook in the centerposition.

EXTENSION. —You extend the arrestinghook by moving the control handle in thecockpit from the up to the down position.This removes tension from the cable and allowsthe uplatch mechanism to deflect to the oppositeextreme of travel (aft). The arresting hookis then free to extend by action of the verticaldamper cylinder and its own weight. At thesame time, a switch in the control handle actuates,de-energizing both the arresting hook relay andthe selector valve solenoid. This action permitsfluid from the vertical damper cylinder to return.The surge damper prevents return line high-pressure surges, caused by hook extension, fromdamaging other subsystems by reducing thesurge.

RETRACTION. —Moving the cockpit leverfrom the down to the up position puts tensionback onto the cable. This action moves the uplatchmechanism forward to receive the arresting hookand latch it in the retracted position. As the leveris brought up, a switch in the lever mechanismactuates, sending current through the de-energizedarresting hook relay to energize the selector valve

vertical damper cylinder to raise the hook. Whenthe hook reaches the retracted position, it actuatesthe up limit switch, which completes the circuit,energizing the arresting hook relay. This, in turn,breaks the circuit to the selector valve solenoidand stops the flow of hydraulic fluid to the verticaldamper. The arresting hook relay has a 1.1 secondtime delay to assure the hook is up and locked

solenoid. Hydraulic pressure then goes to the before removal of hydraulic pressure.

Figure 4-16.-Arresting gear control and indicating system.

Nosewheel Steering System

As you read this section, refer to figure 4-17.The nosewheel steering system is an electricallycontrolled, hydraulically operated system. Itprovides a nonlinear relationship between therudder pedals and the angular position of thenosewheel. The system provides the pilot withadequate directional control of the aircraft duringground operation. It consists of a hydraulicsteer-damp unit, solenoid-operated shutoff valve,command potentiometer, steering feedbackpotentiometer, steering amplifier, and an electricalcontrol system. Control of the electrical power tothe system is through either of the ground safetyswitches on the left or right main landing gear.The landing gear handle switch, a push-buttonswitch on the pilot’s stick grip, and the rudderpedals also control electrical power.

When the steering system electrical circuitsenergize, the steering system aligns for steeringoperation. At this point, the solenoid-operatedhydraulic shutoff valve opens to supply fluid tothe steering actuator. Simultaneously, all relatedcircuitry for controlling the steering servo valveactivates. The electrical section of the steeringsystem is essentially a bridge circuit. One side ofthe bridge circuit runs through the commandpotentiometer; the opposite side runs throughthe feedback potentiometer. Output of eachpotentiometer goes to the steering systemamplifier. Amplifier output currents flow to thehydromechanical servo valve coils, and the netsignal actuates the servo valve, causing nosewheelsteering action.

During operation, the nosewheel steeringsystem attempts to maintain symmetry of theelectrical bridge circuit. The circuit is symmetricalwhen both potentiometers supply equal voltagesto the amplifier and the servo valve is at a nullposition. The bridge circuit becomes unbalancedwhen initiating a turn request by repositioning ofthe command potentiometer wiper. Currentdifferential in the servo valve windings reflects anunbalanced bridge circuit. Movement of the servovalve ports hydraulic pressure to the appropriateside of the actuator piston to cause the nosewheelto turn.

STEER-DAMPER UNIT. —The steer-damperunit is an electrically controlled, hydraulicallyoperated package on the nose gear strut assembly.The unit provides both nosewheel steering and therequired shimmy damping effect. The packageconsists of a check valve, servo valve, bypass

valve, two unidirectional restrictors, the steeringactuator, and fluid compensator.

The check valve prevents reverse flow fromthe unit to the shutoff valve. The servo valvecontrols the actuator position by controlling fluidflow to and from the actuator in response to signalvariations from the amplifier. The bypass valvecloses off the interconnecting passages betweenboth ends of the actuator whenever hydraulicpressure is available to the unit. This permits theactuator to act as a steering unit instead of adamping unit. The unidirectional restrictorsprovide a restricted reverse flow to dampennosewheel shimmy. The steering actuator is abalanced piston-type hydraulic actuator, whichprovides the force to turn the nosewheel. Also,with the restrictors, the steering actuator providesthe shimmy damper action. The fluid com-pensator is in the return passage in the unit andtraps a quantity of fluid at 40 to 100 PSI. Thecompensator supplies fluid to the actuatorthrough the bypass valve and the restrictor whenthe unit is being used as a shimmy damper, andextra fluid is necessary to prevent cavitation ofthe actuator. Since the compensator traps fluidin the actuator, it includes thermal reliefprovisions to prevent excessive pressure buildupwithin the steer-damper unit.

SHUTOFF VALVE. —The steering shutoffvalve is a three-way, two-position, normallyclosed, solenoid-operated valve. The valvecontrols hydraulic system pressure to the steer-damper unit. When the valve de-energizes, fluidflow is cut off from the steer-damper unit. Whenthe valve energizes (during normal steering orarrested landings), pressure flows to the checkvalve, bypass valve, and servo valve in thesteer-damper unit.

COMMAND POTENTIOMETER. —Thesteering system has a rotary-type, pedal-position(command) potentiometer. This potentiometermechanically links to, and is driven by, the rudderpedals. It provides a nonlinear steering response.The potentiometer sends a signal to the nosegear steering amplifier showing the degree anddirection of turn commanded by the pilot. Movingthe potentiometer, with the steering systemoperating, unbalances a bridge circuit causing thesteering amplifier to signal the servo valve to turnthe nosewheel.

FEEDBACK POTENTIOMETER. —Thesteering feedback potentiometer assembly consists

Figure 4-17.-Nosewheel steering system.

of a potentiometer that attaches to and is drivenby the drive arm on the nose gear spindle. As thenose gear moves in response to the pilot’scommand, the feedback potentiometer feeds backa signal to the steering amplifier. When thefeedback potentiometer signal matches thecommand potentiometer signal, the amplifieroutput causes the servo valve to neutralize andstop movement of the nosewheel. The feedbackpotentiometer assembly contains a swivel dis-connect switch, which opens when the nosewheelturns 750 to 800 either side of straight ahead. Thisaction electrically de-energizes the circuit toprevent reverse steering or damage to the aircraft.

AMPLIFIER. —The steering amplifier is asmall transistorized differential amplifier. Theamplifier detects the differential positions of thecommand potentiometer and the steering feed-back potentiometer. Any difference in signalsreceived results in a signal going to the servo valveto port hydraulic pressure to the steering actuator.This causes the nosewheel to turn and the steeringfeedback potentiometer to move. When thefeedback potentiometer signal matches thecommand potentiometer signal, the nosewheelstops turning. In addition, the amplifier containsa circuit that provides a centering signal, whichholds the nosewheel in the straight-ahead positionduring arrested landings.

OPERATION. —When the nosewheel steeringswitch button is depressed, power goes to theelectrical control system and the solenoid-operatedshutoff valve. As the valve opens, hydraulicpressure flows to the servo valve on the steer-damper unit. Signals go to the amplifier from thecommand potentiometer and from the feedbackpotentiometer on the steering linkage. If thesesignals are equal, the amplifier signals the servovalve to stay in the neutral position. When thenosewheel (and consequently the feedbackpotentiometer) does not correspond to rudderpedal position (and command potentiometer), thesignals going to the amplifier are different. Thiscauses the amplifier to send a signal to the servovalve.

The signal sent to the servo valve causes thevalve to port pressure to the steering actuator inthe steer-damper unit. The hydraulic pressurecauses the actuator to move the nosewheel (andthe feedback potentiometer) to a positioncorresponding to rudder pedal position (commandpotentiometer). When the nosewheel reaches aposition corresponding to the pedal position, the

signals to the amplifier are equal. The amplifiernow signals the servo valve to a neutral position.When the servo valve goes to a neutral position(with the steering system energized and hydraulicpressure available), it blocks off both pressure andreturn passages to the steering actuator. Thus, theactuator is hydraulically locked in position. Thesteering actuator will remain locked in positionuntil the servo valve receives a signal to turn thewheel or hydraulic and/or electrical power isremoved from the system. When the steeringsystem is not in use, the steer-damper unitperforms the functions of a shimmy damper. Thesystem accomplishes shimmy damping by trappinghydraulic fluid on both sides of the steeringactuator piston and forcing this fluid from oneside of the actuator to the other side through therestrictor.

Catapulting System

The catapulting system (fig. 4-18) providescatapult handling and attachment capabilities forcarrier operations. The system consists of acatapult launch bar, a launch bar actuatingcylinder and gimbal, swivel joints, a cockpit-controlled selector valve, leaf centering spring,leaf retracting springs, and a catapult tension barsocket. The launch bar is swivel mounted on thenose gear outer cylinder and can extend andretract during taxi operations. The launch barautomatically retracts after catapulting. A launchbar warning light comes on during any of thefollowing conditions:

The launch bar control switch is inEXTEND.

The selector valve is in bar extend position(solenoid A energized).

The launch bar is not up and locked withweight off the landing gear.

The launch bar control switch is inRETRACT and the launch bar actuator isnot up and locked.

Accessories for the catapulting system includea tension bar and a catapult holdback bar. Thecatapult tension bar socket mounts on the nosegear axle beam and provides for attachment ofthe tension bar for tensioning the aircraft beforecatapulting.

Figure 4-18.-Catapulting system control and indicating system.

Placing the launch bar control switch inEXTEND completes a power circuit through theweight-on-gear switch, applying 28-volt dc to thelaunch bar selector valve extend solenoid (solenoidA). The launch bar valve position switch alsocompletes a circuit through its closed contacts toapply 28-volt dc to the launch bar warning light.The warning light comes on and remains on aslong as the switch is in EXTEND. When theselector valve actuates, hydraulic pressure extendsthe launch bar actuator. A plunger on the end ofthe valve mechanically actuates the launch barvalve position switch. Switch actuation completesa parallel circuit to the warning light. When thecontrol switch is in the OFF position, with weighton the landing gear, the warning light should gooff. If the control switch is in RETRACT, thewarning light should come on until the launch baris up and locked. If the light remains on, theselector valve did not cycle from the bar extendposition, and pressure is still on the launch baractuator extend side.

The hydraulic pressure to the launch baractuator unlocks locking fingers inside theactuator and extends the actuator to lower thelaunch bar. As the locking fingers unlock, theyclose the contacts of the launch bar up-lockswitch inside the actuator. After catapulting, the

weight-on-gear switch moves to the weight-offposition. This applies 28-volt dc to the launch barselector valve retract solenoid (solenoid B) and tothe launch bar warning light through the energizedcontacts of relay K1. K1 remains on until thelaunch bar is up and locked. Power to the launchbar selector valve retract solenoid providesautomatic hydraulic retraction and locking of thelaunch bar after catapulting. The launch baractuating cylinder, locking in the fully retractedposition, de-energizes K1, turning off the launchbar warning light. The approach light circuitgoes through de-energized relay K1, giving anadditional indication that the launch bar is up andlocked.

The launch bar control switch goes toRETRACT and is held thereto retract the launchbar hydraulically. This completes a power circuitto apply 28-volt dc to the launch bar selector valveretract solenoid (solenoid B). The energizedselector valve directs hydraulic pressure to retractthe launch bar. The control switch returns to OFFwhen released, de-energizing the selector valve.

Speed Brake System

Speed brakes are moveable control surfacesused for reducing the speed of aircraft. Some

manufacturers refer to them as dive brakes, otherscall them dive flaps. On some aircraft, they areon the sides or bottom of the fuselage; on othersthey attach to the wings. Regardless of theirlocation, speed brakes serve the same purpose onall aircraft—they keep the speed from buildingup too high in dives. They can also slow downthe speed of aircraft preparing to land. Speedbrakes have electrical control and hydraulicoperation. Figure 4-19 shows a typical speed brakesystem, and you should refer to it while readingthis section.

EXTENSION. —Extension of the speedbrakes is done by placing the speed brake controlswitch (located on the throttle lever grip) to theOUT position. The OUT position is a momentary-contact position, the switch being spring loadedto return to STOP when released. Holding thecontrol switch in the OUT position energizes both

selector valve solenoids (solenoids A and B). Thisaction connects hydraulic pressure to the actuatorextend side and connects the return to the actuatorretract side. Any desired brake position may beattained and will be held by a hydraulic lockwithin the selector valve. The speed brake outwarning light illuminates when either the left orright speed brake is not fully retracted. Thewarning light circuit completion is through eitherthe left or right speed brake retract positionswitch.

NOTE: Figure 4-19 shows the speed brakecontrol switch being held in the OUT position.This allows hydraulic pressure to port to theextend side of the speed brake actuators, forcingthe speed brakes to the extend position.

RETRACTION. —In normal operation, thespeed brakes retract by moving the speed

Figure 4-19.-Speed brake system.

brake control switch to the IN position. Thisde-energizes solenoid A by removing the powerto the solenoid and energizes the speed brakerelay. The action of the speed brake relay removesthe power from solenoid B.

With both selector valve solenoids de-energized, the selector valve permits pressure flowto the retract side of the speed brake. It also allowsreturn flow from the extend side to the hydraulicret urn. The speed brake warning light goes outwhen both speed brakes are in the fully retractedposition.

EMERGENCY RETRACTION. —To accom-plish emergency speed brake retraction, place theemergency speed brake switch in the EMER-RETRACT position. This switch de-energizessolenoids A and B of the selector valve (normalsolenoid positions for retracting the speed brakes),which connects both the extend and retract sidesof the speed brake actuators to the system return.With the removal of hydraulic pressure, the speed

brakes are shut by the airstream. This action isnecessary only if the speed brake relay does notenergize when the speed brake control switch isin the IN position. If an electrical failure (suchas a popped circuit breaker) occurs with the speedbrakes extended, retraction is the same asactuation of the emergency retract switch.

Canopy System

As you already know, solenoid valves havemany applications in naval aircraft. The canopyhydraulic system shown in figure 4-20 is anelectrohydraulic system using a solenoid selectorvalve. The canopy selector valve is a four-way,two-position spool valve, which actuates eitherelectrically or manually.

The canopy system consists of a sliding canopymounted over the cockpit area and the com-ponents required for normal operation andemergency jettison of the canopy. The entiresystem is hydraulically operated except for the

Figure 4-20.-Canopy

system.

electrical canopy jettison device. Hydraulic powerfor operation of the canopy is from either thecombined hydraulic system or the hand pumpsystem.

When a canopy switch is in the CLOSEposition, it completes a circuit from the 28-voltdc bus to a terminal on the isolation switch.Current flows from the terminal, through theclosed contacts of the canopy switch, to solenoid2 of the canopy selector valve. The selector valvethen energizes to the close position. Now,hydraulic pressure from either the combinedhydraulic system or the hand pump system flowsthrough the selector valve into the canopy closeline. Pressure through a flow regulator to the rodend of the canopy actuating cylinder causes thepiston and rod to retract closing the canopy.

When the canopy switch is in the OPENposition, power flows from the 28-volt dc busthrough the control circuit breaker, the isolationswitch terminal, the opposite contacts of thecanopy switch, to solenoid 1 of the canopyselector valve. The selector valve energizes to theopen position, reversing the sequence of pressureand return flow. Hydraulic pressure flows through

the canopy open line to the canopy seal valve andhydraulically trips the seal valve. Thus, the airpressure in the canopy seal is dumped, allowingthe seal to deflate. Pressure in the canopy openline continues its flow through a flow regulatorto the backhead end of the cylinder. This actionextends the cylinder, opening the canopy.Hydraulic fluid in the opposite end of the cylinderreturns through the canopy close line to theselector valve, across the valve, and into thecombined hydraulic system. Fluid then returns tothe main reservoir of the system.

For canopy jettison (fig. 4-21), the emergencycanopy jettison switch, or either of the twoemergency outside jettison switches initiates afiring circuit. Closing any of the three switchescompletes the circuit to the electrically firedcanopy jettison cartridge on the canopy cylinderbackhead end. The cartridge discharges throughthe canopy cylinder, causing the canopy tojettison.

PNEUMATIC POWER SYSTEM

The pneumatic power system supplies com-pressed air for various normal and emergency

Figure 4-21.-Canopy jettison.

pneumatically operated systems. The compressedair is held in storage cylinders in the actuatingsystems until required by actuation of the system.These cylinders and power system manifoldreceive an initial charge of compressed air ornitrogen from an external source. In flight, theair compressor replaces the pressure and volumelost through leakage, thermal contraction, andsystem operation.

The air compressor receives supercharged airfrom the engine bleed air system. This ensures an

adequate air supply to the compressor at allaltitudes. The air compressor is driven by anelectric or a hydraulic motor. The system underdiscussion in this chapter is hydraulically driven.(See figure 4-22.)

The aircraft hydraulic system provides powerto operate the hydraulic-motor-driven air com-pressor. The air compressor hydraulic actuatingsystem consists of a solenoid-operated selectorvalve, flow regulator, hydraulic motor, and motorbypass line check valve. When energized, the

Figure 4-22.-Pneumatic system.

selector valve allows the system to pressurize andrun the hydraulic motor. When de-energized, thevalve blocks off hydraulic pressure, stopping themotor. The flow regulator compensates for thevarying hydraulic system flow and pressures bymetering the fluid flow to the hydraulic motor.Thus, excessive speed variation and/or over-spending of the compressor is prevented. A checkvalve in the motor bypass line prevents systemreturn line pressure from entering the motor andstalling it.

The air compressor is the pneumatic system’spressurizing air source. The compressor activationor deactivation is by the manifold pressure sensingswitch, an integral part of the moisture separatorassembly.

The moisture separator assembly is thepneumatic system’s pressure sensor-regulator andrelief valve. The manifold pressure switch governsair compressor operation. When manifoldpressure drops below 2,750 PSI, the pressuresensing switch closes, energizing the separator’smoisture dump valve and hydraulic selector valve,activating the air compressor. When manifoldpressure reaches 3,150 PSI, the pressure sensingswitch opens, de-energizing the hydraulic selectorvalve, deactivating the air compressor and dumpvalve. The deactivated dump valve ventsoverboard any moisture in the separator. Theseparator includes a thermostat and heatingelement. The thermostatically controlledwraparound blanket heating element preventsmoisture from freezing within the reservoir in low-temperature atmospheric conditions.

The safety fitting, at the moisture separatorinlet port, protects the separator from internalexplosions due to hot carbon particles or flamesthat the air compressor may emit.

A chemical drier further reduces the moisturecontent of the air emerging from the moistureseparator.

An air charge valve provides the entirepneumatic system with a single external groundservicing point. An air pressure gauge, near theair charge valve, is for servicing the pneumaticsystem. This gauge shows manifold pressure.

The ground air charge line air filter preventsentry of particle impurities into the system fromthe ground servicing power source.

Q1. The science of liquid pressure and flow isknown as

Q2. As an AE your primary concern withhydraulics will be to

Q3. In a hydraulic system, what unit createsfluid flow?

Q4. What does the hydraulic actuating unitconvert fluid pressure into?

Q5. Why does the arresting gear hook relay havea 1.1 second delay?

Q6. What system provides the pilot withadequate directional control during groundoperations?

Q7. During emergency retraction, what is usedto close the speed brake?

Q8. The pneumatic system compressor comeson when system pressure drops below whatvalue?

AIRCRAFT ENVIRONMENTALSYSTEMS

Learning Objective: Recognize terms,definitions, components, and aircraftenvironmental systems operating principlesand features including controls fortemperature, pressure, and icing.

The proper operation of today’s aircraftrequires maintaining a proper environment not

only for personnel but also for equipment onboard. Similarities exist in the electronicequipment controlling these systems and thecomponents within these systems.

The environmental systems on most aircraftinclude cockpit/cabin air conditioning andpressurization, equipment cooling, windshieldanti-icing and defogging, and equipment pressuri-zation systems. Some aircraft accomplish equip-ment cooling by teeing off with ducting from thecockpit/cabin system; other aircraft use a separatesystem. These systems are not the exclusiveresponsibility of the AE, but rather are theresponsibility of other ratings with the AEassisting.

TERMS AND DEFINITIONS

The system that maintains cabin air tempera-tures to meet the requirements for pilot efficiencyis the air-conditioning system. You must becomefamiliar with some terms and definitions tounderstand the operating principles of air-conditioning systems.

The following terms are self explanatory:engine heat, solar heat, electrical heat, a n dbody heat of personnel. These sources ofheat make cabin air conditioning necessary,and another source is ram air temperature.Ram air temperature is the frictional tempera-ture increase created by ram compression onthe skin surface of an aircraft. This factorbecomes serious only at extreme airspeeds.For example, for an aircraft flying at 45,000feet, at 1,200 MPH, the ram air temperaturewould be about 2,000°F on some parts ofthe aircraft. This extreme temperature, plus theheat from other sources, would cause cabintemperature to rise to about 190°F. The maximumtemperature that a crew member can endure andstill maintain top physical and mental efficiencyis about 80°F. Prolonged exposure to atemperature greater than 80°F will seriouslyimpair his/her mental and physical abilities.Furthermore, under low-speed operating con-ditions at low temperature, cabin heating may benecessary.

NOTE: Some of the terms were discussed inchapter 1, “Elementary Physics.” Although theyare briefly defined in this chapter, a review ofchapter 1 will be helpful.

Definitions for some of the terms relative totemperature control are as follows:

Absolute temperature. Temperature measuredalong a scale that has zero value at that pointwhere there is no molecular motion (–273.1°C or–459.6°F).

Adiabatic. A word meaning no transfer ofheat. The adiabatic process is one in which no heattransfers between the working substance and anyoutside source.

Ambient temperature. The temperature in thearea immediately surrounding the object underdiscussion.

Ram air temperature rise. The increase intemperature created by the ram compression onthe surface of an object traveling at high speedthrough the atmosphere. The rate of increase isproportional to the square of the speed of theobject.

Temperature scales.

1. Celsius (C)—A scale on which 0° representsthe freezing point of water, and 100° is the boilingpoint of water at sea level.

2. Fahrenheit (F)—A scale on which 32°represents the freezing point of water, and212° is the boiling point of water at sealevel.

CABIN SYSTEM

The primary function of the cabin air-conditioning and pressurization system is tomaintain cockpit temperature and pressure withinparameters for crew safety and comfort. To dothis, the system forces a mixture of dehumidifiedrefrigerated air and hot engine bleed air throughcockpit diffusers. The temperature of themixture is automatically maintained through acontinuously selective range by a temperaturecontrol system, consisting of temperature sensorswith associated flow control valves and anelectronic controller. A pressure regulator andsafety valve control cabin pressure. The manualdump control is available if the safety valvemalfunctions. A block diagram for the airflow of

Figure 4-23.

a cabin air-conditioning and pressurization systemis shown in figure 4-23.

System Controls

The air-conditioning control panel is shownin figure 4-24. Three switches and one controlknob on the panel control the air-conditioningoperation.

The air-conditioning master switch is atwo-position (NORM and OFF) toggle switch,guarded to the NORM position. When the switchis in NORM, the main bleed air shutoff valvesopens if the engine is providing air at 8 PSI ormore. Also the air-conditioning caution light willgo to off. Hot engine bleed air is now availablefor operation of the various environmentalsystems. Placing the switch in the OFF posi-tion closes the shutoff valve, thus shuttingoff hot engine bleed air to the environmentalsystem. A spring-loaded guard over the air-conditioning master switch prevents the switch

Figure 4-24.-Temperature control panel.

from being placed in the OFF position bymistake.

The cockpit switch is a three-position toggleswitch marked ON, OFF, and RAM AIR. Whenthe engine is running and the air-conditioningmaster switch at NORM, placing the cockpitswitch to ON engages the air-conditioning andpressurization system. Placing the switch to ONenergizes the dual temperature control valve andopens the cabin bleed air shutoff valve. Dependingon the setting on the automatic temperaturecontrol thumbwheel, conditioned air flows intothe cockpit. When the cockpit switch goes to theOFF position, the cabin bleed air shutoff valvecloses and drives the dual temperature controlvalve to the full hot position. This prevents anyairflow to the cockpit. The RAM AIR positioncloses the cabin bleed air shutoff valve and drivesthe dual temperature control valve to the full hotposition. You now control the ram air flowmanually using the MAN/RAM AIR switch.

The MAN/RAM AIR switch is a four-position toggle switch marked AUTO, HOLD,COLD, and HOT. The AUTO position enablesthe dual temperature control valve to acceptinputs from the automatic temperature controlthumbwheel. Place the switch to HOLD when thedual temperature control valve is removed fromautomatic control. The switch is spring loaded toHOLD when not in AUTO. Momentarily holdingthe switch in COLD or HOT alters the positionsof the hot and cold sides of the dual temperaturecontrol valve to change the cockpit air tempera-ture. Upon releasing the switch, it springs backto HOLD, and the dual temperature control valveremains fixed. The COLD and HOT positionsshould be toggled intermittently to avoidovershooting the desired temperature. With theselection of RAM AIR on the air-conditioningcockpit switch, you manually control airflow byusing the RAM AIR switch. The RAM AIRswitch opens and closes the ram air valve.

The automatic temperature control thumb-wheel enables the crew to adjust the cockpit airtemperature. With the air-conditioning controlsappropriately positioned, you can move thethumbwheel to any setting between 0 and 14.Rotating the thumbwheel automatically regulatesthe openings of the hot and cold sides of the dualtemperature control valve, thus varying thecockpit air temperature. You can select atemperature within a range of 60°F to 80°F.

System Operation

As you read this section, refer to figure 4-25.Placing the cockpit switch to ON causes contact3 to remove power from the cabin bleed airshutoff valve. This valve opens, allowing hot airto flow to the dual temperature control valve.Simultaneously, a second portion of the cockpitswitch, ganged to the first, connects the center tapof the MAN/RAM AIR switch through energizedcontact of the air-conditioning control relay, theCOLD side of the cabin duct thermal switch, andcontact 4 of the COCKPIT SWITCH to thecontrol phase of the dual temperature controlvalve. Contact 1 of the cockpit switch powers thecoil of the air-conditioning control relay when thecockpit switch is in the ON or OFF position.Cockpit switch contact 2 (in the ON or OFFposition) supplies 120 volts to the close windingof the cabin ram air valve. A phase shiftingcapacitor powers the open winding of the samevalve, which drives the two-phase induction motorto the full close position.

In the air-conditioning mode (cockpit switchON), hot engine bleed air flows through thecabin shutoff valve to the dual temperaturecontrol valve, where it mixes with refrigerated,dehumidified air. The cabin temperature controlsystem automatically maintains the cabin at aselected temperature by continuously modulatingthe dual temperature control valve. The modu-lating signals are cabin temperature errors that thecabin temperature sensor translates into errorvoltages. The cabin and ventilation suit tempera-ture controller modulates and amplifies thesesignals. Then, they go to the control winding ofthe dual temperature control valve. The hot sideof the valve starts to close and the cold side opensuntil the selected cabin temperature is achieved.To prevent temperature hunting by the cabintemperature sensor, the cabin duct dual tempera-ture sensor anticipates changes in cabin tempera-ture by sensing changes in the duct inlettemperature.

The dual temperature sensor transmits errorsignals to the cabin and ventilation suittemperature controller when cabin air supplytemperature changes. The size of the transmittedsignals depends on the rate of temperaturechanges. When the cabin duct inlet temperatureexceeds 121.1°C (250°F), the overtemperaturesensor portion of this sensor transmits a fixedsignal. The signal goes through the temperaturecontroller to the dual temperature control valvecontrol winding. The valve drives to the full cold

position until duct temperature returns to normal,then automatic control resumes.

The cabin duct thermal switch provides anadditional override to the cabin temperaturesensor. This temperature-sensitive switch closeswhen the duct inlet temperature exceeds 121.1°Cor the dual temperature sensor malfunctions.During either of these conditions, 120-volt ac fromthe ENVIRONMENT/SEATS CB goes throughthe HOT side of the cabin duct thermal switchand contact 4 of the cockpit switch to the controlwinding of the dual temperature control valve.The valve drives to the full cold position. Whencabin duct inlet temperature falls to 121.1°C, theswitch goes to the cold position and automatictemperature control again resumes.

In the automatic mode, positioning theAUTO TEMP CONT thumbwheel selects cockpittemperature. This fixes the output voltage of thebridge circuit in the cabin and ventilation suitcontroller. The controller amplifies the signaland transmits through contact 4 of the air-conditioning relay to the AUTO position of theMAN/RAM AIR switch. The signal travels backthrough contact 3 of the air-conditioning relay tothe COLD side of the cabin duct thermal switch.The signal then goes through the cockpit switchto the control winding of the dual temperaturecontrol valve.

To gain manual control of the cabin air-conditioning system, momentarily position theMAN/RAM AIR switch at COLD or HOT. Inthe COLD position, power from the ENVIRON-MENT/SEATS CB goes through contact 1 of theair-conditioning control relay, the MAN/RAMAIR switch, through contact 3 of the air-conditioning control relay, the COLD contacts ofthe cabin duct thermal switch, and contact 4 ofthe cockpit switch to the control winding of thedual temperature control valve. This voltage lagsthe fixed-phase voltage by 900 and drives the two-phase induction motor. This allows morerefrigerated air and less hot air to enter the cabin.The valve will continue to move until the switchis released. After the switch is released, it returnsto HOLD and the valve remains fixed.

When the MAN/RAM AIR switch is toggledto the HOT position, power from the ENVIRON-MENT/SEATS CB goes through the reversingtransformer, contact 2 of the air-conditioningcontrol relay, the MAN/RAM AIR switch,contact 3 on the air-conditioning control relay,the COLD contacts on the cabin duct thermalswitch, and contact 4 of the cockpit switch to thecontrol winding of the dual temperature control

Figure 4-25.

valve. Because the voltage goes through the reversingtransformer, it shifts 180°, and the control windingvoltage leads the fixed-phase voltage by 90°. Thiscauses the valve to drive in a direction to decreasecold air and increase hot air to the cockpit. Uponreleasing the switch, the valve remains fixed.

Under any of the previously describedconditions—in either the automatic or manual mode—failure of electrical power to the cabin and ventilationsuit temperature controller, or to the dualtemperature control valve, causes the valve to remainin the position it was maintaining at the time ofelectrical failure.

The cabin may be ram air ventilated by placingthe cockpit switch at RAM AIR. With the canopyclosed, this action de-energizes the air-conditioningcontrol relay, closes the cabin bleed air shutoff valve,and drives the dual temperature control valve to thefull hot position. This effectively blocks the flow ofboth hot and cold bleed air to the cabin. Either crewmember may control the degree of ram air ventilationby depressing the MAN/RAM AIR switch to eitherHOT OR COLD. The HOT position drives the cabinram air valve toward the closed position, restrictingthe ram air flow. The COLD position drives the cabinram air valve toward the open position, increasing theram air flow. You can select any position of the cabinram air valve from full open to full closed.

Cabin pressurization automatically starts at8,000 feet. The cabin pressure remains at 8,000 feetuntil reaching a pressure differential of 5 PSI. Thecabin pressure now maintains a pressure differentialof 5 PSI.

The cabin pressure regulator maintains thepressure schedule. This valve allows the air passedinto the cabin air-conditioning system to exhaust intothe nose compartment. From sea level to 8,000 feet,the exhaust flow is unrestricted. Above 8,000 feet, thevalve closes down, restricting the exhaust flow untilreaching a cabin pressure equivalent to 8,000 feetaltitude. The cabin remains at this pressure as theaircraft climbs until reaching a cabin-to-ambientpressure differential of 5 PSI. Beyond this altitude,the valve regulates the exhaust flow to maintain thisconstant pressure differential throughout theremaining altitude range.

A cabin pressure safety valve provides amaximum limit to cabin pressure in the event of aregulator malfunction. This valve limits the cabin-to-ambient pressure differential to 5.5 PSI. The CABINDUMP switch permits either crew member to dumpcabin pressure through the cabin safety valve. Placingthe CABIN DUMP switch to ON completes a circuit tothe cabin dump valve, which opens the cabin safetyvalve.

Figure 4-26.—Equipment cooling flow diagram.

EQUIPMENT COOLING SYSTEM

Ram air is the primary means of ventilationfor the forward and aft equipment compartments.The ram air thermal switch, equipment coolingvalve, forward compartment ram air valve, andthe aft compartment ram air valve controlventilation of these compartments.

The equipment cooling system flow diagramis shown in figure 4-26. When the aircraft is inflight, the equipment cooling valve closes and theforward and aft compartment ram air valves

open. Under these conditions, the right forwardand aft equipment compartments are ram airventilated. When temperature in the right wingram air duct reaches 46.1°C (115°F), theequipment cooling valve opens, and the forwardand aft compartment ram air valves close. Thispermits moist, cooled bleed air into the rightforward and aft equipment compartments,ensuring adequate cooling when ambient airtemperatures are excessive.

With an engine running, the equipmentcooling system (fig. 4-27) automatically engages

Figure 4-27.-Equipment cooling schematic diagram.

when the AIR COND MASTER switch is atNORM. Under these conditions, the circuit fromthe 28-volt dc essential bus to the solenoid of theengine bleed air shutoff valve is complete. Bleedair now flows to the air cycle refrigeration unit.The discharge from the cooling turbine suppliesthe cooled bleed air used by the equipment coolingsystem.

When the aircraft is airborne and the ram airtemperature is below 46.1°C (115°F), voltagefrom the 120-volt ac primary bus runs throughthe overpressurization relay, and the unoperatedram air thermal switch to the equipment coolingvalve, the forward compartment ram air valve,and the aft compartment ram air valve. Theequipment cooling valve runs to the fully closedposition, and the forward compartment and aftcompartment ram air valves run to the openposition. Ram air flows to the right forward andaft equipment compartments and the cooled,bleed airflow is shutoff.

When the ram air temperature exceeds 46. 1°C,the ram air temperature thermal switch operates.Voltage from the 120-volt ac primary bus goesthrough the overpressurization relay, and theoperated ram air thermal switch to the equipmentcooling valve, the forward compartment ram airvalve, and the aft compartment ram air valve. Theequipment cooling valve runs to the open position,and the forward compartment and aft compart-ment ram air valves run to closed position.Undried, cooled bleed air flows to the forwardand aft equipment compartments. The ram air tothese compartments is shutoff.

When an overpressurization condition occurs,the pressure switch operates. This closes theoverpressurization relay. Voltage from the120-volt ac primary bus runs through theoverpressurization relay to the equipment coolingvalve, the forward compartment ram air valve,and the aft compartment ram air valve. Theequipment cooling valve goes to the open positionand the forward and aft compartment ram airvalves run to the closed position. Under theseconditions, the overpressure dumps into theforward and aft equipment compartments.

The computer pressure regulator maintains apressure of 2 to 3 PSI in the equipment coolingline. The pressure-regulated, partially-dried cooledbleed air from the regulator automatically mixeswith controlled quantities of hot bleed air. Theservo-controlled computer temperature controlvalve controls the mixture process. This valvemodulates in response to temperature signals fromthe computer temperature sensor in the computer

inlet duct to maintain a computer ducttemperature of 4.4° ± 2.8°C (40° ± 5°F). Thetemperature controlled air flows through theballistics computer set. Sonic venturis in theballistics computer outlet duct and the transmittermodulator inlet duct limit the flow of cooling air.

A computer cooling shutoff valve upstreamof the computer temperature control valveprovides a safety override for the computertemperature control valve. If the computer ducttemperature exceeds 65.6°C (150°F), thecomputer thermal switch opens. This de-energizesthe computer overheat indicator relay andcompletes the circuit to the COMPUTER OVER-HEAT caution light. Voltage from the 28-voltessential dc bus runs to the computer overheatindicator relay. This energizes the computercooling control relay, thereby closing thecomputer cooling shutoff valve. If the computerduct temperature drops below 65.6°C, thecontacts of the computer thermal switch close,energizing the computer overheat indicatorrelay. This action interrupts the circuit to theCOMPUTER OVERHEAT caution light. Thecomputer cooling control relay remains energizedthrough the computer emergency cooling switchand its own contacts. Placing the CMPTREMER COOL switch momentarily to RESETde-energizes the computer cooling control relay,completing the circuit to open the computercooling shutoff valve. If the COMPUTEROVERHEAT caution light should cycle on andoff indicating a constant overheat condition, placethe CMPTR EMER COOL switch to ON,permanently closing the computer cooling shutoffvalve. This allows uncontrolled, cooled bleed airto duct to the ballistics computer.

Air exhausted from the ballistics computer setand from the cockpit circulates through the lesscritical electronic equipment compartments. Thisair supplies direct cooling requirements forcommunication, navigation, and identification(CNI) equipment in the aft equipment compart-ment.

CABIN AND VENT SUITTEMPERATURE CONTROL SYSTEM

The cabin and vent suit temperature controlleris a transistorized electronic device. It operates on120 volts, 400 hertz. Maximum power con-sumption is 28 watts.

Electrically, the controller consists of twochannels—the cabin temperature channel and theventilated suit channel. Both channels operate in

basically the same manner. Bridge circuitscompare temperature selector and temperaturesensor resistances, which represent selected andactual temperature conditions and generate aresultant dc error voltage. The dc error voltageis then modulated and amplified. The resultingac signal is the controller output, and it has phaseand voltage characteristics proportional to themagnitude of the dc error signal. It goes to thetwo-phase servomotor of the appropriate controlvalve, where it modulates the valve to maintainthe selected temperature.

The cabin temperature channel (fig. 4-28) ofthe temperature controller uses three bridgecircuits to maintain cabin temperature. Selectionof the desired temperature with the cabintemperature selector varies resistance of oneleg of the cabin temperature bridge. The varyingcabin temperature changes the resistance of thecabin sensor in the second leg of the bridge circuit.Thus, a selected temperature must have a changeof cabin temperature to provide bridge balancebetween the cabin sensor and selector. Whenever

bridge imbalance exists, the resulting dc voltageis ac modulated by the modulator circuit,amplified by the cabin amplifier, and fed to thecontrol valve servomotor to either increase ordecrease cabin temperature.

The function of the cabin duct limit bridge isto limit the temperature of the cabin inlet air. Adiode-biasing network permits the cabin duct limitbridge to override the cabin temperature bridgeupon reaching the duct temperature limit. Theresulting dc voltage goes to the modulator circuit,the cabin amplifier, and to the control valveservomotor to decrease cabin temperature.

The cabin duct anticipator bridge functionswith sudden changes of air temperature in thecabin air inlet duct. To do this, the error voltageis capacitor coupled to the modulator. When thecabin temperature is being held at the selectedtemperature with constant cabin inlet air tempera-tures, the anticipator bridge is in balance. Shouldduct temperature suddenly change, with all otherconditions remaining the same, the anticipatorbridge is unbalanced. This unbalance causes an

Figure 4-28.-Cabin temperature channel.

error voltage to go to the modulator circuit. Theresulting amplified signal regulates the cabincontrol valve to return the duct air to its originaltemperature. Error voltages caused by cabintemperature bridge or cabin duct limit bridgeimbalance override the cabin duct anticipatorbridge, provided the duct air temperature erroris gradual or small.

One additional error voltage is capacitorcoupled to the modulator circuit. The voltagechange is proportional to the rate of change ofthe feedback potentiometer of the cabin dualtemperature control valve. This voltage changeis applicable only when the valve is beingregulated (the feedback potentiometer is rotating).Therefore, the error feedback voltage seen at thecontrol input is proportional to the feedbackpotentiometer rate of change. Once regulation ofthe valve starts, the potentiometer rate-of-changevoltage reduces initial starting voltage to thecontrol valve actuator motor, slowing valveactuator rotation.

The vent suit channel (fig. 4-29) of thetemperature controller uses one bridge circuit tomaintain suit temperature. Operation of thisbridge is similar to that of the cabin temperaturebridge. The suit temperature selector resistancein one leg balances againsttemperature sensor resistanceleg. Imbalance between these

the suit ductin the oppositelegs of the suit

temperature bridge results in a dc error voltage.This voltage is ac modulated by the modulatorcircuit, amplified, and goes to the suit temperaturecontrol valve, altering suit air temperatures.

Rate-of-change voltage from the feedbackpotentiometer in the ventilated suit temperaturecontrol valve is capacitance-coupled to themodulator circuit. As in the cabin temperaturechannel, this voltage is present only when the valveis operating. It reduces the initial starting voltageto the valve once actuator rotation has started,thus slowing actuator rotation.

ANTI-ICING AND DEICINGEQUIPMENTS

The anti-ice and defrost system on someaircraft having the air-conditioning and pressuri-zation system described in this chapter use the aircycle systems as an air source. The anti-ice anddefrost equipment consists of the windshield anti-ice system (fig. 4-30.) and the windshield defrostsystem. Each receives its hot air supply from thesame manifold. The anti-icing switch is on thepilot’s temperature control panel.

In a typical system, a windshield overheatthermostat and a shutoff valve in the windshielddefrosting duct operate together to preventwindshield overheating. The thermostat opensthe valve automatically when the temperature

Figure 4-29.-Vent suit channel.

Figure 4-30.

becomes too high. This action diverts the hotdefrosting air to the air-conditioning outlet at thefloor of the cockpit until the temperature drops.Windshield overheating occurs only if the cabinair temperature high limit pickup fails.

Windshield Anti-Icing andDefogging System

An electrical anti-icing and defogging systemis now in the windshield panels of current navalaircraft. The panels are constructed of two piecesof semitempered plate glass laminated with a vinylplastic core. The core acts as a safety device toprevent shattering in a collision with birds whenit is in the heated condition. The resistance heatingelement for anti-icing and defogging consists ofa transparent, electrically conductive film, evenlydistributed over the inner surface of the outer paneof glass. The system includes the followingadditional components:

• A windshield wire terminal box, locatedbetween the windshield panels

• A temperature-sensing element embeddedin each panel

• Two windshield autotransformers and aheat control relay

• A dual windshield control unit

• A windshield heat control toggle switchlocated in the cockpit

The system receives power from the ac busesthrough the windshield heat control circuitbreakers. When the windshield heat control switchis in HIGH, 115 volts, 400 hertz goes to the leftand right amplifiers in the dual windshield controlunit. The windshield heat control relay thenenergizes, applying two phases of ac power at200 volts, 400 hertz to the windshield heatautotransformers. These transformers provide218-volt ac power to the windshield heatingcurrent bus bars through the dual windshieldcontrol unit relays. The sensing element ineach windshield has a POSITIVE temperaturecoefficient of resistance and forms one leg of abridge circuit. (Some systems use a thermistor,which has a negative temperature coefficient,as a sensing element to control windshieldtemperature.)

When windshield temperature is above cali-brated value, the sensing element has a higherresistance value than needed to balance the bridge.This decreases the flow of current through theamplifiers, and the relays of the control unitare de-energized. As the temperature of thewindshield drops, the resistance value of thesensing element also drops. Now, the currentthrough the amplifiers again reaches sufficientmagnitude to operate the relays in the control unit,thus energizing the windshield heaters.

When the windshield heat control switch is inLOW, autotransformers provide 115 volts, 400hertz ac to the left and right amplifiers in the dualwindshield control unit. In this condition, thetransformers provide 121-volt ac power towindshield heating current bus bars through dualwindshield power relays. The sensing elements inthe windshield operate in the same manner asdescribed for high heat operation. The calibratedunits maintain windshield temperature between40°C and 49°C (105°F to 120°F).

Wing and Tail Anti-Icing

Some aircraft have a thermal anti-icing systemthat prevents formation of ice on the leading edgeof the wing panels and tail surfaces. This systemuses hot air to heat the leading edges. The hot aircomes from a combustion heater in the wings andtail section, or from the compressor section of ajet engine. An electrically operated pump suppliesfuel for the anti-icer heaters. Heater demanddetermines the amount of fuel the pump delivers.In some aircraft, thermostats automaticallydetermine heater demands. Various types andcombinations of solenoid-operated valves controlfuel flow and airflow in the system.

Wing and Tail Deicing

Some aircraft may have air-inflated, rubberdeicer boots on the leading edges of wing and tailsurfaces. Air pressure or vacuum is alternatelyapplied to these boots and cracks off any ice thathas formed. Once the ice has cracked, the forceof the airstream peels it back and carries it away.Pressure for inflating the air cells in the deicerboots is normally from engine-driven pumps. Airpressure or vacuum is alternately applied eitherby the use of motor-driven rotary distributionvalves or by the combination of an electronictimer and solenoid distributor valves.

Empennage Deicing

The P-3 empennage deicing system combinesboth anti-icing and deicing for the verticaland horizontal stabilizers. The system useselectrical power to prevent or remove theaccumulation of ice. Electrical heating ele-ments are in the leading edges and surfacesof the empennage. The system has ac powerand dc control. Parting strips in the leadingedge of each stabilizer accomplish anti-icingrequirements. The parting strips remain ononce the system actuates. Twenty cyclicheat areas accomplish deicing requirements.During normal operation, each of the cyclicareas heat for 8 seconds and are off for168 seconds. Overheat sensors and a thermal-sensitive relay protect the system. Any timethe system detects an overheat, power is shutoff automatically to prevent damage to the metalstructure.

Pitot Tube Anti-Icing

To prevent the formation of ice over theopening in the pitot tube, a built-in electricheating element (fig. 4-31) is in use. A switchon the pilot’s console controls power to theheaters. System power is from either the acor the dc bus. Exercise caution when groundchecking the pitot tube since the heater can notoperate for long periods unless the aircraft is inflight. Also, the danger exists for groundpersonnel to be accidentally burned.

Figure 4-31.-Pitot tube anti-icing circuit.

The frictional temperature increase createdby ram compression on the skin surface ofthe aircraft is known as

What is theconditioning

primary purpose of an air-and pressurization system?

Where is the resistance heating elementinstalled in electrically heated windshields?

What do aircraft equipped with wing deicesystems use to prevent ice buildup on theleading edge?

What does the P-3 empennage anti/deicingsystems use to prevent ice buildup?

CHAPTER 5

AIRCRAFT POWER PLANT ELECTRICALSYSTEMS

A naval aircraft requires a mechanism topropel it through the air. Aviation Machinist’sMates (ADs) maintain the power plant and itsassociated controls. There are, however, severalelectrical systems (such as starting, ignition, firewarning, fire extinguishing, anti-icing, fuel, andindicating systems) that support the power plant.AEs maintain these systems.

STARTING EQUIPMENT

Learning Objective: Identify the types ofstarting equipment used to start aircraftengines.

Jet engine starters provide high starting torqueto initially overcome the engine rotor weight andhigh speed to increase rotor RPM until the engineis self-sustaining. The following paragraphsdescribe the various starting systems used onturbojet, turboprop, and turbofan engines.

CONSTANT SPEED DRIVE/STARTER

The starting system on the A-6 Intruderaircraft is the constant speed drive/starter(CSD/S). It has three separate modes ofoperation:

1. Start mode.2. Constant speed drive mode. This mode

maintains main generator frequency controlregardless of engine RPM or generator loads.

3. Air turbine motor generator mode. Thismode is used for auxiliary and emergencyoperation of a main generator when the mainengine is shut down.

Air for starting or emergency/auxiliaryoperation of the generator is received from theother engine or from an external source.

The CSD/S unit (fig. 5-1) consists of twosections and connects to the engine input drive

shaft through a generator. The turbine motorsection contains an axial flow turbine wheel andvariable area nozzle that directs air and controlsturbine wheel speed. The planetary differentialtransmission section is a variable gear assemblywith infinite gear ratios.

In the start mode of operation, an externalair source drives the turbine wheel. The gearreduction in the transmission section converts thehigh speed of the turbine to high torque. Thehollow core of the generator shaft delivers torqueto the engine input drive shaft.

After the engine gains a self-sustaining speed,the ring brake releases and the system changes tothe constant speed drive mode. In this mode,electromechanical, pneumatic, and pneumatic-hydraulic components act to vary the gear ratioof the planetary differential transmission. Thesecomponents maintain an output of 8,000 RPMfrom the engine input drive shaft.

If it becomes necessary to shut down anengine, airflow routes from the remaining operat-ing engine to the CSD/S. This places the systemin the air turbine motor-generator drive mode ofoperation. In this mode, power to drive thegenerator comes entirely from the turbine drivewheel. A planetary differential transmissionsection disconnects the generator and turbine fromthe engine.

AIR TURBINE STARTER

The air turbine starter is a lightweight unitdesigned to start turbojet, turboprop, andturbofan engines when supplied with compressedair. The unit consists primarily of a scrollassembly, rotating assembly, reduction gearsystem, overrunning clutch assembly, and outputshaft. An overspeed switch mechanism limitsmaximum rotational speed. Compressed air,supplied to the scroll inlet, goes to the turbinewheel through the nozzle in the scroll assembly.The reduction gear system transforms the highspeed and low torque of the turbine wheel to low

Figure 5-1.-Constant speed drive/starter.

speed and high torque at the output shaft. Whenat the desired starter rotational speed, theflyweights in the governor assembly throw out andopen the limit switch. This section sends a signalthat shuts off the supply air. At a higher,predetermined rotational speed, the overrunningclutch assembly releases the output shaft from therotating assembly.

A source of compressed air flows to theshutoff valve inlet duct to drive the starter. Thestarter is a turbine air motor equipped with aradial inward-flow turbine wheel assembly,reduction gearing, splined output shaft, and aquick-detaching coupling assembly. The completeassembly mounts within one scroll assembly andgear housing (fig. 5-2). As you read this section,you should refer to figures 5-2 and 5-3.

The air turbine starter converts energy fromcompressed air to shaft power. This power goesto a splined output shaft at speed and torquevalues for starting aircraft engines.

Initial control of the air shutoff valve (fig. 5-3)is by a normally open, momentarily closed, startswitch and a relay box. After pressing the startswitch, the sequence of operation of the valveand starter is automatic. A normally closed,momentarily open, stop switch provides a meansof manually stopping the starter when motoringan engine without fuel or in emergencies.

When the external start switch momentarilycloses, a double-pole, single-throw, holding relayin the relay box actuates. This action completeselectrical circuits to the air shutoff valve and thestarter. When the external start switch opens, the

Figure 5-2.-Air turbine starter.

holding relay receives a positive potential throughthe normally closed external stop switch. The relayalso receives a negative potential through theclosed overspeed control provided within thestarter. The relay continues to hold until theexternal stop switch removes the positive potential

or until the closed overspeed control removes thenegative potential.

When high-pressure air is at the normally closedregulating valve inlet and the start switch energizesthe holding relay, the regulating valve opens.Thus, admitting compressed air to the starter. The

Figure 5-3.-Air turbine starting system diagram.

control mechanism regulates the compressed air through the nozzle vanes. The airflows againstto specified conditions. The mechanism senses the blades of the turbine wheel to rotateupstream and downstream conditions, and it the turbine wheel. The reduction gear systempositions the valve butterfly to supply the desired (fig. 5-2), which transmits power to the driveflow. shaft, converts the high-speed, low-torque output

The compressed air enters the inlet port of the of the turbine wheel to low-speed, high-torquestarter and expands as it flows radially inward output.

As compressed air enters the starter inlet portthrough the pressure regulating valve, the turbinewheel rotates, transmitting torque to the drive jawthrough reduction gearing. This torque transmitsto the drive shaft through the splined drive shaft.When aircraft engine speed exceeds the starterdrive shaft speed, the speed of the starter outputshaft (directly connected to the engine drive) alsoexceeds the drive jaw speed. When this conditionoccurs, the pawls begin to ratchet. When theoutput shaft (driven by the aircraft engine) attainsenough speed, centrifugal force releases the pawlscompletely from the drive jaw. This action releasesthe starter from the aircraft engine. When thestarter reaches cutoff speed, the internal overspeedcontrol actuates, breaking the electrical circuit tothe holding relay. Then, the pressure regulatingvalve butterfly closes and prevents compressed airfrom entering the starter.

NOTE: If engine shaft speed fails to exceedstarter driving mechanism speed before the starterreaches cutoff speed, the cutout switch actuates,shutting down the starter.

A starting operation begins by momentarilyclosing the start switch to energize the pressureregulating valve circuit. Sequence of the regulatingvalve and start operation becomes automatic. Thestart continues until engine light-off occurs,and the engine overspeeds the starter drivingmechanism or until output shaft speed reaches thecalibrated cutoff point. In either condition,disengagement of the starter from the engine orinterruption of the supply air to the starter isautomatic. Starter operation also stops when thestop switch momentarily opens, which closes thepressure regulating valve and interrupts the supplyair to the starter.

NOTE: During operation of the starter, youshould stand clear of the plane of rotation of thehigh-speed rotating turbine wheel. Only qualifiedADs should install and service the unit and itspressure regulating valve.

List the three modes of operation of the A-6aircraft constant speed drive/starter unit.

In the air turbine starter, what unit limitsrotational speed?

The air turbine starter converts energy fromwhat type of air to shaft power?

The external stop switch de-energizes the airturbine start control holding relay by

IGNITION SYSTEMS

Learning Objective: Recognize operatingparameters and characteristics of aircraftengine ignition systems.

Three things are necessary to cause a fire—acombustible material (such as aircraft fuel),oxygen, and heat. A fire will not start without allthree, and removing any one of the three puts thefire out. All internal combustion engines use fireto produce mechanical energy, and the piston-typeengine uses the higher degree of fire—anexplosion.

The gas turbine (jet) engine also produces itsenergy through the use of fire. However, itsoperation is considerably different from the pistonengine. Rather than a series of independentexplosions, a jet engine produces a continuousburning fire. Ignition is necessary only during thestart cycle to ignite the fire.

Electronic ignition systems provide internalcombustion for turboprop, turbofan, and turbojetengines. Unlike reciprocating engine systems,timing is not a factor in turbine-power ignitionsystems. All that is needed is a series of sparkswith enough intensity to cause combustion.

The exciter develops voltage of sufficientamplitude to produce a spark. The exciter unitcontains a capacitor or capacitors to develop thevoltage and current necessary to supply a sparkplug (called an igniter). The resultant spark is ofhigh heat intensity, capable not only of ignitingabnormal fuel mixtures but also of burning awayany foreign deposits on the plug electrodes. Theexciter is a dual unit and produces sparks at eachof two igniter plugs.

The igniter plugs are, in general, similar to thespark plugs on reciprocating engines. The maindifferences were changes necessary to operate athigher energies, voltages, and temperatures of jetengines. In general, the igniter plug is larger, moreopen in construction, and the gap is much wider

than spark plugs of familiar design. Figure 5-4shows a typical jet igniter plug.

Control of jet ignition systems is throughrelays or switches that operate automaticallyduring the engine start cycle. Fuel or oil pressureswitches or centrifugal speed switches energize arelay to begin ignition. Ignition stops by actuationof a centrifugal switch at a speed between 45percent and 65 percent of rated engine speed.

CAPACITOR-DISCHARGE IGNITIONSYSTEM

This ignition system has three major com-

assemblies. The exciter unit’s hermetical sealpermanently protects internal components frommoisture, foreign matter, inadvertent maladjust-ments, pressure changes, and adverse operatingconditions. This type of construction eliminatesthe possibility of flashover at high altitude dueto pressure change and gives positive radio noiseshielding. The complete system design, includingleads and connectors, gives adequate shieldingagainst leakage of high-frequency voltage thatinterferes with radio reception of the aircraft. Itsupplies energy to two surface-type spark igniters.

Figure 5-5 is a functional schematic of thesystem. You should refer to this figure whenstudying the theory of operation of a capacitor-

ponents—one ignition exciter and two lead discharge system.

Figure 5-4.-Cross-sectional view of a jet igniter plug.

The ignition system gets its input power fromthe low-voltage dc power supply of the aircraftelectrical system. Its function is to produce highenergy, capacitance-type sparks at the sparkigniters in the engine.

Input power from the low-voltage supplyconnects through a noise filter to a dc motor. Themotor drives a multilobe cam that operates twobreakers, and a single-lobe cam that operates twocontractors. One breaker and one contactorassociate with each side of the system. The twosides are identical. The following descriptionapplies to either side.

When the multilobe cam closes the breaker(fig. 5-5), input current flows through theautotransformer’s primary winding, establishinga magnetic field. Then the breaker opens, the flowof current stops, and the field collapse inducesabout 1,000 volts in the secondary. This voltagecauses a pulse of current to flow into the storagecapacitor through a rectifier that limits the flowto a single direction. You can see that each timethe breaker opens, the capacitor receives a chargeof electricity, and the action of the rectifierprevents any loss of this charge. After 34 suchpulses accumulate a charge on the capacitor, thecontactor closes by mechanical action of thesingle-lobe cam. Then, the capacitor dischargesits stored energy through a triggering transformer,and the energy dissipates at the igniter. The

triggering transformer increases the output voltageof the ignition unit and ensures reliability of thesystem.

The igniter spark rate varies in proportion tothe voltage of the dc power supply, which affectsthe speed of the motor. However, asshaft operates both cams, the storagealways accumulates its energy fromnumber of pulses before discharge.

the samecapacitorthe same

CAUTION

Due to the high voltage and amperage ofthis system, you should use extremecaution around the equipment.

ELECTRONIC IGNITION SYSTEM

The development of more powerful jet enginesdemands a reliable, maintenance-free ignitionsystem. This chapter doesn’t cover all ignitionsystems; rather, the system described representsmost modern systems. An electronic ignitionsystem has an advantage over the capacitor-discharge system; it has no moving parts andbreaker points or contractors that can becomepitted or burned.

Figure 5-5.-Functional schematic of capacitor-discharge system.

The engine ignition system (fig. 5-6) providesthe necessary electrical energy and control to beginengine combustion during aircraft armamentfiring and starting, and for automatic reignitionin case of engine flameout.

Engine Ignition Exciter

The engine ignition exciter is a dual-circuit,dual-output unit that supplies a high-voltage.high-energy electrical current for ignition. Theexciter consists of a radio frequency interferencefilter and two power, rectifier, storage, and outputelements. The exciter mounts on the forward partof the compressor section of the engine.

Engine Control Amplifier

The engine control amplifier is the electroniccontrol center of the engine. It controls thefunction of the ignition system as well as otherengine operational functions. The amplifiermounts on the compressor section aft of theengine front frame.

Engine Ignition Leads and Igniter Plugs

The ignition leads are high-tension cables,which transmit electrical current from the exciterto the igniter plugs. The igniter plugs mount inthe combustion chamber housing.

Figure 5-6.-Jet engine electronic ignition system.

Engine Alternator Stator

The engine alternator stator is an engine-driven single-phase, ac electrical-output unitmounted on the engine accessory gearbox.It supplies electrical power to the engine,independent of the aircraft electrical system. Itcontains three sets of windings. Two windingssupply electrical power to the ignition exciter, andthe third supplies electrical power to the controlamplifier.

Ignition Operation

As you read this section, refer to figure 5-6.With the ignition switch ON, the engine crankingfor starting, and throttle advanced to 10-degreepower lever angle (PLA) position, current flowsfrom the alternator stator to power the controlamplifier. At the same time, the PLA ignitionswitch in the fuel control closes. The gas generator(Ng) speed logic circuit closes the ignition relayto provide ignition when Ng is within the 10 to48 percent Ng range. With the relay closed, itcompletes a circuit from the alternator statorignition windings, through the ignition exciter,to the igniter plugs. Current flows from thealternator, through the control amplifier, to theignition exciter. At the ignition exciter, currentis intensified and discharged as a high-voltageoutput, and conducts through the igniter cablesto the igniters. Current crossing the gaps in theigniters produces a continuous high-intensityspark to ignite the fuel mixture in the combustionchamber. When engine speed reaches 8,500 RPM,Ng and interturbine temperature (ITT) reachesoperating range, a signal from the T5 temperaturedetectors flows through the T5 circuit to thecontrol amplifier ignition logic circuit. The controlamplifier ignition relay opens and ignition ends.Combustion then continues as a self-sustainingprocess.

Ignition automatically reactivates when eithera flameout occurs or when aircraft armamentfires. When T5 temperature drops more than800°F (427°C) from T5 selected by PLA, the T5detectors signal control amplifier T5 flameoutlogic to close the amplifier ignition relay. Thisactivates ignition system operation. Ignitioncontinues until engine operating temperature isagain normal and the 800°F temperature errorsignal cancels, causing the control amplifier to endignition operation.

An armament-firing protection circuit pre-vents flameout from armament gas ingested by

the engine during armament firing. When firingaircraft armament, a signal from the armamenttrigger switch activates the armament-firing logiccircuit in the control amplifier. The amplifier logiccircuit causes ignition operation to activate.Ignition operation ends after a 1-second time delayin the amplifier logic circuit following release ofthe armament firing trigger.

Q1. Why is timing not a factor in jet engineignition systems?

Q2. What unit in an ignition system developsenough voltage to produce a spark?

Q3. At what engine RPM does ignitionnormally stop?

Q4. In the capacitor-discharge ignition system,what total number of pulses must occurbefore the capacitor discharges?

Q5. Engine ignite automatically reactivateswhen either of what two conditions occurs?

ENGINE TEMPERATURECONTROL SYSTEMS

Learning Objective: Recognize operatingconditions and characteristics of aircraftengine temperature control systems.

On most reciprocating engines, the enginecowl flaps control cylinder head temperature(CHT). In turbine-powered engines, turbinetemperature is controlled differently. Enginetemperature is a measure of power, and tempera-ture is a product of fuel consumption. In mostturbine-powered aircraft, then, the pilot selectsdesired power through a mechanical linkage to thefuel control. As the power increases, so does the

temperature, In turbojet aircraft, this also causesan increase in engine speed. The only electricalcircuits required are those to show temperatureand speed.

The engine temperature control system onturboprop engines lets the operator controlturbine inlet temperature and torque through theuse of power and condition levers. These leversconnect to each engine coordinator throughpushrods, sectors, cables, and pulleys. When theengine is operating in the flight range, enginespeed is constant. Engine power is controlled byincreasing or decreasing fuel flow, which resultsin a corresponding change in turbine inlettemperature. The main components of the enginetemperature control system are—

Power levers,

Condition levers,

Engine coordinators,

Temperature datum controls,

Turbine inlet thermocouples, and

Temperature datum switches.

Figure 5-7 shows the block diagram of an enginetemperature control system. You should refer toit while you study this section.

POWER LEVERS

The power levers (one for each engine) canmove separately or together to control enginepower. The range of power lever settings is fromREVERSE (reverse thrust) to MAX POWER(takeoff). Power lever switches within the cockpitpedestal supply electrical power to other systems.A detent at the FLT IDLE position preventsinadvertent movement of the power levers belowFLT IDLE while airborne. To move the powerlevers to the taxi range, you must raise the leversfrom the detent. During a catapult-assistedtakeoff, a retractable catapult grip helps the pilotmaintain the power levers at MAX POWER.

CONDITION LEVERS

The condition levers are located next to thepower levers on the cockpit pedestal. They havefour positions—FEATH, GRD STOP, RUN, andAIRSTART. Switches at each condition lever

position complete electrical circuits for othersystems. The pilot must raise the detent releasehandle of each condition lever to move the leversto different positions. A detent holds the lever atFEATH, GRD STOP, or RUN. When thecondition lever is in the AIRSTART position, thepropeller unfeathers and the engine starting cyclebegins. The lever is held in the AIR STARTposition until the engine speed reaches 100 percentRPM. Then, you release the lever, which springsback to RUN and remains there for normaloperation. When set to RUN, the condition leverpositions the mechanical linkage to open the fuelshutoff valve. A mechanical stop in the pedestalprevents both condition levers from being set toFEATH at the same time. When set to FEATH,the condition lever electrically and mechanicallycloses the corresponding fuel shutoff valve andfeathers the propeller. At GRD STOP, thecondition lever electrically closes the fuel shutoffvalve to shut down the engine.

ENGINE COORDINATORS

The coordinators are mechanical devices thatcoordinate the power and condition levers,propeller, fuel control, and electronic fueltrimming circuit. One engine coordinator mountson each fuel control. The main components ofa coordinator are a variable potentiometer, adiscriminating device, and a cam-operated switch.A scale calibrated from 0 to 90 degrees attachesto the outside case, and a pointer secures to themain coordinator shaft. Pushrods, connectedfrom the coordinator to a cable sector, transmitpower and condition lever movement to thecoordinator. Power lever movement through thecoordinator changes resistance of the variablepotentiometer and changes the temperature datumcontrol temperature reference signal. Thecam-operated switch changes the temperaturedatum control from temperature limiting totemperature controlling with power lever above66-degree coordinator and engine speed above 94percent RPM. Power lever movement transmitsto the coordinator, propeller, and fuel controlthrough a series of rods and levers. With thecondition lever in FEATH, the discriminatingdevice mechanically positions propeller linkagetoward feather and closes the fuel shutoff valve,regardless of the power lever setting.

The temperature datum control consists ofelectronic units that automatically compensate for

Figure 5-7.-Engine temperature control (turboprop) system block diagram.

changes in fuel density, manufacturing tolerances power lever is below 66-degree coordinatorin fuel controls, and variations in engine fuelrequirements between engines. With the powerlever above 66-degree coordinator (temperaturecontrolling range) and the TEMP DATUM switchin AUTO, the temperature datum controlcompares the actual turbine inlet temperaturesignal and desired temperature reference signal.If there is a difference greater than 1.9°C (4.5°F),the control electrically signals the temperaturedatum valve to reduce or increase fuel flow to theengine. This action brings the turbine inlettemperature to the desired value. A dampingvoltage goes back to the control from a generatorwithin the temperature valve motor, preventingovercorrection and stabilizing the system. Whenengine speed is above 94 percent RPM and the

(temperature limiting range), the normal limitingtemperature automatically becomes 978°C(1,792°F). However, when engine speed is below94 percent RPM, regardless of power leverposition, the limiting temperature is 830°C(1,524°F). This prevents high turbine inlettemperature during starting and acceleration whenthe compressor bleed valves are open.

Dual-unit thermocouples are mounted radiallyin the turbine inlet case of each engine. Thejunction portion of the thermocouples protrudesthrough the case to sense the gas temperaturebefore the gas enters the turbine section. Fourleads, two of Chromel and two of Alumel,connect to each thermocouple to form twoindependent parallel circuits. One circuit connects

to the cockpit turbine inlet temperature indicator.The second circuit supplies the temperature datumcontrol with temperature signals for the electronicfuel trimming circuit. As the gases heat thethermocouples, they generate an electromotiveforce that goes to the cockpit indicator andthe temperature datum control. Because thethermocouple connections are in parallel, thesignal they send is the average temperature of thethermocouples. If one parallel circuit fails, theother circuit continues to operate normally.

TEMPERATURE DATUM SWITCHES

The left and right engine temperature datum(TEMP DATUM) switches are on the engine con-trol panel in the cockpit. Each switch has AUTOand NULL positions. When the switch is in AUTO,the engine RPM is above 94 percent, and the enginecoordinator is above 66 degrees, the temperaturedatum control compares the turbine inlet tempera-ture to a reference temperature. If the temperaturesdiffer, the temperature datum control electricallysignals the temperature datum valve to bypassmore or less fuel from the engine to bring turbineinlet temperature to the selected value.

If the electronic fuel trimming circuit malfunc-tions, position the TEMP DATUM switch toNULL. The circuit de-energizes and the fuelcontrol, through movement of the power lever,controls turbine inlet temperature. Overtempera-ture protection is not available.

List the four positions of the E-2 aircraftcondition lever.

When is the temperature datum control inthe temperature controlling range?

At what RPM is the engine limitingtemperature 830°C?

Thermocouple leads are made of

List the main components of an enginecoordinator.

ENGINE START CONTROL SYSTEM

Learning Objective: Recognize operatingparameters and characteristics of aircraftengine starting systems.

In this section, you will learn about enginestarting, engine ignition, and engine fuel tempera-ture starting systems. By combining these systems,you will understand how they interrelate. Theengine start control system covered in this sectionis the one found in the P-3C aircraft.

MAJOR COMPONENTS

Major components of the engine start controlsystem include the air turbine starter, thespeed-sensitive control, the ignition exciter, theengine fuel pump and filter, the fuel control, thefuel nozzles, the fuel control relay, the startingfuel enrichment valve, the temperature datumvalve, the drain valves, and the compressor bleedair valves. You should refer to figure 5-8 and totable 5-1 as you read about these components.

Air Turbine Starter

You have learned that the air turbine starteris pneumatically driven and mechanicallyconnected to the engine through a gearbox. Theair turbine starter operates on compressed airfrom an external gas turbine compressor (GTC),internal auxiliary power unit (APU), or bleed airfrom an operating engine. The compressed airgoes through a manifold, an engine isolation bleedair valve, and a starter control valve into thestarter’s turbine. The engine start switch, locatedin the cockpit, controls the opening of the startercontrol valve, which allows compressed air toenter the air turbine starter. The speed sensitivecontrol, through a holding solenoid, holds theengine start switch on until engine speed reaches65 percent RPM.

Speed-Sensitive Control

The speed-sensitive control, located on theengine, contains internal switches that activate atthree predetermined intervals. These intervals arerelative to the engine’s normal speed—16 percent,65 percent, and 94 percent of engine RPM.Activation of these switches controls manyoperations in the engine start cycle.

Figure 5-8.-Engine start control system.

Ignition Exciter

The ignition exciter (discussed earlier) is a dualelectronic ignition unit that uses 28-volt dc fromthe ignition relay. The exciter then steps up thevoltage to a proper level for firing the igniterplugs. The exciter unit contains two identicalcircuits, each one independently capable of firingits own igniter plug. The speed-sensitive controlenergizes the ignition relay so the exciter is inoperation between 16 and 65 percent of engineRPM.

Engine Fuel Pump and Filter

The fuel pump and high-pressure filterassembly mount on the rear of the accessoriescase. This assembly consists of a centrifugal boostpump, two gear-type pressure elements, and ahigh-pressure filter.

Fuel entering the pump assembly passesthrough the centrifugal boost pump, which raises

the pressure to a minimum value. Fuel passesthrough the low-pressure filters before going tothe secondary element. There is a differentialpressure switch connected across the inlet andoutlet of the filters. If the pressure differentialexceeds 7.5 PSI, the switch closes and completesa circuit to a filter light at the flight deck. Fuelthen flows to the primary element and throughthe high-pressure filter assembly before enteringthe fuel control. Both the low- and high-pressurefilters have bypass valves that open if the filtersbecome clogged.

The capacity of the pump’s primary elementis 10 percent greater than that of the secondaryelement. If the primary element were to fail, thesecondary element would provide enough flow tooperate the engine.

During engine starting, the elements operatein parallel to provide enough fuel flow at lowRPM; above 65 percent, they operate in series.Parallel operation occurs during starting whenengine speed is between 16 percent and 65 percent

Table 5-1.

RPM, If both elements are operating properly,the paralleling light will be on between 16 percentand 65 percent RPM only. If the secondaryelement fails, the light never comes on; if theprimary element fails, the light is on above 65percent RPM.

Fuel Control

The fuel control is on the accessories drivehousing and mechanically links to the coordi-nator. The fuel control provides a starting fuelflow schedule that, in conjunction with thetemperature datum valve, prevents overtempera-ture and compressor surge.

The fuel control schedule is 20 percent richerthan the nominal engine requirements to accom-modate the temperature datum valve. The valvebypasses 20 percent of the control output whenin the null position. This excess flow gives thetemperature datum valve the capacity to add aswell as subtract fuel. The valve is then able tomaintain the temperature scheduled by thecoordinator and the temperature datum control.

The fuel control includes a cutoff valve forstopping fuel flow to the engine. It actuates eithermanually or electrically. During engine starts, thecutoff valve remains closed until the enginereaches 16 percent RPM. The speed-sensitivecontrol then opens the cutoff valve, permittingfuel to flow to the engine.

Fuel Nozzles

The fuel output from the temperature datumvalve flows through the fuel manifold to the sixfuel nozzles. Fuel flows through both the primaryand secondary nozzle orifices during normaloperation. At low fuel flow rates, fuel flowsthrough the primary orifice only.

Fuel Control Relay

The fuel control relay is a fail-safe-type relaythat energizes when the FUEL and IGNITIONswitch is off, or when the propeller is feathered.With this arrangement, the engine can still operatewith an electrical power failure during flight. Thepilot shuts the engine down by placing the FUELand IGNITION switch in the OFF position, orby feathering the propeller. Power then goes tothe fuel control shutoff valve, which closes andstops all fuel to the engine.

Starting Fuel Enrichment Valve

The primer switch (with two positions, ONand spring-loaded OFF) operates the fuelenrichment valve, providing increased fuel flowduring engine starting. The primer switch mustbe in the ON position and held there before theengine reaches 16 percent RPM. Further, it mustremain on until the fuel control shutoff valveopens at 16 percent RPM or enrichment will notoccur. The enrichment (primer) valve closes whenfuel pressure in the fuel manifold reaches 50 PSI.Fuel enrichment is needed only in very coldclimates.

Temperature Datum Valve

The temperature datum valve is locatedbetween the fuel control and the fuel nozzles. Itis a motor-operated bypass valve that respondsto signals from the temperature datum control.If the power lever positions are between 0 degreesand 66 degrees, the valve remains in null and theengine operates on the fuel flow scheduled by thefuel control. The valve remains in null unless thetemperature datum control signals it to limitturbine inlet temperature. The valve then reducesthe fuel flow (up to 50 percent during starting,20 percent above 94 percent RPM) by returningthe excess to the fuel pump. When turbine inlettemperature is at the desired level, the temperaturedatum control signals the valve to return to thenull position.

In power lever positions between 66 degreesand 90 degrees, the temperature datum valve actsto control turbine inlet temperature to apreselected schedule corresponding to power leverposition. This is the temperature controllingrange. In this range the temperature datumcontrol may signal the valve to allow more (highertemperature desired) or allow less (lowertemperature desired) fuel to flow.

Drain Valves

A spring-loaded, solenoid-operated manifolddrain valve is located at the bottom of the fuelmanifold. It drains the fuel manifold when fuelpressure drops below 8 to 10 PSI. This actionminimizes the amount of fuel dropping into thecombustion liners while the engine unit is beingstopped.

Compressor Bleed Air Valves

The fifth and tenth bleed air valves release airfrom the compressor to reduce the compressorload during engine starts. During starting, thebleed valves are open up to 94 percent RPM. At94 percent RPM, the speed-sensitive valve portscompressor-discharge air to close the bleedvalves.

ENGINE START CYCLE OPERATION

The following sequence of events is typical ofa normal engine start cycle. While readingthis section, you should assume that externalcompressed air and electrical power are beingapplied to the aircraft. Also, assume that all othersystem switches are in the proper position for anengine start. Refer to figure 5-8 and table 5-1throughout this discussion.

The operator places the ENGINE STARTSELECTOR switch to the engine number 1position. Position FUEL and IGNITION switch,engine 1 to the ON position, de-energizing the fuelcontrol relay. This allows power to pass throughthe contacts of the fuel control relay, the 16percent speed-sensitive control switch, and thefuel manifold pressure switch to energize thetemperature datum relay.

When the operator depresses the ENGINESTART switch, the current flows through theengine start switch, engine start selector switch,starter control valve, and speed-sensitive control65 percent switch to ground. Current also flowsthrough the engine start switch holding coil toground through the same 65 percent switch in thespeed-sensitive control.

With power applied to the starter controlvalve, the valve opens. This allows compressedair to flow to the air turbine starter. It also closesthe contacts of the air valve position switch. Theyellow starter valve lights illuminate to show theoperator the starter control valve is open. The airturbine starter now causes engine rotation.

If fuel enrichment is needed, depress thePRIMER switch and hold in the ON position untilthe engine reaches 16 percent. Power then goesthrough the primer relay contacts and thetemperature datum relay contacts, energizing theprimer valve solenoid.

When the engine reaches 16 percent RPM, thespeed-sensitive control mechanically actuates the16 percent switch from 16 percent to 65 percent.Power then goes through the 16 percent switchcontacts, energizing the ignition relay. Power

then flows through the ignition relay contacts,energizing the fuel pump paralleling solenoid, thedrip valve solenoid, and the ignition exciter.Power also goes to the fuel control shutoff valve,allowing fuel to enter the engine fuel manifold.Extra fuel for fuel enrichment (if used) flows tothe temperature datum valve. The temperaturedatum relay remains energized by a holding circuitconsisting of the lower temperature datum relaycontacts and the fuel manifold pressure switch.

As engine speed increases, the fuel pressureincreases. When fuel manifold pressure reaches50 PSIG, the fuel manifold pressure switch opens,de-energizing the temperature datum relay,stopping fuel enrichment. When the secondaryfuel-pump pressure exceeds 150 PSIG, a parallel-ing light illuminates showing parallel operationof the fuel pumps to the operator.

At 65 percent RPM, the speed-sensitivecontrol 65 percent switches open to de-energizethe ignition relay. The ignition relay removespower from the ignition exciter, the fuel pumpparalleling solenoid, and the drip valve solenoid.The fuel pumps now operate in series, and fuelpressure now holds the drip valves closed. Thestarter control valve and the engine start switchlose their common ground, and current flowceases through those circuits. The starter controlvalve closes, stopping airflow to the air turbinestarter, and the engine start switch opens. The airvalve position switch opens, causing the startercontrol valve light to go out. This completes theengine start cycle.

When engine speed increases above 94 percent,contacts in the speed-sensitive control (circuit notshown) de-energize the temperature datum valvetake solenoid. This reduces the fuel take capabilityfrom 50 percent to 20 percent. The fifth and tenthstage bleed air valves also close now. When thepower lever advances above 66 degrees ofcoordinator travel, the temperature datum systemswitches from temperature limiting range totemperature controlling.

During a P-3 engine start, at what percentRPM does the start control switch re-energize?

What are the three switches in the speed-sensitive control?

Q3. How much greater is the capacity of theprimary pump than the secondary pump?

Q4. Why is the fuel control schedule 20 percentricher than nominal engine requirements?

Q5. What unit minimizes the amount of fueldropping into the combustion liners duringengine shutdown?

POWER PLANT ANTI-ICING ANDDEICING SYSTEMS

Learning Objective: Recognize operatingprinciples and characteristics of aircraftpower plant anti-icing and deicing systems.

Naval aircraft deicing lets you fly in any typeof weather. When there is enough moisture in theair and you encounter freezing temperatures, youmust protect the power plant against ice buildup.There are two electrical systems that do this—the anti-icing and deicing systems. Anti-icingsystems prevent ice from forming and deicingsystems remove ice that has already accumulated.

Many types of anti-icing systems are usedtoday. All systems use heated air from the engineto perform the anti-icing function. The use ofheated air causes engine power loss; so useanti-icing only when necessary. In some aircraft,a reversible electric motor opens and closes an airvalve to supply the needed air. In other aircraft,an electrical solenoid positions a pneumatic valveto allow regulated heated air into the engineanti-ice system.

Where the mission of the aircraft dictates thatit fly routinely in adverse weather conditions, afail-safe anti-ice system is used. Fail safe meansthe solenoid-actuated air valve electrically actuatesclosed. If the switch is turned on, or if electricalpower fails, the valve is spring loaded to the openposition. Some systems anti-ice the complete inletduct; while in other systems, only the guide vanesare anti-iced.

GUIDE VANE ANTI-ICING SYSTEM

There are a variety of engine anti-icingsystems in use today. The system covered in this

TRAMAN is representative of several systemsdesigned for Navy aircraft. Look at figure 5-9.Here, you see that the electrical portion of thecircuit serves only to turn the system on or off.

The guide vanes of a turbine-powered enginedirect the flow of inlet air into the compressorsection. At this point, the air is coldest and mostsubject to icing. The biggest problem caused byice forming here is blockage of inlet air, causingair starvation and thus engine failure. Also, thereis a possibility that chunks of ice can be inductedinto the engine. Therefore, turn on the anti-icingsystem at the first indication of any icing conditionor before entering an icing condition. Normally,icing doesn’t occur in supersonic flight becausefriction of the aircraft passing through the aircreates enough heat to prevent ice formation.

The anti-icing valve is a solenoid-operatedbleed valve. With no electrical input to thesolenoid, the bleed valve closes, and there is noanti-icing airflow through the valve. When theengine is operating with the valve solenoid de-energized, the main poppet will remain in theclosed position. When the solenoid energizes, thesolenoid valve unseats and permits air pressurewithin the main poppet to escape through theoverboard vent.

With pressure decreasing in the poppet valvebody, inlet pressure on the main poppet valve faceovercomes spring tension and raises the valvefrom its seat. This permits high-pressure air todischarge through the outlet of the valve to theanti-icing manifold on the engine. The regulatingpiston and spring valve assembly control dischargeair from the anti-icing valve to a preset pressure.

PROPELLER DEICING SYSTEMS

One method of preventing excessive accumu-lation of ice on the propeller blades of turbopropor reciprocating engines is using electric heaters.Figure 5-10 is a simplified schematic diagram ofa system for a two-engine aircraft.

The propeller deicing system consists of athree-position, two-speed selector switch(propeller deicer switch) and an indicator light,a two-speed timer, and two propeller deicer relays(one for each propeller). Also, included (for eachpropeller) are the brush pad bracket assembly,slip-ring assembly, the aft portion of the propellerassembly, and a neoprene rubber heating elementand connector for each blade. Abrasion stripsprotect the blade heaters. The deicing systemoperates at either a slow cycle of 40-75 secondson, 120-225 seconds off; or a fast cycle of 17-22

Figure 5-9.-Inlet guide vane anti-icing system.

Figure 5-10.-Electrical deicing for a propeller system.

seconds on, 51-66 seconds off. The icing condi-tions during flight determine the switch position.

Setting the selector switch to SLOW or FASTpermits dc power from the essential bus toenergize the deicer timer motor and turn on theindicator light. Resistances in the timer determinethe speed of the timer motor. The motor, throughreduction gears, causes the camshaft to rotate.This rotation positions the cam switchesalternately between the right and left contacts.Current flows through these contacts to cycle theirrespective propeller deicer relays. With the relaysbeing energized alternately, current from thethree-phase generator ac buses, through propellerdeicer circuit breakers, flows to the propellerbrush pad bracket assemblies. Carbon brushescontact the copper slip rings, transmitting acpower through the slip rings to the blade heatingelements. Placing the selector switch in the OFFposition stops the propeller deicing operation, andthe indicator light goes out.

The propeller deice timer is a two-speed,automatically controlled timer. It regulates, incycles, the time duration and sequence of electricalimpulses to the propeller blade heating elements.The unit is located in a moistureproof, airtightcase, which isolates the unit from temperatureextremes and vibration. The deice timer consistsof a fractional horsepower, constant-speed dcmotor, including reduction gear, camshaft withthree cams, three cam switches, two fixedresistors, and variable resistor. The unit alsoincludes a filter to minimize radio interference.

With the propeller deice switch set to FAST,direct current flows from the left dc bus, throughthe propeller deicer circuit breaker and switch, tothe timer. This current follows two paths in thetimer. One path, from pins E and F that connectin the timer, directs the current flow to the controlcam switch. The other path, from pin G, directsthe flow through the variable resistor and onefixed resistor to the timer motor, the filter, andto ground. The adjustment of the variable resistordetermines the speed of the motor. The motor,through the 3,000 to 1 reduction gear, rotates thecamshaft and cams. Two single-lobed shift camsand a single two-lobed control cam are on thecamshaft. Positioning on the camshaft is so thetwo single-lobed shift cams are on either side ofthe two-lobed control cam. As the control camrotates, it alternately makes and breaks its rightand left cam switch contacts. This permits theflow of current to the shift cam switches. As thecurrent flows to the other cam switch, rotationof the single-lobed cam makes and breaks the shift

cam switch contacts. This action cycles first theright and then the left propeller deicer relays.

With the propeller deicer switch set to SLOW,the operation is the same as the fast cycle withthe one exception. Direct current enters the timerthrough a different pin (pin H) in the plug andflows through the two fixed resistors and thevariable resistor to the motor. (Dc power to thecontrol cam switch is through the same pins asfor the fast cycle.) Because of the increase inresistance, the motor operates at a slower speed.Thus, with motor speed reduction, rotation of thecamshaft, through the reduction gear, is slower,and the timer now functions at the slower cycle.

Several aircraft have an anti-icing and deicingsystem that prevents the formation of ice (anti-icing) on the forward portion of the propellerspinner. The system removes any ice formation(deicing) from the blades and cuffs, aft portionof the spinner, and spinner islands. This systemoperates similarly to the system describedpreviously, except that the anti-icing elements areon continuously and the deicing elements cycle.Figure 5-11 shows the location of the heatingelements. The system usually contains a safetyfeature for testing the propellers on the ground.This feature provides a low voltage to the heatingelements, which prevents damage to the prop fromoverheating.

In supersonic flight, icing doesn’t occur forwhat reason?

What are the two types of anti-icing/deicingsystems used on aircraft?

At what point in the engine is thecoldest and most likely to ice?

air the

What is the difference between propellerdeicing and anti-icing?

Figure 5-11.-Propeller heating elements.

FIRE WARNING ANDEXTINGUISHING SYSTEMS

Learning Objective: Recognize operatingconditions and characteristics of aircraftengine fire warning and extinguishingsystems.

Some turbine engines operate at temperaturesof more than 1,000°C. Fuel and oil lines runwith in a few inches of these extreme temperatures.For this reason, you must closely monitor the

engine, and immediately take corrective actionwhen an abnormal condition occurs. Performanceof precision work is necessary when maintainingfire warning systems to ensure their reliability. Anundetected fire may cost the lives of the aircrewand possibly millions of dollars in aircraft andequipment.

In multiengined aircraft, a fire warning usuallydictates that the engine be shut down and the fireextinguished. The least that can happen if thereis an erroneous fire warning is the aircraft willabort its assigned mission.

WARNING SYSTEM

The engine fire detector system is an electricalsystem for detecting the presence of fire or dan-gerously high temperatures in the engine(s) areas.The system for each engine consists of a controlunit, test relay, signal lamp, test switch, andseveral sensing elements. The detect or uses acontinuous strip of temperature-sensing elementsto cover the paths of airflow in the engine com-partment. The same engine fire warning systemsare in multiengined aircraft, one for each engine.

Look at figure 5-12. It shows the electricalschematic for a fire warning system that is

Figure 5-12.-Typical engine fire warning circuit schematic.

representative of systems found in modern Navyaircraft. The system is of the continuous-element,resetting type. The sensing element consists of twoconductors separated by a semiconductor. Theouter conductor is at ground potential, and thecenter conductor connects to an amplifier inputin the fire detector control unit. The semi-conductor portion of the element has an inversetemperature coefficient; as the temperatureincreases, the resistance of the sensing elementdecreases.

The fire detector control unit continuouslymonitors the electrical resistance of the firedetector system’s sensing element. The controlunit activates a fire warning light (in some unitsan audible warning is also given) when one of thefollowing conditions exist:

1. The sensing element resistance decreases tothe predetermined level (established by the firealarm setting) due to an increase in temperature.

2. The sensing element resistance decreasesat a predetermined rate due to the rate oftemperature increase in the sensing element.

NOTE: In the past, flights have aborted andcrewmen actually ejected because of a fire warninglight illumination caused by a short circuit in thesystem. Modern aircraft fire warning systems havea short discriminator circuit that differentiatesbetween an overtemp (fire) and a short circuit.If there is a short in this system, the fire warninglight will NOT illuminate.

The rest of this section contains a descriptionof a dual jet engine aircraft fire warning system.Both systems operate identically and are similarin operation to fire warning systems found onother Navy aircraft. The left side is discussed inthis section.

Look at figure 5-12. Electrical power from theL FIRE DET circuit breaker supplies the left firedetector and sensing element circuit through pinA of the detector control unit. The left fire-sensingelement loop connects to pins L and C of thedetector control unit. This completes the sensingcircuit through normally closed contacts of thede-energized relay K1. At normal temperatures,the sensing element resistance is high, reverse-biasing diode CR1. This allows current throughresistor R2 to turn transistor Q1 on. With Q1 on,the base current at Q2 is shut off, turning Q2 off.With Q2 off, the current flows into the base ofQ3. This turns Q3 on and Q4 off. Transistors Q3and Q4 are relay-driving. With Q3 on, relaycoil K2-A energizes, opening K2 contacts,

de-energizing the warning circuit and turning outthe L FIRE warning indicator lights. Normaltemperature conditions energize relay coil K2-A.

When the temperature rises, the sensingelement resistance decreases, shunting the currentfrom resistor R2 through diode CR1, turningtransistor Q1 off. This switches transistor Q2 on,transistor Q3 off, and transistor Q4 on. Withtransistor Q4 on, relay coil K2-B energizes, andrelay coil K2-A de-energizes. This transfers(switches) contacts of relay K2, energizing thewarning circuit. Then, 28 volts of dc powers thewarning circuit through pin K of the detector,turning on the L FIRE warning indicator lights.If the temperature drops, the warning circuitde-energizes, causing the fire warning indicatorlights to go out.

Transistors Q5 and Q6 make up the shortdiscriminator circuit. Its principle of operationworks on the rate of change of sensor resistance.In a fire, the resistance rate of change is slow. Inan electrical short condition, the resistance ratedrops abruptly. There are two timing circuits—one is made up of resistors R5 and R6 andcapacitor C2, and the other of resistors R15 andR16 and capacitor C5. These circuits are preset.This allows a slow change of sensor resistance tolet transistor Q1 switch transistor Q2 beforetransistor Q5 switches transistor Q6. A fast changeof sensor resistance allows transistor Q5 to switchtransistor Q6 before transistor Q1 can switchtransistor Q2. If transistor Q6 switches first,transistor Q2 is unable to switch. This action holds

Figure 5-13.-Container and dual valve assembly.

the rest of the circuits in the de-energized (noalarm) condition. If transistor Q2 switches first,the contacts of relay K2-B remove transistor Q6from the discriminator circuit.

EXTINGUISHING SYSTEM

The fire-extinguishing system on many aircraftprovides control for fires within the engines andnacelles. The extinguishing system is an electricallycontrolled, high rate discharge (HRD) system.Normally, you, the AE, will troubleshoot theHRD system electrical circuits only.

The extinguishing agent container is a weldedsteel sphere, 9 inches in diameter and cadmiumplated for corrosion prevention (fig. 5-13). Eachcontainer has a charge of bromotrifluoromethaneand is pressurized with nitrogen. Bromotri-fluoromethane (CF3Br in liquid form) is nontoxicand classed as a nonpoison; however, it readilyvaporizes, is odorless, and can be harmful. Youshould handle it carefully. Don’t let it comein contact with your skin; frostbite or low-temperature burns may result.

As it leaves the system, vaporization changesthe liquid to a gas, displacing the oxygen within

the compartment. The lack of oxygen in thecompartment won’t support combustion or life.Don’t enter an area where bromotrifluoromethanehas been discharged until it is safe to do so.

Each container has two valve assemblies fordischarging the agent. Each valve assemblycontains an explosive, electrically controlledcartridge. When a fire-extinguishing dischargeswitch actuates, it completes a circuit (fig. 5-14).The cartridge electrically fires, allowing the slugto rupture the frangible disk in the neck of thecontainer. When the frangible disk is ruptured,nitrogen pressure expels the extinguishing agent.

CAUTION

The cartridge mentioned in the aboveparagraph is an explosive device. Useextreme caution when working on or nearthis device.

Refer to “Cartridge Activated Devices(CADS) Manual,” NA 11-100-1. You needto have an ordnance certification beforeworking on this system.

Figure 5-14.-Engine fire-extinguishing circuit schematic.

Q1. The circuit that differentiates between anovertemp (fire} and a short circuit in afirewarning system is known as a

Q2. The fire warning element has an inversetemperature coefficient. What does thisstatement mean?

Q3. In the fire warning circuit (fig. 5-12), whathappens when Q6 switches before Q1?

Q4. What chemical is used in the engineextinguishing system HRD bottles?

FUEL TRANSFER SYSTEMS

Learning Objective: Recognize operatingparameters and characteristics of aircraftfuel transfer systems.

The F/A-18 aircraft fuel transfer system isdescribed in this section of the TRAMAN. TheF/A-18 carr ies fue l internal ly in fourinterconnecting fuselage (bladder) tanks and twointernal wing (wet) tanks. External fuel is carriedin three 315- or 330-gallon tanks. All tanks maybe refueled on the ground through a single-pointrefueling receptacle. Airborne, they can berefueled through the in-flight refueling probe. Theinternal wing tanks—tank 1 and tank 4—aretransfer tanks. The tanks are arranged so internalfuel gravity transfers (at a reduced rate) even ifthe transfer jet ejectors fail. Regulated enginebleed air pressure is used to transfer fuel from theexternal tanks and also provides a positivepressure on all internal fuel tanks. Float-type fuellevel control valves control fuel level duringrefueling of all tanks. Fuel level control shutoffvalves in tanks 1, 2, 3, and 4 control fuel levelsduring external fuel transfer. During internal wingtransfer, fuel level control shutoff valves controlfuel levels in tanks 1 and 4.

Fuel level sensors control the fuel level in tanks2 and 3 (engine feed tanks) during fuel transferfrom tanks 1 and 4. All internal and external fuel(except engine feed tanks) can dump overboardthrough flame arrester protected outlets in eachvertical fin. All internal fuel tanks vent throughoutlets in the vertical fins. The external tanksvent overboard through pressure relief valvesin the individual external tanks. A fuel quantityindicating system provides fuel quantity indi-cations in pounds.

FEED TANKS

The internal transfer system design keeps fuelin the feed tanks (tanks 2 and 3) at all enginepower settings. Fuel being transferred from tanks1 and 4 flows to the feed tanks, where the fuellevel is maintained by fuel level sensors.

WING TANKS

Wing tanks transfer fuel to tanks 1 and 4.They are an integral part of the wing structure.Wing tanks are sealed by filling channels withsealant injected through fittings on the outside ofthe wings.

TRANSFER MOTIVE FLOW

The internal fuel transfer system is poweredby motive flow pressure, generated by twoairframe-mounted accessory drive-mounted(AMAD) motive flow/boost pumps contained ina closed loop circuit.

Flow pressure passing through the left andright engine motive flow check valves combinesto create transfer motive flow pressure. Transfermotive flow pressure operates the wing transferejectors and tanks 1 and 4 transfer jet ejectors.Transfer motive flow pressure also closes therefuel/defuel shutoff valve and defuel valve.

WING TRANSFER

Wing transfer starts when the fuel level dropsbelow the high level pilot valves in tanks 1 and/or4. This opens the fuel level control shutoff valve,allowing the wing transfer jet ejectors to transferfuel through the refuel/transfer manifold. As thefuel level in the wings drops below the transfermotive flow pilot valves, the wing transfer motiveflow shutoff valves close, stopping transfer fromthe wings. Refuel/transfer check valves in tanks

1 and 4 keep fuel from entering the refuel line andthe feed tanks.

As fuel from the hot fuel recirculation systemincreases wing fuel, the wing motive flow pilotvalve and motive flow shutoff valve open,allowing wing transfer. A flapper check valve ateach wing ejector inlet prevents transfer from onewing to the other.

Transfer motive flow pressure to each wingejector is controlled by the normally open wingdamage shutoff valve in tank 4. If wing damageoccurs, the pilot sets the INTR WING switch toINHIBIT on the cockpit EXT LT control panel.This closes the wing damage shutoff valve,preventing loss of fuel through a wing transfermotive flow line and stopping wing transfer.

If normal wing transfer does not occur, allwing fuel can be gravity transferred to tank 4 with5 degrees of roll. A check valve in each gravitytransfer line prevents reverse flow.

FUSELAGE TRANSFER

Fuselage transfer (transfer from tanks 1 and4 to tanks 2 and 3) starts when the fuel levelsensors open the transfer shutoff valves in tanks2 and 3.

Fuel flow (transfer) from tanks 1 and 4transfer jet ejectors enters a fuselage transfermanifold that supplies fuel to a transfer shutoffvalve in feed tanks 2 and 3 and to the dump valve.Transfer from the tank 1 or 4 ejector alone isenough to keep both the feed tanks full atmaximum engine demand.

When a transfer tank is empty, the transferpilot valve and transfer shutoff valve close,preventing transfer motive flow pressure fromentering the transfer line. Fuel transfer betweenfuselage transfer tanks is prevented by checkvalves in the inlets of each transfer jet ejector.

Fuel levels in the feed tanks are maintainedby a fuel level sensor and transfer shutoff valvewithin each feed tank.

If transfer from tanks 1 and 4 to the feed tanksfails, fuel gravity transfers through an alwaysopen interconnecting line in the bottom of tank4 and through an orifice in the interconnect valvein tank 1. If motive flow pressure to either enginefuel boost jet ejector or engine fuel turbine boostpump is interrupted, tanks 2 and 3 fuel gravityfeed through the ejector to the engine.

If the left engine shuts down and/or leftmotive flow boost pressure is lost, the tank 1 andtank 2 pressure-operated interconnect valves open.This allows fuel to gravity feed from tanks 1 and

2 to tank 3. Reverse flow from tank 3 is preventedby a flapper check valve on tank 3 interconnectvalve.

If the right engine is shut down and/or rightmotive flow boost pressure is lost, the tank 3pressure-operated interconnect valve opens. Tank4 interconnect line is always open. Fuel gravityfeeds from tanks 3 and 4 to tank 2. Reverse flow isprevented by the flapper check valve on tank 2interconnect valve and tank 3 flapper check valve.

On some series aircraft, motive flow pressureto the tank 1 fuel low-level shutoff and pressure-operated interconnect valve controls gravity feedto tank 2. The fuel low-level shutoff valveenergizes closed when fuel in tank 2 is below 700to 900 pounds. The closed fuel low-level shutoffvalve stops motive flow fuel to the pressure-operated interconnect valve, allowing the flapperto swing open. Once the flapper is open, fuel intank 1 can gravity feed to tank 2. The positions ofthe tank 2 and tank 3 pressure-operated intercon-nect valves are tested using the fuel check panel.

CENTER OF GRAVITY (CG)CONTROL SYSTEM

The fuel quantity gauging intermediate devicecontinuously compares the ratio of fuel betweentanks 1 and 4. When tank 1 transfers fuel at afaster rate than tank 4, the transfer control valvein tank 1 is energized closed, stopping transferfrom tank 1. Tank 1 will not resume transfer untiltank 4 transfers (depletes) fuel to within theparameters defined by the intermediate devices.If fuel distribution in tanks 1 and 4 has causedthe aircraft CG to be further aft than desired, aCG caution will display on the left digital displayindicator.

Once tank 1 depletes below 150 pounds offuel, the intermediate device will stop monitoringtanks 1 and 4 fuel ratios. Tank 1 will then transferfuel until the transfer pilot valve closes the shutoffvalve.

Q1. In the F-18 fuel transfer system, what fueltanks feed the engines?

Q2. When does fuel transfer from the wingstart?

Q3. When does the fuel transfer center ofgravity control system energize?

OIL TEMPERATURE CONTROLSYSTEM

Learning Objective: Recognize operatingprinciples and characteristics of aircraftengine oil temperature control systems.

The cooling capacity of the oil cooler system(fig. 5-15) in an aircraft depends on the airflowthat passes through the cooler. Airflow iscontrolled by an oil cooler door actuator, whichvaries the oil cooler air exit duct.

The door actuator is a split-field, reversibledc motor. It includes a magnetic brake forstopping it quickly when it reaches the limits oftravel. The control switch has four positions—OPEN, CLOSE, AUTOMATIC, and OFF. In theOPEN or CLOSE position, electrical power goesto the actuator, which opens or closes the oilcooler door. When the switch is in AUTO-MATIC, a thermostat controls the actuator.

The thermostatic control unit is in the oilreturn line. The unit contains two floating contactarms and a central contact arm that actuates bya bimetallic coil immersed in the oil return line.

Figure 5-15.-Oil temperature control circuit.

One of the floating contacts is in the door opencircuit and the other is in the door closed circuit.The two arms rest on a cam, which a smallmotor constantly rotates. Thus, the floatingcontacts are constantly vibrating toward thecentral contact.

When oil temperature rises above normal, thethermostatic element causes the central contactto move toward the door o p e n contact. Asthe contact vibrates, it intermittently closes thedoor open circuit. As the actuator intermittentlyenergizes, the door slowly opens. When the oiltemperature returns to normal, the central contactmoves back to a neutral position.

When oil temperature falls below normal, thecentral contact moves in the opposite direction,closing the door. To prevent excessive huntingof the system, a tolerance is maintained byan adjustment of the cam on the floatingcontact.

When the oil temperature rises high above thenormal value, the central contact lifts the floatingcontact clear of the cam, completing a continuouscircuit. The door then moves to the full openposition where a limit switch in the actuatorbreaks the circuit.

Figure 5-16 shows another type of engine oiltemperature regulator. This regulator has amercury-filled thermostat, and relays auto-matically control the position of the engine oilcooler doors. When the engine temperature is lowrequiring more heat, the two relays energize,allowing the oil cooler door to close. As thetemperature increases, the thermostat completesa path to ground, bypassing the relay coils andde-energizing them. Power then goes through thecontact of one of the relays, opening the actuatorcoil and causing the oil cooler door to open andreduce engine temperature.

VARIABLE EXHAUST NOZZLECONTROL SYSTEM

Learning Objective: Recognize operatingconditions and features of aircraft enginevariable exhaust nozzle control systems.

As you read this section, refer to the simplifiedschematic diagram of a typical variable exhaust

Figure 5-16.-Automatic oil temperature control circuit.

nozzle (VEN) control system (fig. 5-17). This isthe control system for the F-18 aircraft, andit is a converging-diverging nozzle system.When operating, this system varies the exhaustescape area size to obtain desired thrust, whilemaintaining safe operating conditions throughoutthe engine. The VEN control system consists ofelectrical, hydraulic, and mechanical componentsthat position the VEN while maintaining exhaustgas temperature (EGT).

MAJOR COMPONENTS

The VEN control system has nine majorcomponents. As you read about these com-ponents, locate each component in figure 5-17 asyou read about it in the text.

VEN power unit. The VEN power unitprovides hydraulic power to actuators for posi-tioning the VEN area.

Main fuel control (MFC). The MFC providesa regulated flow of fuel to the fuel nozzles.

Electrical control assembly (ECA). The ECAcomputes, schedules, and controls engine opera-tion.

Fan speed transmitters. The transmitters areeddy-current sensors mounted in line with thesecond stage fan blades. A permanent magnet,rotating at the RPM of the fan blades, inducesa voltage into a coil indicative of the fan speed.

Afterburner control (ABC). The ABCschedules fuel to the afterburner pilot and mainspray bars.

Afterburner (AB) flame sensor. This sensorprovides an electrical signal to the ECA. Thissignal must coincide with the AB no-light/lightcondition to start the afterburner.

Thermocouple harness. This device senses theexhaust gas temperature (EGT).

VEN position transmitter. The transmitterprovides feedback to the ECA to ensure the VENis in the correct position. It also gives feedbackto the engine monitor indicator (EMI) to indicatepercent of nozzle position.

VEN actuators. These actuators hydraulicallyoperate to position the VEN.

OPERATING PRINCIPLES

The VEN schedule is in response to movementof the throttle. Throttle setting repositions thepower lever angle (PLA) cam in the MFC,providing a linear variable differential transformer(LVDT) signal to the ECA. The ECA biases theVEN area schedule according to inputs from thefollowing sources:

Fan inlet temperature transmitter

Air data computer (ADC) ambient pres-sure

Fan/low-pressure turbine speed

Compressor/high-pressure turbine speed

Main fuel control metering valve

Afterburner

control metering valve

flame sensor

permission signal

Thermocouples

VEN area position transmitter

AB permission switch

The VEN area closes when the engine shutsdown and the throttle is in the off position. Asthe throttle is moved to the idle position, the VENarea rapidly moves to an almost full openposition. This aids engine starting and lowersengine thrust, allowing higher idle speeds andreducing engine acceleration time. As the throttleadvances past the idle position, the VEN areaschedule closes the VEN area, increasing thrust.As the VEN area decreases, EGT increases.

The ECA adjusts the VEN area for varyingatmospheric conditions. When the ECA receivesan ambient pressure signal from the air datacomputer, it means the aircraft is at a pressurealtitude of 9,000 feet or greater. The ECA nowincreases low-pressure turbine dischargetemperature and fan/low-pressure turbine speedlimits, providing more thrust.

Below 4,000 feet pressure altitude, theschedules trim back to reduce fuel consumption

Figure 5-17.

and to increase hot section life. This trim signalvaries in size between 4,000 and 9,000 feet. Toensure correct positioning of the VEN area,the VEN position transmitter LVDT providesfeedback to the ECA. Any error between theactual VEN area position and its required positiongoes to the bias signal to readjust the VEN areato its correct position.

As the throttle advances into military (MIL)power (100 percent) and afterburner (AB) ranges,the VEN area maintains the low-pressure turbinedischarge temperature within established limits.Throttle position establishes this limit. This limitis adjustable for ambient pressure for fan inlettemperature and for actual low-pressure turbinedischarge temperature values from the thermo-couple harness.

As the throttle enters AB range, the VEN areareopens slightly above the normal throttle setting,and low-pressure turbine discharge temperatureresets to a lower value. These conditions are helduntil the flame sensor signals the ECA of an ABlight-off. The ECA then releases its hold on theVEN area and re-establishes actual low-pressureturbine discharge temperature values. The VENarea will adjust to maintain actual low-pressureturbine discharge temperature limits.

The VEN power unit supplies hydraulic powerfor positioning the VEN. The power unit activateson an electrical signal from the ECA to the powerunit torque motor. The torque motor drives aservo, which supplies high oil pressure to thesynchronized actuators to open or close the VEN.

When the engine is shut down and thethrottle is in the off position, the variableexhaust nozzle is in what position?

What unit in the variable exhaustadjusts the VEN for atmospherictions?

nozzlecondi-

PROPELLER SYNCHROPHASINGSYSTEM

Learning Objective: Recognize operatingprinciples and features of propellersynchrophasing systems.

The propeller synchrophaser system discussedhere is common to the P-3 and C-130 aircraft andsimilar to the E-2 aircraft. In this section, you willlearn about the electrical operation of controllingand synchrophasing the hydromatic propellers ofmultiengined aircraft.

PROPELLER GOVERNOR

A propeller governor is a control device thatcontrols engine speed by varying the pitch of thepropeller. Increasing the propeller pitch adds loadon the engine and increases propeller thrust. Thisload reduces engine speed. Conversely, decreasingthe propeller pitch reduces engine load, whichincreases engine speed. Therefore, engine speedis a function of propeller pitch. Furthermore, ifthe propeller governor setting remains unchanged,any variation of power produced by the enginetranslates into a corresponding variation ofpropeller thrust. The system does this by varyingthe propeller pitch while engine speed remainsconstant. The best engine efficiency is when enginespeed is constant; therefore, because it controlsengine speed, the propeller governor achievesengine efficiency.

The pitch of hydromatic propeller bladeschanges by porting hydraulic fluid onto thepropeller piston in the propeller dome. The actionon the piston transmits through a geared cammechanism that rotates the propeller blades to thepitch desired.

The governor is the constant speed controldevice used with the hydromatic propeller.The output oil of the pump goes to eitherthe inboard or the outboard side of the propellerpiston.

There are two separate ranges of propelleroperation—the flight range and the groundoperating range. The flight range includes thetakeoff roll after the power levers advanceforward for takeoff. For the ground operatingrange, power levers return aft of the flight-idledetent. In the ground operating range (taxi range),power lever position determines propeller bladeangle. A hydromechanical system, with linkageto the power lever, meters oil pressure to eitherthe increase or decrease side of the propellerdome. As the power lever moves forward towardFLIGHT IDLE, a simultaneous increase in blade

angle and fuel flow occurs, providing increasedpower, As the power lever moves aft fromFLIGHT IDLE, blade angle decreases and fuelflow decreases, reducing power. Fuel flow beginsto increase when the blade angle decreases to thepoint that the propeller is delivering negativethrust. Reverse power continues to increase untilthe power levers reach the full aft position. Duringoperation in the ground operating range, there isno electronic governing.

In the flight range of operation, the powerlever is forward of the flight-idle detent. Inthis range, a flyweight governor, driven bypropeller rotation, mechanically controls propellerspeed. In normal operation, the pitch-changeoil goes through the feather valve to eitherthe increase or- decreasepropeller.

SYNCHROPHASING

pitch portion of the

The synchrophaser has different functions,depending upon the mode of governing selectedby the flight crew. The synchrophaser does notfunction in the mechanical governing mode, butthe mechanical governor controls the blade pitchand so propeller RPM. In a normal governingmode, the synchrophaser helps the mechanicalgovernor by limiting engine transient speedchanges, or to changes in flight conditionsaffecting propeller speed.

In a synchrophasing governing mode, thesynchrophaser helps the mechanical governor bymaintaining all propellers at the same RPM. Itdoes this by maintaining a preset phase relation-ship between the master propeller number 1 bladeand the number 1 blades of the slave propellers.This serves to reduce noise and vibration in theaircraft. In the synchrophasing governing mode,the synchrophaser also provides the limiting oftransient speed changes as it does in normalgoverning mode.

The synchrophaser consists of four maincomponents—a pulse generator, a phase and trimcontrol, a speed-bias servo assembly, and thesynchrophaser.

Pulse generator. The pulse generator providesthe information needed by the synchrophasersystem to produce speed and phase control of theaircraft propellers.

Each propeller has a pulse generator thatconsists of a permanent magnet on the propellerspinner and a stationary coil installation in thegovernor control. Each time the permanentmagnet passes by the coil it generates a pulse. Inother words, each revolution of the propellergenerates one pulse.

Phase and trim control. The phase and trimcontrol functions as a means of setting phaserelationships between mast er and slave propellers.It also trims the master engine.

The phase and trim control consists of sevenpotentiometers that receive a fixed dc voltage fromthe synchrophaser. The wiper of the master trimpotentiometer supplies a voltage through themaster select switch to the synchrophaser to trimmaster engine speed. The other six wipers connectto relay contacts that separate the wipers into twogroups of three per group. One group correspondsto engines 1, 3, and 4 when engine 2 is master.The other group corresponds to engines 1,2, and4 when engine 3 is master. These wipers supplybias voltages to the phase correction circuits ofthe synchrophaser to set propeller phase anglesof other than 0 degree.

Speed-bias servo assembly. The speed-biasservo assembly functions as a means of translatingsynchrophaser electrical signals into a mechanicalbias on the mechanical governor speeder spring.

The synchrophaser supplies the servomotorwith a reference voltage that is 90 degrees out ofphase with the aircraft 400-Hz source. Thesynchrophaser also supplies a control voltage thatis either in phase or 180 degrees out of phasewith the aircraft 400-Hz source. Therefore, thein-phase control voltage lags the reference voltageby 90 degrees. This lag results in counterclockwisemotor rotation when viewed from the output gearof the electric brake. The 180-degree out-of-phasecontrol voltage leads the reference voltage andcauses clockwise rotation. The amplitude of thecontrol voltage determines motor speed andtorque output.

The motor drives a reduction gear train,which, in turn, drives a potentiometer wiper andthe electric brake. The potentiometer receives afixed dc supply from the synchrophaser across itsresistive element. When the motor rotates, thewiper transmits a corresponding feedback voltage

to signal winding number 2 of the magneticmodulator. The electric brake has clutch-controlled input and output shafts. The outputshaft drives a lever, which biases the speederspring in the propeller governor. Energizingthe clutch decouples the two shafts, lockingthe output shaft and leaving the input shaft freeto turn.

The synchrophaser. The synchrophaser hasfour channels, which correspond to the aircraft’sfour engines. Figure 5-18 is a schematic of thesynchrophaser. For explanation purposes, onlytwo channels (one slave channel and one masterchannel) are shown. Each channel has a push-pullpower amplifier feeding the control winding ofits corresponding servomotor in the speed-biasservo assembly. Magnetic modulators using dccontrol current furnish a phase and amplitudecontrolled ac signal to the push-pull amplifierinput. The synchrophaser changes all signal inputsto dc voltages proportional to the error beforethey are applied to the modulator. The modu-lators are the signal summing devices for the twooperational modes of the synchrophaser.

The magnetic modulators function on a coresaturation basis. Each modulator consists of a dcbias winding, a 400-Hz excitation winding, andtwo control windings (signal winding number 1and signal winding number 2). With no signalsapplied elsewhere, the 400-Hz excitation vohageappears as a 400-Hz output of negligible ampli-tude due to the bias winding current. Any currentin either or both signal windings will change theoutput.

The size of the signal windings current controlsthe amplitude of the output; the current directioncontrols the phase of the output. Thus, currentfrom pin 10 to 9 in winding number 2 and currentfrom pin 8 to 7 in winding number 1 of anymodulator produces a voltage 180 degrees out ofphase from the excitation voltage. Current in theopposite direction in the signal windings producesan in-phase voltage. Simultaneous currentsflowing in opposite directions in the two signalwindings produce a signal that is the algebraic sumof the two signals. Then, the modulator producesa 400-Hz signal, which is either in phase or 180degrees out of phase with the excitation voltage.This signal is amplified and fed to the servomotorcontrol winding. The 400-Hz voltage in thereference winding of the servomotor goes through

a series capacitor, giving the voltage a 90-degreephase shift from the aircraft power source.Appropriate signals to the modulators causeclockwise or counterclockwise rotation of themotor because of phase difference in the speedbias motor windings. The use of the twosignal windings in the modulators, along withappropriate relay switching, permits the twomodes of synchrophaser operation.

OPERATIONAL MODES

The normal governing mode provides im-proved engine response to transient RPM changes.In this mode, the synchrophaser receives signalsfrom the power lever anticipation potentiometersand the engine tachometer generators. Signalsfrom these result in a temporary resetting of themechanical propeller governor. This resettingadjusts for power lever changes and engine speedchanges, thus limiting engine overspeeds orunder speeds.

The synchrophasing governing mode synchro-nizes engine speeds, regulates propeller phaseangles, and maintains the limiting features of thenormal governing mode. In the synchrophasinggoverning mode, one engine (2 or 3) is the master.The master engine operates in normal governingmode, while the other three engines (slaves) followchanges in speed or phase of the master withinpreset limits.

Normal Governing Mode

In normal governing mode, the propellergovernor switch is in the NORMAL positionand the power lever switch closes, providingreference voltages to the servomotors. Thesynchrophase master switch is OFF and the PROPRESYNCH switch is in NORMAL. All relays arede-energized, resulting in the speed and phaseerror circuits being grounded. Each phase- andspeed-error signal side of every magnetic modu-lator signal winding number 2 (pin 10) ends ona dummy load (fig. 5-18) within the synchro-phaser. The other side (pin 9) connects to thefeedback circuit in the speed-bias servo assembly.The controlling signals go to signal windingnumber 1 (pin 8) of each modulator. All channelsfunction identically while in the normal governingmode.

Figure 5-18.-Synchrophaser control schematic diagram.

THROTTLE LEVER ANTICIPATION. —Any power lever movement (fig. 5-19) causesa change in dc voltage at the anticipationpotentiometer wiper, which serves as a voltagedivider for the RC circuit. The charging voltagefor the capacitor is directly proportional to theposition of the power lever. The change in chargeon the capacitor is directly proportional to the rateat which the power lever moves. If the power levermovement is to decrease engine power, the

capacitor charges up to a more positive voltagevalue. This results in a current from pin 7 to pin8 in signal winding number 1 of the magneticmodulator. A lagging voltage surge appears in theservomotor control winding, causing counter-clockwise rotation. This rotation resets themechanical governor towards decrease pitch tocompensate for the reduced power setting. As theservomotor rotates, the feedback potentiometerbegins canceling the error signal by causing a

Figure 5-19.-Power lever anticipation and speed derivative circuits in normal governing mode.

current in signal winding number 2. The magneticfield of this current is in opposition to themagnetic field of the signal current in windingnumber 1. This stops the servomotor. As theanticipator capacitor continues to charge to itsnew peak value, the current in signal windingnumber 1 decays to zero. The feedback potenti-ometer is still applying voltage to signal windingnumber 2. This results in a leading voltage to theservomotor control winding that returns themotor to its original position. This positioncorresponds to a zero-volt feedback potentiometerposition. In retarding the throttle lever veryrapidly, the peak voltage will overcome the reversebias on diode CR620. This will limit the signalvalue to prevent overcompensation toward a flatblade pitch. For an increase in engine power, thecapacitor discharges. This causes a current in theopposite direction in signal winding number 1,which results in a temporary resetting towardincreased pitch. The amount of reset in either casedepends on the rate at which the lever moves.Mechanical stops in the speed-bias servo assemblylimit speed resets to plus 10 and minus 10 percentregardless of the applied signal. Furthermore, stopsin the propeller control valve housing linkagereduce the limits to plus 6 and minus 4 percent.

LIMITING ENGINE TRANSIENT SPEEDCHANGES. —The speed-derivative circuit in thesynchrophaser (fig. 5-19) senses changes in engineRPM and produces output signals, which dampenthe engine RPM changes. The speed-derivativecircuit does this by translating the frequencychanges received from one phase of the tach-ometer generator into signal voltages. Themagnitudes of the signal voltages vary at the ratethe tachometer generator frequency changes. Thesignal voltage goes to signal winding number 1of the magnetic modulator, where it is summedand sent to the push-pull amplifier. Afteramplification, the signal goes to the servomotorcontrol winding. The servomotor adjusts thespeeder spring tension, which begins a change inpropeller pitch, thus dampening the change inengine RPM.

The action of the speed-derivative circuit isfurther described as follows: The voltageproduced on the collector of transistor Q603 isproportional to the output frequency of thetachometer generator. When engine RPM isconstant, the voltage on the collector is constantand capacitor C621 charges through resistor R634and signal winding number 1 of the magneticmodulator. Current in signal winding number 1

decays to zero as the charge on capacitor C621reaches the potential on the collector of transistorQ603. When engine RPM changes, a changein the collector voltage of transistor Q603proportional to the change in tachometer gene-rator frequency occurs. This causes capacitorC621 to change its charge at the rate in which thetachometer generator frequency is changing. Thisproduces a current in signal winding number 1.The current size varies at the rate at which theengine is varying offspeed. Its direction is suchthat the amplified signal in the servo-bias assemblycontrol winding drives the servomotor in adirection to dampen the drift in engine RPM.Speed-error signals in signal winding number 1from the speed-derivative circuit cancel in thesame manner as anticipation signals from theanticipation circuit cancel.

The speed-derivative and power leveranticipation circuits are much more sensitiveto engine RPM changes than the mechanicalgovernor flyweight speeder spring. The governingaction of the flyweight and speeder springimproves the mechanical governor’s response tochanges in power lever settings and engine RPM.

Synchrophaser Mode

In adding synchrophasing to normal governingmode, the master switch selects either engine 2 or3 as the master engine (fig. 5-18). With masterengine selection, relays energize, removing dummyloads from signal winding number 2 of allmagnetic modulators, except the master channelmodulator. Also, the outputs of the speed-errorand phase-error circuits of each synchrophaserchannel (except the master) are taken from groundand connected to signal winding number 2 of theirrespective modulators. While the slave engines arein synchrophasing mode, the master engineremains in normal governing mode. Essentially,pulses from the master engine form into sawtoothwaves and are compared to pulses from each ofthe slave engines in the slave channel samplingcircuits. If slave pulses are not in phase with themaster pulse (sawtooth), error detection occursin the respective synchrophaser channel. Theerrors then go to signal winding number 2 of thechannel magnetic modulator. Here, they aresummed with any error signals that exist in signalwinding number 1. The result of the error signalsis amplified and fed to the control winding of therespective speed-bias servomotor. This alters thetension of the slave governor speeder spring,correcting for engine speed differences and

Figure 5-20.

propeller blade angle errors. (Figure 5-20 shows onechannel of the synchrophasing circuit.)

The following paragraphs describe phase andspeed error sensing.

SAWTOOTH FORMER.—The pulse from themaster pulse generator is transformer-coupled to thesawtooth former. The slave pulse generators aretransformer-coupled to the channel sampler circuits.Figure 5-21, views A and B, shows the master pulseand the resultant sawtooth formed in the sawtoothformer.

SAMPLING CIRCUITS.—Pulses from the threeslave engines couple to the grids of the samplingcircuit tubes, while the sawtooth voltage goes to theplate of one tube and the cathode of the other tube inall sampling circuits. The positive-going portion of theslave pulse places the tubes in a conductive state.(Sampling is the same in all channels.) Refer to figure5-20. If the sawtooth is at zero potential when theslave pulse occurs, neither tube conducts, so the phasedifference and speed

error circuits receive no signal. With the sawtooth inthe positive region, tube V205 conducts. Thecorresponding voltage changes go to the grid of phasedifference tube V206A.

You can see the nature of the sampling action byreferring to figure 5-21. The time interval betweenmaster pulses is the time interval for 360 degrees ofpropeller rotation (fig. 5-21, view A). The time intervalof the sawtooth, which the master pulse generates, is360 degrees. The half-time interval, or 180-degreeposition, is the sawtooth zero potential point (fig. 5-21,view C). When the slave pulse occurs at the zero point,the propellers are on phase with a 180-degree phasedifference between them. All references to slavepulses are with respect to the 360-degree interval. Thepoint of occurrence of the slave pulse determines thesignal size in the sampler circuit (fig. 5-21, view D).This signal represents the phase difference betweenpropellers.

NOTE: Since the propellers are four-bladed, therelative blade position between the master and slavepropellers is exactly the same when the slave

Figure 5-21.—Master and slave pulse comparisons.

propellers differ from the master by one-halfrevolution. Therefore, consider a 180-degreephase difference as an on-phase condition in pulsecomparisons <

PHASE DIFFERENCE CIRCUITS. —Lookat figure 5-20. The phase difference circuits receivethe voltages the sampling circuits generate at thegrids of the respective tubes. With a 180-degreephase difference signal at the grid of V206A, thevoltage at the wiper of potentiometer RP251adjusts to null or zero volts. This voltage changesproportionally with the grid signal from thesampling circuit, and hence represents the phasedifference between propellers. The followingparagraphs discuss the effect of this voltage onsynchrophaser output.

OFF-SPEED CIRCUITS. —When the pro-peller goes off-speed, the sampling circuit sensesthe condition as a sharp voltage change. (Whenthe slave propeller is on-speed, the slave pulsesoccur at the same time interval in each successivesawtooth cycle.) When an underspeed or over-speed condition develops, the slave pulse occursat a different position for each sawtooth cycle.At one point, the slave pulse falls up or down the

sawtooth (fig. 5-22). Think of this as a phase errorthat occurs too rapidly for the phase-error circuitto compensate. During an underspeed condition,the voltage going to the phase difference tubesuddenly changes from positive to negative (fig.5-22). For an overspeed condition, the voltagesuddenly changes from a negative to a positive.

As you read this paragraph, refer to figure5-20. The action of the off-speed circuit isas follows. The underspeed voltage change,registered in the plate circuit of the phasedifference tube, couples to the off-speed circuitthrough capacitors C253 and C254. The voltagechange in the plate circuit is positive and thereverse bias on diode CR251 is reinforced.However, the reverse bias on diode CR252 ismomentarily overcome, allowing capacitor C255to charge positively. Capacitor C255 dischargesslowly through its parallel resistive network andmaintains a grid bias on the off-speed tube V206B.(For an overspeed, diode CR251 conducts andcharges capacitor C255 negatively.) Successivesharp voltage changes add to the charge oncapacitor C255 until correction of the off-speedcondition occurs. Like the phase differencecircuit, the off-speed circuit has a zero-voltadjustment potentiometer in the cathode circuit.

Figure 5-22.-Off-speed pulse comparison: (A) slave underspeed; (B) slave overspeed.

The pulsating voltage generated at the wiper bythe changing grid bias on the tube is thespeed-error voltage. The effect of this voltage isdiscussed later.

SIGNAL SUMMING. —As you can see bylooking at figures 5-20 and 5-23, the phase-difference and speed-error voltages couple to thesignal summing point. Also connected to thispoint are the limiting circuit, the phase-controlpotentiometer, and magnetic modulator signalwinding number 2. Any difference in potentialbetween the signal summing point and thefeedback potentiometer causes a current inthe modulator winding. This results in motormovement, which, in turn, causes the feedbackpotentiometer to null whatever potential is at thesumming point.

PHASE-ERROR CORRECTION. —Whenthe slave propellers aren’t on-phase, phase-difference voltage acts through the averagingcircuit to the summing point causing speed-bias

servomotor rotation to correct the off-phasecondition. This action occurs provided the phasecontrol potentiometer (fig. 5-23) is set for zerovolts. In this case, the phase difference voltageactually represents phase error. However, it isdesirable to have a slave propeller maintain aspecific angle of lead or lag to the masterpropeller. The phase control potentiometer thenadjusts to cancel the phase-difference voltage atthe summing point. For example, to maintain aslave lead of 10 degrees from the 180-degreeposition of the master pulse, the phase controlpotentiometer adjusts to cancel the phase-difference voltage. The phase-difference voltageis generated when the slave propeller leads themaster propeller by 10 degrees. Thus, whenthe slave propeller leads by 10 degrees, netpotential at the summing point is zero, andthe speed-bias servomotor does not move.When the slave propeller changes from the10-degree lead condition, the phase-differencevoltage changes. The net potential at the summingpoint is the difference between the phase control

Figure 5-23.-Simplified schematic of signal summing.

potentiometer setting and the phase-differencevoltage. This is how phase-error voltage and itssize depend on the slave propeller’s degree of leador lag from the 10-degree lead condition. Themanual phase control is adjustable to allow a slavelead or lag of 45 degrees.

SPEED-ERROR CORRECTION. —Duringthe phase-error correction, Zener diodes VR405and VR408 isolate the speed-error potentiometerfrom the signal summing point. When anoff-speed occurs, these diodes conduct andconnect the speed-error circuit to the summingpoint. For large speed errors, the speed-errorsignal is a pulsating dc of very low frequency. Atthe same time, the phase-error signal is a rapiddc step voltage consisting of sharp potentialchanges, which trigger the speed-error circuit. Thedc step voltage changes developed in the cathodecircuit of V206A (fig. 5-20) are averaged andeffectively blocked by resistor R268 and capacitorC251. The sharp potential changes passing resistorR268 and capacitor C251 leak to ground byresistor R272 and capacitor C256, leaving thespeed-error signal in control. When the enginereaches on-speed condition, but is still out ofphase, the speed-error signal drops off until apoint where the Zener diodes stops conductingand the phase-error signal assumes control.

TWO-PERCENT LIMITING. — T h epotential at the summing point has a limit of plusor minus 5 volts. The limiting circuit (R442, R443,R444, R445, CR409, and CR410 as shown in fig.5-23) accomplishes this by clipping any signaloutside of the range limit. In the off-speed circuit,this limited signal provides only enough outputto drive the speed-bias servomotor to correct atwo-percent speed change. This occurs because themotor movement results in a feedback voltage thatcancels the speed-error signal. The slave propellercannot follow large speed changes of the masterengine. This prevents a slave from following anoverspeeding or underspeeding master.

until it matches the feedback signal. When thisoccurs, the potential at the summing point is zeroand the motor stops moving, leaving a portionof the error uncorrected. This is insignificant forerrors that occur about a set point—leadingand lagging errors. However, for errors thatcontinually occur in one direction, the erroraccumulates, and can become large enough toreduce system efficiency. To overcome this, aprocedure provides a lock on the mechanicalgovernor bias, while the servomotor recenters ata zero-volt feedback position; this is resyn-chrophasing.

As you read this paragraph, refer to figure5-19. While in synchrophasing mode, place theprop resynchrophase switch in the RESYNCHposition. This energizes the electric clutch-brakeson the slave engine speed-bias servo assemblies.The brakes lock the output shafts and the clutchesdecouple the input and output shafts. This actionleaves the input shafts free to turn. Locking theoutput shaft retains whatever mechanical bias ispresent on the mechanical governor. At the sametime, relay K503 energizes. This opens the signalwinding number 1 circuits on the magneticmodulators. (The master channel remains innormal governing mode through a parallel groundprovided by the master relay to signal windingnumber 1 of the master channel modulator.) Aftera time delay sufficient to assure brake and clutchactuation before allowing the recentering action,relays de-energize, removing the phase- and speed-error circuits from the slave channel magneticmodulators and connecting the dummy loads.The only input to the slave channel magneticmodulators is now the feedback potentiometeracting into the dummy load through signalwinding number 2. The generated signal thendrives the motors until the feedback voltage iszero. Upon releasing the prop resynchrophaseswitch, all circuits return to the synchrophasingmode. Correction of the accumulated phase errorcan now occur.

TROUBLESHOOTING HINTSResynchrophasing

The need for resynchrophasing arises from thenature in which phase angle correction circuitsoperate. As phase errors occur, the servomotorrotates to correct the error. At the same time, thefeedback potentiometer moves and cancels aportion of the error signal. As the phase-errorcorrection occurs, the phase-error signal decreases

When trouble occurs in the propeller systemelectrical controls, the first troubleshooting stepyou perform is determining which part of thesystem is malfunctioning. You can determine thisby completing an operational check. The next stepis to refer to the electrical schematic and thetroubleshooting chart in the aircraft maintenanceinstructions manual, and then analyze the trouble.

Q1. What are the two ranges of propelleroperation?

Q2. What mode of operation is the synchro-phaser in when maintaining all propellersat the same RPM?

Q3. What unit provides a means of setting phaserelationships between the master and slavepropellers in the synchrophase system?

Q4. The synchrophaser speed derivative circuitis used to dampen changes in

Q5. The slave propellers are prevented fromfollowing an overspeed or underspeedmaster propeller by what synchrophasercircuit?

PROPELLER CONTROL SYSTEM

Learning Objective: Recognize operatingprinciples and features of aircraft propellercontrol systems.

The propeller control system is similar tosystems already discussed in this chapter. It is anintegral part of, or works with, the propellersynchrophaser system and the propeller governingsystem. The discussion that follows concernspropeller pitchlock, the negative torque system,and feathering, unfeathering, and autofeatheringoperations. You should refer to figure 5-24throughout this discussion of the components inand operation of the propeller control system.

PROPELLER

The four-bladed propeller converts engineshaft horsepower to thrust. The propeller consistsof two principal sections—the rotating section and

the nonrotating section. The rotating sectionconsists of the blades, hub, spinner, and the domethat houses the pitch changing mechanism. Thenonrotating section contains the pressure andscavenge oil pumps, the governor controlmechanism, and the spinner afterbody. It isa constant-speed, full-feathering, reversingpropeller, having the added features of pitchlockand a combination synchronizing and synchro-phasing system.

Low-Pitch Stop Assembly

A mechanical stop assembly in the propellerdome limits low pitch blade angle to maintain aminimum blade angle. When the power leversposition is below 28 degrees, a cam operatedbackup valve in the propeller housing collapsesthe low-pitch stop levers. The blades move towardthe reverse position as directed by the power leverbeta schedule.

Beta Follow-up System

The beta follow-up system provides a variablehydraulic low-pitch stop. At the FLIGHT IDLEpower lever position, the beta follow-up stop is10-degree blade angle; the mechanical low-pitchstop is 13 degrees. The purpose of the betafollow-up stop is to provide a secondary low-pitchstop. As the power lever moves toward theTAKEOFF position and the blade angle increases,the mechanical low-pitch stop remains at 13degrees. The beta follow-up control programs thehydraulic stop in relation to power lever positionto a maximum of 22.5 degrees of blade angle.If a sudden engine failure occurs and thenegative-torque system fails to operate, betafollow-up prevents excessive reduction in bladeangle and associated violent yawing.

Negative Torque System

The negative torque system (NTS) protectsthe aircraft from excessive drag by limitingthe negative torque from the propeller to apredetermined value range of -150 to -500 shafthorsepower. During a negative torque condition,NTS provides a mechanical signal that overridesthe propeller governing action to increase bladeangle. When the propeller reaches a positionwhere it is no longer developing negative torque,the propeller governor regains control of thepropeller. If the negative torque condition persists,a cycling action continues from the mechanical

Figure 5-24.-Propeller control circuit.

signal to the propeller governor, and vice-versa,until some corrective action is taken.

ANTS INOP warning light illuminates whenthe 45-degree air start blade angle circuit energizes.The air start blade angle system limits negativehorsepower during in-flight restart operations ifNTS fails. The blades, upon reaching 45 degrees,energize the air start blade limit circuit, whichdrives the feather valve to the feather position.The blades moving toward the feather positionopen the circuit, de-energizing the feather solenoidvalve. The result of this action is a cycling of theblades around the air start blade angle switch. Forground unfeather operation, bypass the air startblade limit switch by actuating the feather pumppressure cutout override switch while unfeatheringthe propeller.

Pitchlock

A hydraulic, speed-sensitive pitchlockmechanism is in the propeller assembly to preventoverspeeding. It is completely automatic andprevents further decrease in blade angle whenengaged. Automatic engagement occurs in twoways: by loss or control oil pressure and whenRPM exceeds 103.5 percent. The pitchlockgovernor, sensing the overspeed, allows thepitchlock teeth to engage, which prevents furtherdecrease in blade angle. Due to teeth design, theblade angle can increase if the normal governingcontrol is restored.

BLOCKOUT RANGES. —There are twoblade angle ranges where pitchlock does notoperate. One range is from plus 17 degrees tominus 14 degrees. This allows for RPM surges asblade angles reduce during approach and landing.The second is blade angles between 57 degrees and86 degrees (full feather). This is necessary todecrease blade angle for air starting. Blade angleswill be less than 57 degrees at 405 knots or limitMach speed.

PITCHLOCK RESET. —A reset system willreset pitchlock up to 109 percent RPM at bladeangles above 10 degrees with the power leverbelow 28-degree coordinator position. This109 percent RPM setting of the pitchlock ismomentary. It is necessary when the power levermoves very rapidly into the beta range and bladeangle reduction is not quick enough. Pitchlockreset comes into use during an aborted takeoffor landing. In this condition the increase in fuelscheduled will cause engine speed to increaseabove 103 percent RPM.

FUEL GOVERNOR AND PROPELLERPITCHLOCK TEST SWITCH. —Four two-position (TEST-NORMAL) switches, one foreach propeller, are on the engine check panel.Upon placing a switch to TEST, the propellerspeed-bias servomotor receives a continuousoverriding signal. The signal drives the mechanismfull travel toward decrease pitch. This resets thepropeller governor to a speed of 106 percentRPM, permitting a ground check of the pitchlockand fuel governor functions.

Propeller Feathering

Feathering aligns the propeller to the airstreamto minimize drag during engine shutdown con-ditions. Feathering is started by pulling the engineemergency shutdown handle, autofeathering, ordepressing the feather switch button. The feathervalve ports hydraulic fluid to increase pitch onthe propeller blades. This action bypasses all othercontrol functions.

NORMAL FEATHERING. —To accomplishnormal feathering, pull the engine emergencyshutdown handle. Pulling the handle mechanicallypositions the feather valve and electricallyenergizes the feather button solenoid. Also,current goes to the auxiliary pump and to thefeather valve solenoid, which hydraulicallypositions the feather valve to feather the propeller.When the propeller fully feathers, oil pressurebuildup operates a pressure cutout switch, causingthe auxiliary pump and feathering solenoid tode-energize. The feathering button light then goesout.

Feather Pump Pressure Cutout OverrideSwitch. —There is a push-button switch next toeach feather button. Actuating the override switchbypasses the feather pump pressure cutout switch.This permits continued operation of the featherpump if the feather pump operation finishesbefore the propeller reach the full feather.

Feathering Switches (Buttons). —Four guardedfeather switches (buttons), one for each engine,provide an alternate method for feathering thepropellers. Pressing a button to FEATHER cutsoff fuel electrically, and energizes the feathersolenoid and the feather pump motor to featherthe propeller. The propeller unfeathers when thebutton goes to the unfeather position. The centerposition is for normal propeller operation. A lightin the feather button illuminates when the circuitto the feather pump is energized.

AUTOMATIC FEATHERING. —The auto-matic feathering system performs its function byenergizing the feathering button holding coil(pulling in the feathering button). This occurswhen the autofeather arming switch is in theARMED position and the engine loses power,which results in a large decrease in propellerthrust. This system is for use during takeoff only,and it functions above 60°±2° coordinatorsetting. Only one propeller will feather auto-matically.

Autofeather Arming Switch. —There is anautofeather ARMED-OFF switch on the auto-feather and RPM control panel. It provides thecontrol for the autofeather system. During normaloperations and when the switch is in the ARMEDposition, all power plants have autofeatherprotection. With the switch in the ARMEDposition and electrical power to the power leverquadrant, autofeather arming switches close andthe autofeather armed lights illuminate. Both thearming switch and power lever quadrant switch(activated at 60 degrees) must close before thethrust-sensitive signal device can cause propellerautofeathering.

Autofeather System Indicators. —Four greenindicator lamps, one for each propeller, illuminateto show the arming of each individual propellerautofeathering circuit. The lights are on the pilot’soverhead control panel. Also, should a propellerautofeather, its light will remain on and the otherswill go out.

UNFEATHERING ON THE GROUND. —The propeller is unfeathered by holding thefeathering button in the unfeather position. Theair start switch (in the propeller) limits the bladeangle decrease to 45 degrees. For further bladeangle decrease, you hold the pressure cutoutoverride button in.

PROPELLER CONTROL OPERATION

In this section, you will read about a typicalpropeller control operation. Before the operationis presented, you are given the following con-ditions:

The engines are running.

The engines are operating in the taxi range(power levers between 0 degree and 34degrees on the coordinator).

Each power lever controls the propeller bladepitch and engine fuel flow through hydro-mechanical linkages. If a power lever is below 28degrees and the blade pitch angle is above 10degrees, the pitchlock reset valve solenoid(fig. 5-24) activates, resetting pitchlock up to 109percent RPM.

As the power levers advance above 34 degreesinto the flight-idle range, the flyweight propellergovernor assumes propeller pitch control. Thisgovernor increases or decreases blade pitch tomaintain 100 percent RPM by directing fluidthrough the feather valve. The temperaturedatum system controls proper engine temperatureoperation.

When ready for takeoff, the pilot placesthe AUTOFEATHER switch in the ARMEDposition. As the power levers advance beyond 60degrees, the number 3 power lever switches closeto the HIGH POWER position. If a propeller’sthrust drops below 500 pounds, the thrust-sensitive switch closes. This energizes theautofeather relay. Power goes through theautofeather relay contacts to energize the featherswitch pull-in and hold-in solenoid. Theautofeather relay loses its power from the featherswitch, but the feather switch remains energizedthrough a holding circuit. Power runs from theenergized feather switch through the contacts ofthe feather cutout relay to the feather pump powerrelay, to the feather pump motor. Power also goesto the feather valve solenoid through the feathercutout relay. The feather valve solenoid moves thefeather valve, which ports oil to increase pitch,causing the propeller to feather. The fail-safe-typefuel control relay activates by power from thefeather switch, causing the engine fuel shutoffvalve to close.

As the propeller feathers, the beta cam switchcloses when the blade angle goes past 74 degrees.The feather motor cutout switch closes when thepropeller reaches full feather and the oil pressureincreases. This completes a circuit that causes thefeather cutout relay to energize, de-energizing thefeather valve solenoid. Besides cutting out thefeather pump power relay, the energized feathercutout relay maintains its own holding circuit.This completes the autofeathering cycle.

When shutting down the engine with theEMERGENCY ENGINE SHUTDOWN handle,power from the emergency engine shutdown relaysis directed to the feather switch pull-in and hold-insolenoid. Each emergency shutdown handle hastwo switches and two relays to make sure theelectrical systems energize. The engine is shutdown

and feathered as described above. Also, pressingin the feather switch causes the feather cycleoperation to take place.

When the flight crew restarts an engineduring flight, pull the feather switch out to theUNFEATHER position and hold it there. Thefeather cutout relay is de-energized, and powergoes to the feather pump power relay. The featherpump motor pressure builds up and oil flows tothe decrease pitch side of the propeller dome,causing the propeller to start unfeathering. Duringrestarts the feather valve solenoid energizes duringNTS inoperative conditions only.

When blade angle decreases to less than 45degrees, the beta cam air start switch closes,completing the path for current through the airstart control relay. The air start control relayenergizes the feather valve solenoid. This causesthe propeller blades to increase toward the featherposition. The beta cam air start switch opens at48-degree blade angle, causing the air start controlrelay and feather solenoid valve to de-energize.This causes cycling of blade angle around the airstart beta cam switch (45 degrees) and causes theNTS INOP warning light to flash. The light blinksbecause of the blade angle cycling action. At thisindication, the flight crew pulls the emergencyshutdown handle to abort the restart.

The propeller is used to convert engine shafthorsepower to

What mechanical unit maintains aminimum blade angle?

What does the beta follow-up systemprovide?

What system resets pitchlock RPM from103.5 percent to 109 percent during anaborted takeoff?

What valve bypasses all other controlfunctions to feather the propeller?

Q6. What propeller system is used only duringtakeoff?

Q7. During ground propeller unfeathering, howdo you decrease blade angle below 45degrees?

APPROACH POWERCOMPENSATOR SYSTEM

Learning Objective: Recognize operatingparameters and features of aircraftapproach power compensator systems.

The approach power compensator (APC)system (fig. 5-25) automatically controls enginepower to maintain the best angle of attack duringlanding approaches. This permits the pilot todirect most of his attention toward flying theapproach. During an APC approach in light tomoderate turbulence, the set can maintainairspeed within a range of ±4 knots.

The major components of the system are acontrol panel, accelerometer, computer, controlamplifier, a potentiometer to detect changes inelevator position, and a rotary actuator, whichmoves the throttle linkage. The aircraft angle-of-attack transducer supplies angle-of-attack signalsto the computer. A compression switch on thelanding gear breaks the APC circuit upontouchdown.

With the landing gear down and weight offthe gear, the APC is energized by placing theengage switch to ON. The switch is magneticallyheld on until the aircraft touches down, thethrottle overrides the system, or the pilot placesthe engage switch to OFF. The APC will alsorelease automatically if either the override switchor the compression switch fails.

With the set engaged, the accelerometer,angle-of-attack transducer, and the ambienttemperature selector switch supply informationto the computer. The APC computer also receiveselevator position signals from a potentiometermounted in the control valve linkage of theelevator actuator. When normal acceleration is1 g and the angle of attack is optimum for alanding approach, the computer sends no cor-rective signal and the throttle position does notchange. The computer interprets deviations fromthese values as either offsetting each other or as

Figure 5-25.-Approach power compensator simplified schematic.

requiring a change of power setting. When a the APC operates with an approximate 11 percentpower change is necessary, the computer sends an increase in gains from STD. This results in APCelectrical signal to the rotary actuator through the performance that is essentially the same regardlesscontrol amplifier. The rotary actuator motor then of ambient air temperature, with the followingdrives the engine control linkage to accomplish temperature settings: COLD below 4°C (40°F),the required power change. The system is capable STD from 4°C to 27°C (80°F), and HOT aboveof driving the throttle linkage at speeds up to 25 27°C.degrees per second. The APC can vary the engineRPM from full throttle to 67 percent RPM attemperatures above 0°F, or to 58 percent RPMat temperatures below 0°F.

Since engine performance varies with ambient Q1.air temperature, the APC includes a three-positiontemperature switch to compensate for this effect.At low ambient air temperatures, the thrustchange per degree of throttle movement is greaterthan at high ambient air temperatures. When theAPC temperature switch is in COLD, the APCoperates with a 15 percent reduction in gains Q2.from those with the switch in standard (STD).Conversely, with the temperature switch in HOT,

To maintain the best angle-of-attackduring landing approaches, the approachpower compensator automatically controls

The approach power compensator circuitde-energizes when the pilot turns it off, theaircraft touches down, and

Q3. What does the approach power compen-sator use to compensate for differences inambient temperatures?

BLADEFOLD SYSTEM

Learning Objective: Recognize operatingprinciples and characteristics of helicopterbladefold systems.

Bladefold on the H-60 helicopter is anelectromechanical system. The bladefold indexdrive unit positions the main rotor head beforefolding. There are four bladefold actuators—onefor each blade. A tail rotor blade positioneractuator positions the tail rotor for pylon andstabilator folding. The blade deice distributorassembly on the main rotor contains sequencingelectronics and power switching circuits forbladefold and pitchlock actuators.

Power goes to the main rotor distributorthrough the main rotor slip ring. The bladefoldcontrol panel contains system operating controlsand status/cue light capsules.

You can manually fold the pylon by using afolding pole. A lockpin switch on the pylonindicates when the pylon is locked in flightposition. The tail cone-pylon fold point has acrack open switch. This switch indicates pylondisconnection from the tail cone. Two stabilatorlockpin switches indicate when stabilator lockpinsare locking the stabilator panels in flight position.

In addition, switches on each rotor hubindicate fold or spread status, and pitchlockengage or disengaged status. The switch signalsgo to the bladefold control panel and distributor.

MASTER POWER DISTRIBUTION

DC power for the system is from the No. 2dc primary bus. This power goes through theweight-on-wheels and low oil pressure relays.Bladefold cannot operate unless the helicopter ison the ground and the rotor head not moving.

The transmission pressure switch energizes thetransmission low oil pressure relay. With bothweight-on-wheels and rotor head stopped, the28-Vdc master power goes to the blade deicejunction box.

The 28 Vdc at the blade deice junction boxgoes to two hydraulic system interlock relays.These relays actuate by pressure switches in theNo. 1 and No. 2 hydraulic systems.

To protect the flight controls, pressure can goto only one stage of the main rotor servos duringfold operation. One hydraulic system must be shutoff by the SERVO switch on either collective stickgrip. Pressure switches operate relays to ensurethat only one hydraulic system is on. Relayoperation permits master power to the bladefoldcontrol panel only when one hydraulic system ison.

The 28-Vdc rotor brake pressure switch sensesrotor brake condition. The pressure switchactuates a relay, which applies a signal to thebladefold control panel. This signal is for thebladefold control panel indexing logic. TheROTOR BREAK-APPLY-RELEASE indicatorson the bladefold control panel advise pilots whento apply or release the rotor brake.

FOLD SEQUENCE

The first step in folding the main rotor bladesis to index the rotor head to a known position.The bladefold index drive unit accomplishes thistask. A de-operated linear actuator in the unitengages the indexing motor output gear to therotor brake disc gear. An index command fromthe bladefold index drive unit drives rotor headto indexed position. A 2.5-degree maximumcutaway segment of an index slip ring in the slipring assembly senses this position.

With the BLADEFOLD switch in the FOLDposition, and the rotor brake disengaged, the rotorcan dr ive in e i ther the c lockwise orcounterclockwise direction. The rotor will drivein the direction of the shortest distance to theindex position. To do this, index drive slip ringconstruction is in two segments separated at theindex position, 180 degrees from index position,and by providing a reversing relay in the indexdrive unit.

The index motor rotates at 0.5 RPM. As therotor head reaches the 1.0-degree index position,index drive motor power switches off and thedrive motor brake de-energizes, applying thebrake. Power removal and brake applicationprevent the rotor head from overshooting theindex position.

When rotor head position is on the insulatedarea between index drive slip ring segments 180degrees from the index position, head rotation willbe counterclockwise toward the index position.

The index unit can also act as a backup systemto the rotor brake. You can use the index unit toengage the brake disk and to hold the rotor headstationary when performing maintenance on the

rotor brake. To do this, use the GUST LOCK-LKD-UNLKD switch on the miscellaneous switchpanel. The switch enables the index unit to extendand engage the rotor brake disk, without applyingpower to the index motor. A brake and gear traininside the motor prevents motor output shaftrotation with no power to the motor. The indexunit effectively acts as a lock to hold the rotorhead in place.

Power goes to the GUST LOCK switch fromNo. 2 dc primary bus interlocked throughweight-on-wheels and transmission low oilpressure relays, so the GUST LOCK can onlyengage on the ground. When the index unitactuator retracts fully, a limit switch within theactuator activates a second relay inside the panel.This relay turns on a GUST LOCK caution lighton the caution advisory panel.

DC power extends or retracts the index unit.Three-phase ac power operates the index motor.

When the rotor head is indexed, logic circuitsin the bladefold control panel turn on anINDEXED status light. The logic also senses thecondition of the main rotor blades, tail pylon,stabilator, and rotor brake. If main rotor bladesare spread and indexed, pylon and stabilator arein flight position. With the rotor brake on, a pitchblades command goes to the DAFCS computer.

The computer flashes the BAR ALT push-button legend on the AFCS CONTROL panel.When the operator presses this push-button for2 seconds, the TRIM push-button legend goes on.This causes the computer to send flight controlpositioning signals to the pilot-assist servos. Thepilot-assist servos move the flight controls to acomputer memorized position.

With controls positioned, the computer sendsan enable signal to the pitchlock logic in thebladefold control panel indicated by the flashingTRIM push-button legend. The control paneldelivers a pitchlock command through the slipring to the bladefold deice distributor. This powergoes to the distributor for pitchlock actuators andbladefold actuators. This 115-Vac power also goesto the pitchlock actuator on each rotor blade hub.Each actuator advances a Iockpin that locks eachblade pitch control horn.

When all control horns lock, a signal goesthrough the slip ring to the bladefold controlpanel. This turns on a PITCH LOCKED statuslight.

As each blade’s pitch lockpin engages itsrespective pitch control horn, the bladefoldactuator operates. Each actuator, through

gearing, unlocks its respective blade by pulling outtwo blade lockpins.

Once the blade lockpins pull out, wormgearing drives a segment gear, which folds backeach blade. Switches on each blade sense whenthe blade lockpins pull out and each blade reachesa folded position. When all four blades actuatethe bladefold switches, a signal goes from logicin the distributor. This signal flows through theslip ring to the bladefold control panel, bringingon a FOLDED status light.

Each blade has individual logic circuits in abladefold module. Should any individual logiccircuit fail, the associated blade will not fold. Thisdoes not affect the other blades in the system, andthey will fold normally. A spread blade can foldmanually.

If a loss of command signals or power occurs,this will prevent all blades from folding. A cyclecaution logic circuit incorporated in the distributormonitors failures that are common to all blades.The cycle caution output goes to test points in thebladefold module for troubleshooting and faultisolation of the system.

SPREAD SEQUENCE

You accomplish main rotor blade spreadingby setting the BLADEFOLD switch to SPREAD.The switch routes a spread command through theslip ring to the blade deice distributor. The spreadcommand also goes to the junction box andenergizes a relay. This causes the power to thedistributor circuits to reverse phase, allowing thebladefold actuators to run in the spread direction,spreading each rotor blade. As each blade reachesthe full spread position, two blade lockpins driveinto the blade hinge lugs.

Two switches sense lockpin positions. Whenboth switches actuate, a relay in the distributorenergizes and reverses the phase of power to theactuators. The actuators run in the fold conditionagain for a brief period. This relieves stresses builtup on the segment gear in each bladefold actuator.

Running the actuators in the fold directioncauses the blade lockpins to back slightly out ofthe hinge lugs. One of the blade lockpin switchesactuates before the other. This switch remainsactuated even though the lockpins are backing outof the hinge lugs. The other switch deactuateswhen the lockpins back out. When one lockpinswitch actuates, and the other is deactuated,distributor logic determines achievement of thefull spread condition. This de-energizes thebladefold motor actuator on each blade.

Because the lockpins pull out only a shortdistance, they still safely lock the blades in flightposition. Logic circuits in the distributor sense thefully spread conditions, spread command, andpitch lockpin IN condition for each blade. At thispoint, a relay actuates that pulls out the pitchlockpin in each blade pitch control horn.Simultaneously, a signal goes to a control panelspread complete logic circuit. This circuit actuatesa 2-second time-delay relay. The relay shuts offhydraulic pressure to the pilot-assist servos andprovides a logic signal to the DAFCS computer.

The logic signal to the DAFCS computer tellsthe computer to update flight control positioninformation from the transducers for 2 seconds.This is accomplished when one stage of hydraulicpressure is off, preventing preloading of the flightcontrols. Any preloading of the controls wouldcause an error in position information to thecomputer.

After the 2-second delay, the computer blinksthe RDR ALT push-button legend on the AFCSCONTROL panel. This informs the operator thatthe computer has accepted the flight controlposition update. The computer will place the flightcontrols in the new memorized position during thenext fold cycle.

PYLON AND STABILATOR FOLD

The first step in manually folding the pylonis to complete bladefold operation. By doing this,you prevent rotor blade and pylon damage.

The crack open switch receives 28 Vdc, whichcomes through contacts of the weight-on-wheelsrelay. When the pylon opens about 5 degrees, thecrack open switch applies power to the tail rotorblade positioner actuator. The actuator extends,indexes, and gust locks the tail rotor. It extendsuntil the internal extend limit switch disconnectspower. At the same time, another internal switchapplies an actuator extended signal to thebladefold control panel logic circuits.

If main bladefolding is operating at the sametime as pylon folding, the actuator extended signalstops main fold. This prevents the main bladesfrom striking the pylon.

Two stabilator lockpin switches close when thestabilator locks in the flight (spread) position. Thisapplies 28 Vdc to the automatic engagementcircuit in the stabilator amplifier.

With the stabilator lockpin switches innonflight position, power is removed from theNo. 1 and No. 2 stabilator amplifier auto engagecircuitry. The auto engage circuitry engages when

the stabilator AUTO CONTROL PUSH TORESET pushbutton is depressed. DAMAGEMAY OCCUR TO THE FUSELAGE IF THESTABILATOR IS IN THE FOLD POSITION.At the same time, the bladefold control panel logiccircuits receives a ground signal.

With the main rotor blades properly folded,the bladefold control panel logic circuitry causesthe SPREAD INCOMPLETE caution capsule tolight under the following conditions:

• Power goes to the fold or spread relayswithout being commanded by the blade-fold control panel.

• Pylon or stabilator panel spread but notlocked.

• Tail rotor blade positioner actuatorextends with the pylon in the flightcondition.

Q1. In the H-60 bladefold system, thetransmission low oil pressure relay isenergized by a

Q2. When the rotor head is on the insulated areabetween index drive slip ring segments 180degrees from the index position, in whatdirection will the rotor head travel towardsindex?

Q3. During bladefold operations, if one bladefails to fold, will the other blades fold?

Q4. During blade spread operations, theactuators run in the fold condition afterlockpin position switches activate to

Q5. During pylon and stabilator fold, at whatpoint does the crack open switch power thetail rotor blade position actuator?

VARIABLE INLET DUCT RAMPSYSTEM

Learning Objective: Recognize operatingprinciples and features of variable inletduct ramp systems.

High-speed aircraft usually operate in pre-determined Mach numbers instead of specificairspeeds. The Mach number is the ratio of thespeed of an object to the speed of sound in thesame medium and at the same temperature. Sonicvelocity and Mach number vary with air tempera-ture; therefore, at standard day conditions, theairspeed that corresponds to a given Mach numbervaries with changes in altitude.

In aircraft that fly at speeds of Mach 2.0 andgreater, the air velocities at the inlet duct are muchhigher than the engine can efficiently use. Thevelocity of the air at the inlet duct entrance mustbe decreased before entering the engine to preventengine stall. Some aircraft accomplish this byusing a variable inlet duct ramp. The duct rampis a moving part of the surface attached to theleading edge of the engine inlet duct. It positionsthe supersonic shock waves to allow only subsonicair to enter the duct. Through wind tunnel tests,the exact ramp position necessary at any particularspeed is determined.

ENGINE INLET BLEED AIRSYSTEM

Learning Objective: Recognize operatingprinciples and characteristics of aircraftengine inlet bleed air systems.

The engine inlet bleed air system of the F/A-18provides the best inlet airflow for subsonic andsupersonic flight. The inlet bleed air doors controlinlet airflow by opening at 1.33 Mach to bleedboundary layer air off the compression ramp andclose at 1.23 Mach. This makes sure the doors willnot oscillate when the aircraft travels at 1.33Mach. A bleed air channel controls the fuselageboundary layer airflow near the wing roots.

The bleed air doors open when three-phase,115-Vat power is applied, and the controllingFLIGHT CONTROL COMPUTER (FCC) has aloss of power. After normal engine shutdown, thebleed air doors normally close. In some modelsof the F/A-18 aircraft when external power

is on or the APU is operating in the groundmaintenance mode and FCCs off, the bleed airdoors will open. When FCCs are turned on, thebleed air doors will close. The FLT CONTR GNDPWR RELAYS No. 11 and No. 12 prevent thecycling of the bleed air doors during groundoperations.

The Air Data Sensor, DT-600/ASW-44,computes the free air Mach number and sends thesignal to FCCA and FCCB. When the Machnumber is below 1.23, the FCCs supply a ground,energizing the retract circuits. Each actuatortorque motor energizes, and the actuator brakesrelease. The torque motors retract the bleed airdoors (if doors are not retracted) until the actuatorretract limit switches close. The retract limitswitches de-energize the retract circuits, and theactuator brakes are applied by spring pressure.

When the Mach number is above 1.33, theFCCs remove the ground, de-energizing thecontrol relays and energizing the extend circuits.Each actuator’s torque motor energizes and theactuator brakes release. The torque motors extendthe bleed air doors (if doors are not extended) untilthe actuators extend limit switches close. Theextend limit switches de-energize the extendcircuits, and the actuator brakes are applied byspring pressure.

The extend limit and retract limit switches ineach bleed air door actuator makes a ground tosignal the signal data converter when the inletbleed door fully extends or retracts. When groundis removed, the signal data converter measures thetime in transient until the other limit switch closes.If the time in transient exceeds 8 seconds, thesignal data converter signals the digital computer.The digital computer commands the cockpit DDIto display a caution.

Q1. The ratio of the speed of an objectto the speed of sound in the same mediumand at the same temperature is known as

Q2. The F-18 engine bleed air system duct doorsare controlled by

CHAPTER 6

AIRCRAFT INSTRUMENTS

When the first aircraft came into existence, themain goal was to launch the aircraft and keep itairborne as long as possible. At first, it was notpossible to keep the aircraft in the air for longerthan a few minutes. However, as engines andaircraft structures were improved, the aircraft wasable to remain aloft for a longer time. Alongwith these improvements came the need forinstruments. The first aircraft instruments werefuel and oil pressure instruments. These instru-ments warned the pilot of engine trouble so theaircraft could be landed before the engine failed.Later, when the aircraft could fly over con-siderable distances, weather became a problem.This led to the development of instruments thathelped pilots fly through snowstorms, thunder-storms, and other bad weather conditions.

The instruments used in aircraft years ago arereasonably simple compared with those in currentaircraft. The jet aircraft has brought manycomplex problems to instrument engineering.

Instrumentation is basically the science ofmeasurement. Measurements that are common onall aircraft are posit ion, direction, speed, altitude,engine condition, fuel on board, and fuelconsumption. In addition, jet aircraft instrumentsinclude Mach number, angle of attack, and tailpipe temperature indicators.

There are two ways of grouping aircraftinstruments—by their operating principles and bythe job they perform. Instrument operatingprinciples include gyroscopic, pressure ortemperature sensing, magnetism, electrical energy,or a combination of any of these. This chapterdeals with instruments and indicating systems inrelation to the jobs they perform—flight, engine,and equipment instruments. The flight instru-ments discussed are those instruments that provideaircraft performance information to the pilot.These instrument systems include the airspeed,altimeter, vertical speed, attitude, turn and bank,and angle-of-attack. Along with the headingindicator, these instruments provide primary flightreference to the pilot. The heading indicator is a

flight instrument, which is discussed in the inertialnavigation systems chapter.

FLIGHT INSTRUMENT SYSTEMS

Learning Objective: Recognize operatingprinciples and features of various aircraftflight instrument systems, including thepitot-static, airspeed indicator, angle-of-attack, gyroscope, and miscellaneous flightinstrument systems.

To maintain instruments properly, you, as anAE, must know the basic principles of the flightinstrument systems. AEs frequently work withequipment and systems that use the principles ofdensity and pressure.

You must consider density and pressure whendiscussing altimetry and airspeed. Although verylight, air has weight and is affected by gravity.By its weight, air exerts pressure on everythingit touches. Since air is a gas, it exerts pressure inall directions. The weight of the air pressing downfrom above determines the air pressure at anygiven altitude.

The weight of the atmosphere presses themolecules closer together, making them morenumerous per unit of volume. This action takesplace at the bottom of the atmosphere, or whereit rests upon the earth’s surface. Therefore, theair at the bottom of the atmosphere is more densethan at higher altitudes. Air pressure at sea levelon an average day will support a column ofmercury 29.92 inches high (fig. 6-1).

Atmospheric pressure is a force per unitarea, and force is equal to mass multiplied byacceleration. Therefore, a pressure change occursif either the mass of the atmosphere changes orthe molecules within the atmosphere accelerate.Although altitude exerts the dominant control,temperature and moisture alter pressure at any

Figure 6-1.-Mercurial barometer.

given altitude. Figure 6-2 shows the standardpressure and temperature at given altitudes.

Conditions are very seldom standard fortemperature or pressure; therefore, you mustcorrect the formula to find density altitude or trueairspeed. Let’s consider an airfield under theinfluence of a low-pressure climatic condition,where the temperature is very hot. Together, thesetwo conditions may reduce the density of the airto such an extent that it affects aircraft engineperformance. This makes takeoff capabilitymarginal, especially for a helicopter. The densityof air also directly affects aircraft movementthrough the air, and thus the true airspeed of theaircraft. The denser the air, the more difficult itis for the aircraft to move through it.

PITOT-STATIC SYSTEM

The aircraft pitot-static system (fig. 6-3)includes instruments that operate on the principleof the barometer. The system consists of a pitottube, static air vents, and three indicators, whichconnect with pipelines that carry air. The threeindicators are airspeed, altimeter, and the verticalspeed indicator (VSI).

The airspeed indicator shows the speed of theaircraft through the air and the altimeter shows

Figure 6-2.-The standard atmosphere.

the altitude. The VSI indicates how fast theaircraft is climbing or descending. They all operateon air that comes in from outside the aircraftduring flight.

The pitot tube mounts on the outside of theaircraft (fig. 6-3) at a point where the air is leastlikely to be turbulent. The tube points in aforward direction parallel to the aircraft’s line offlight. One general type of airspeed tube mountson a streamlined mast extending below the noseof the fuselage. Another type mounts on a boomextending forward of the leading edge of the wing,Although there is a slight difference in theirconstruction, they operate identically.

Pitot stands for impact pressure, the pressureof the outside air against the aircraft flyingthrough it. The tube that goes from the pitot tubeto the airspeed indicator applies the outside airpressure to the airspeed indicator. The airspeed

Figure 6-3.-Pitot-static system.

indicator calibration allows various air pressuresto cause different readings on the dial. Thepurpose of the airspeed indicator is to interpretpitot air pressure in terms of airspeed in knots.

Generally, static air vents (fig. 6-3) are smallcalibrated holes in an assembly mounted flushwith the aircraft fuselage. Their position is in aplace with the least amount of local airflowmoving across the vents when the aircraft is flying.

Static means stationary or not changing. Thestatic part of the pitot-static system also introducesoutside air. However, it is at its normal outside

atmospheric pressure as though the aircraft werestanding still in the air. The static line applies thisoutside air to the airspeed indicators, thealtimeter, and the vertical speed indicator.

Airspeed Indicators

Readings from an airspeed indicator areuseful. They are used to estimate ground speedand to determine throttle settings for the mostefficient flying speed. These readings also providea basis for calculating the best climbing andgliding angles. They warn the pilot if diving speedapproaches the safety limits of the aircraft’sstructure. Since airspeed increases in a dive anddecreases in a climb, the indicator is an excellentcheck for maintaining level flight. Figure 6-4

Figure 6-4.-(A) Airspeed indicator; (B) maximum allowableairspeed indicator.

shows a cutaway view of a typical airspeedindicator.

An airspeed indicator has a cylindrical, airtightcase that connects to the static line from thepitot-static tube. Inside the case is a small aneroiddiaphragm of phosphor bronze or berylliumcopper. The diaphragm is very sensitive tochanges in pressure, and it connects to theimpact pressure (pitot) line. This allows air fromthe pitot tube to enter the diaphragm. The sideof the diaphragm fastens to the case and is rigid.The needle or pointer connects through a seriesof levers and gears to the free side of the

The airspeed indicator is a differential pressureinstrument. It measures the difference between thepressures in the impact pressure line and in thestatic pressure line. The two pressures are equalwhen the aircraft is stationary on the ground.Movement through the air causes pressure in theimpact line to become greater than that in thestatic line. This pressure increase causes thediaphragm to expand. The expansion or con-traction of the diaphragm goes through a seriesof levers and gears to the face of the instrumentto regulate needle position. The needle shows thepressure differential in MPH or knots. (All speeds

diaphragm. and distances are in nautical miles.)

Figure 6-5.-Airspeed/Mach indicator.

MAXIMUM ALLOWABLE AIRSPEEDINDICATOR. —Figure 6-4, view B, shows theface of a maximum allowable airspeed indicator.The dial face measurements are in knots from 50to 450 with an expanded scale below 200 knots.The dial has an indicating pointer and a maximumsafe airspeed pointer. The maximum safe airspeedpointer moves as the maximum safe airspeedchanges because of static pressure changes atdifferent altitudes.

No matter where the pitot-static tube islocated, it is impossible to keep it free from allair disturbances set up by the aircraft structure.You must make allowances for this installationerror when reading the indicator. Temperature isanother cause of error. Also, imperfect scaling ofthe indicator dial with respect to the air-speed-differential pressure relationship will causean error in reading. You can make simpleadjustments to the instrument mechanism tocorrect the tendency to read fast or slow.

MACH NUMBER INDICATORS. —In somecases, the term Mach number is used to express

aircraft speed. The Mach number is the ratio ofthe speed of a moving body to the speed of soundin the surrounding medium. For example, if anaircraft is flying at a speed equal to one-half thelocal speed of sound, it is flying at Mach 0.5. Ifit moves at twice the local speed of sound, itsspeed is Mach 2. (The term Mach number comesfrom the name of an Austrian physicist, ErnestMach, pioneer in the field of aerodynamics.)

Figure 6-5 shows the front view of a typicalairspeed and Mach number indicator. Theinstrument consists of altitude and airspeedmechanisms incorporated in a single housing. Thisinstrument gives the pilot a simplified presentationof both indicated airspeed and Mach number.Both indications are read from the same pointer.The pointer shows airspeed at low speeds, andboth indicated airspeed and Mach number at highspeeds. Pitot pressure on a diaphragm moves thepointer, and an aneroid diaphragm controls theMach number dial. The aneroid diaphragm reactsto static pressure changes because of altitudechanges. Figure 6-6 is a mechanical schematic ofan airspeed and Mach indicator.

Figure 6-6.-Airspeed/Mach number indicator mechanical schematic.

The range of the instrument is 80 to 650 knots measure the altitude. In standard day conditions,indicated airspeed and from 0.5 to 2.0 Mach pressure altitude and true altitude are the same.number. Its calibrated operating limit is 50,000feet of altitude. A stationary airspeed dial masks ABSOLUTE ALTITUDE —Absolute altitude

the upper range of the movable Mach dial at low is the distance between the aircraft and the terrain

altitudes. The stationary airspeed dial is graduated over which it is flying. It is referred to as the

in knots. The instrument incorporates a landing altitude above ground level (AGL). AGL is

speed index and a Mach number setting index. You usually useless information because of variations

can adjust both indexes by a knob on the lower in the terrain. However, it is useful when flying

left-hand corner of the instrument. You can adjust near the ground; for example, in a takeoff or

the landing speed index over a range of 80 to 150 landing pattern. You find AGL by subtracting the

knots. The index operates with the knob in its elevation of the terrain beneath the aircraft from

normal position. You may adjust the Mach number the altitude read on the altimeter (MSL). A radar

index over the entire Mach range. The index altimeter indicates actual altitude above the

adjusts by depressing the knob and turning it. terrain; you call this indication radar altitude.

Altimeter

An altimeter is an instrument that measuresstatic pressure. Before you can understand howthe altimeter works, you need to understandaltitude. Remember, even though the altimeterreads in feet, it is actually measuring pressure.

The word altitude is vague, so it needs furtherdefining. The term altitude includes altitude abovemean sea level (MSL) and altitude above groundlevel (AGL). It also, includes pressure altitude,indicated altitude, density altitude, and elevation.

MEAN SEA LEVEL. —Since about 80percent of the earth’s surface is water, it is naturalto use sea level as an altitude reference point. Thepull of gravity isn’t the same at sea level all overthe world because the earth isn’t perfectly roundand because of tides. To adjust for this, anaverage (or mean) value is set; this is the meansea level. Mean sea level is the point where gravityacting on the atmosphere produces a pressure of14.70 pounds per square inch. This pressuresupports a column of mercury in a barometer toa height of 29.92 inches. This is the reference pointfrom which you measure all other altitudes. Seefigures 6-1 and 6-2. The altitude you read froman altimeter refers to MSL.

ELEVATION AND TRUE ALTITUDE. —Elevation is the height of a land mass above meansea level. Elevation is measured with precisioninstruments that are far more accurate than thestandard aircraft altimeter. You can find elevationinformation on charts or, for a particular spot,painted on a hangar near an aircraft ramp or taxiarea.

True altitude is the actual number of feetabove MSL. A ruler or yardstick is used to

PRESSURE ALTITUDE. —It is not possibleto have a ruler extending from an aircraft andreaching to sea level to measure altitude. Tomeasure altitude, instruments sense air pressureand compare it to known values of standard airpressure at specific, measured altitudes. Thealtitude you read from a properly calibratedaltimeter referenced to 29.92 inches of mercury(Hg) is the pressure altitude.

Refer back to figure 6-2. If a pressure altimetersenses 6.75 pounds per square inch pressure with thealtimeter set to sea level and barometric pressure29.92 inches of mercury, the altimeter indicates20,000 feet. This does not mean that the aircraft isexactly 20,000 feet above mean sea level. It meansthe aircraft is in an air mass exerting a pressureequivalent to 20,000 feet on a standard day. Youcan see that pressure altitude is not true altitude.

INDICATED AND CALIBRATED ALTI-TUDE. —Unfortunately, standard atmosphericconditions very seldom exist. Atmosphericconditions and barometric pressure can varyconsiderably. A pressure change of one-hundredth(0.01) of an inch of mercury represents a 9-footchange in altitude at sea level. Barometricpressure changes between 29.50 and 30.50 are notuncommon (a pressure change of about 923 feet).Indicated altitude is the uncorrected reading ofa barometric altimeter. Calibrated altitude is theindicated altitude corrected for inherent andinstallation errors of the altimeter instrument. Onan altimeter without such errors, indicated altitudeand calibrated altitude are identical. Assume thatthis is the case for the rest of this discussion.

When flying below 18,000 feet, the aircraftaltimeter must be set to the altimeter setting(barometric pressure corrected to sea level) of aselected ground station within 100 miles of theaircraft. Altitude read from an altimeter set tolocal barometric pressure is indicated altitude. The

accuracy of this method is limited because youmust assume a standard lapse rate; that is, for agiven number of feet of altitude, an exact changein pressure occurs. This seldom happens, whichlimits the accuracy of the altimeter. Above 18,000feet, all altimeters are set to 29.92 (pressurealtitude). Although the altimeter is not accurate,as long as all aircraft have the same barometricpressure setting, aircraft vertical separation iscontrolled.

DENSITY ALTITUDE. —A very importantfactor in determining the performance of anaircraft or engine is the density of the air. Thedenser the air, the more horsepower the enginecan produce. Also, there is more resistance to the

aircraft in flying through the air, and the airfoilsproduce more lift. Pressure, temperature, andmoisture content all affect air density. Measure-ments of air density are in weight per unit volume(for example, pounds per cubic foot). However,a more convenient measurement of air density forthe pilot is density altitude. This is the altitudeassigned to a given density in the standardatmosphere.

Density altitude does not show on an instru-ment. It is usually taken from a table or computedby comparing pressure, altitude, and temperature.Although moisture content affects air density, itseffect is negligible.

Figure 6-7 shows the different types ofaltitude. At Navy Memphis, the field elevation is

Figure 6-7.-Different types of altitude.

322 feet. The altimeter should read 322 feet whenproperly set to the local altimeter setting (in thiscase 29.67). If the altimeter is set to 29.92(standard day pressure at sea level), the altimeterwould read about 550 feet. The density altitudeis such that the aircraft and engine perform thesame as if they were at a standard day altitudeof 1,180 feet. This means the distances requiredfor takeoff and landing are longer for higherdensity altitudes (higher temperatures). Theabsolute altitude (AGL) for the aircraft atMemphis is zero feet because the aircraft istouching the surface.

Look at the altitude for Stapleton Inter-national Airport in Denver, Colorado (fig. 6-7).Notice that the difference between pressurealtitude and indicated altitude remains the sameas at Memphis. However, at the same temperaturethe density altitude is much greater at Denver.

Several kinds of altimeters are used today.They are all constructed on the same basicprinciple as an aneroid. They all have pressure-responsive elements (aneroid wafers) that expandor contract with the pressure changes of differentflight levels. The heart of a pressure altimeter isits aneroid mechanism (fig. 6-8), which consistsof one or more aneroid wafers. The expansion orcontraction of the aneroid wafers with pressurechanges operates the linkage. This action movesthe indicating hand/counter to show altitude.Around the aneroid mechanism of most altimetersis a device called the bimetal yoke. As the nameimplies, this device is composed of two metals.It performs the function of compensating for the

Figure 6-8.-Simplified aneroid mechanism.

effect that temperature has on the metals of theaneroid mechanism.

The altimeter discussed in the followingparagraphs is a simple one. Several complexaltimeters are discussed later in this chapter, alongwith the automatic altitude system.

COUNTER POINTER PRESSURE ALTIM-ETER. —The purpose of the counter pointerpressure altimeter (fig. 6-9) is to show aircraftheight. By studying the dial of the indicator, itis easy to understand the procedure for determin-ing the height of the aircraft. A description of themechanical operation of this altimeter follows. Asyou read about the operation, refer to figure 6-10.

Atmospheric changes cause movement of thetwo aneroid diaphragm assemblies (1). Theseassemblies move two similar rocking shaftassemblies (2) mutually engaged with the mainpinion assembly (3). This movement goes to thehandstaff assembly (4), which operates the handassembly (5) and drives the counter mechanism(6) through a disk (7). Because of the specialdesign of the hand assembly (5), the counterindication is never obscured. An internal vibrator(8) minimizes friction during the instrument’soperation.

You make barometric corrections by turningthe externally located knob (9). The knob engagesthe barometric dial (10) and the main plateassembly (11) that supports the entire mechanism.You make adjustments so the reading on thebarometric dial corresponds to the area baro-metric conditions in which the aircraft is flying.

Figure 6-9.-Counter pointer pressure altimeter.

Figure 6-10.-Mechanical schematic of a counter pointer pressure altimeter.

Vertical Speed Indicators (VSI)

A vertical speed indicator shows the rate atwhich an aircraft is climbing or descending. It isvery important for night flying, flying throughfog, clouds, or when the horizon is obscured.Another use is to determine the maximum rateof climb during performance tests or in actualservice.

The rate of altitude change, as shown on theindicator dial, is positive in a climb and negativein a dive or glide. The dial pointer (fig. 6-11)moves in either direction from the zero point. Thisaction depends on whether the aircraft is goingup or down. In level flight the pointer remainsat zero.

The vertical speed indicator is contained in asealed case, and it connects to the static pressureline through a calibrated leak. Look at figure 6-12.You can see that changing pressures expand orcontract a diaphragm, which moves the indicatingneedle through gears and levers. The instrumentautomatically compensates for changes intemperature. Although the vertical speed indicatoroperates from the static pressure source, it is adifferential pressure instrument. The differencein pressure between the instantaneous staticpressure in the diaphragm and the static pressure

Figure 6-11.-Vertical speed indicator (VSI).

Figure 6-12.-Mechanical schematic of a VSI.

trapped within the case creates the differentialpressure.

When the pressures equalize in level flight, theneedle reads zero. As static pressure in thediaphragm changes during a climb or descent, theneedle immediately shows a change of verticalspeed. However, until the differential pressurestabilizes at a definite ratio, indications are notreliable. Because of the restriction in airflowthrough the calibrated leak, it requires a 6 to9-second lag for the pressures to stabilize.

The VSI has a zero adjustment on the frontof the case. You use this adjustment with theaircraft on the ground to return the pointer tozero. While adjusting the instrument, tap it lightlyto remove friction effects.

AIR DATA COMPUTERSYSTEM (ADC)

Aircraft operating below 0.9 Mach airspeeduse raw pitot and static pressures to developaccurate airspeed, altitude, and vertical speedindications. Aircraft operating in this speed rangeuse the pressures that the pitot-static portssense.

Modern supersonic aircraft operate in a higherspeed range and require more accurate pressures.At high speed, pressures build upon the externalskin of the aircraft. These pressures cause adistortion of the normal flow of air, causing the

pitot-static system to sense false pressures. Thesystem then supplies erroneous information to theflight instruments. The altimeter, for instance,may show an error of more than 3,000 feet. A3,000-foot error in altitude is intolerable andcould put an aircraft in an extremely dangerousposition.

The system that compensates for altitude andother pitot-static errors is the air data computersystem (ADC system). Many variations exist inboth the name of the systems and the method ofdata development. The system in the F-14 isstrictly a digital computer known as the centralair data computer (CADC). The system in the S-3is an airspeed altitude computer set (AACS). TheAACS is a digital computer, but it is very differentfrom the CADC used in the F-14.

Purpose

Many inputs are common to the various typesof ADC systems. Air data computers differ in

how they process input data and distribute outputdata to the various systems using the data. Datarequirements vary with the type and mission ofthe aircraft.

The F-14 CADC system contains a single-processor digital computer with a separateindependent analog backup wing sweep channel.The CADC can make yes and no decisions andsolve mathematical problems. It converts outputsto either digital or analog form, depending onwhich aircraft system is to use the information.The CADC gathers, stores, and processes signalsrepresenting pitot pressure, static pressure, totaltemperature, and angle-of-attack data derivedfrom the aircraft airstream sensors.

Figure 6-13 shows the functional relationshipbetween the CADC (digital processor), its inputs,and the various functions in each aircraft systemusing the data. The digital processor performswing-sweep, glove-vane, and flap-and-slatschedule computations. It also performs limit-control and electrical interlocks, failure detection,

Figure 6-13.-F-14 air data computer system.

and systems test logic. Figure 6-14 shows inputspower. This power is from essential bus number tothe CADC and the CADC outputs to the 2 (phasesA, B, and C). The system also uses systems, whichdepend on or in part upon CADC single-phase,

115-volt ac, 400-hertz power. This functions. Noticein figure 6-14 that the CADC power is also fromone phase of the essential bus system uses three-phase, 115-volt ac, 400-hertz and goes to thebackup channel. Cooling

Figure 6-14.-Air data computer block diagram showing inputs and outputs.

equipment for the ADC system must be availablewhen the system is operating.

NOTE: To understand air data computermaintenance in modern aircraft, you must havea knowledge of digital electronics, including logicdiagrams and flow charts. Review Navy Electricityand Electronics Training Series (NEETS),module 13, NAVEDTRA 14155, and DigitalComputer Basics, NAVEDTRA 10088, beforecontinuing.

The air data computer receives informationfrom pressure-sensitive and temperature-sensitiveunits mounted on external points of the aircraft.Then, it compensates for errors and sends thecorrected information to other systems in theaircraft. Concurrently, it detects any changes inpressure and temperature information. It convertsthese changes into usable signals and sends themalong with the pressure and temperature signals.

The electrical signal outputs are representative ofaltitude, Mach number, true airspeed, angle ofattack, total temperature, and impact pressure.There is also a pneumatic output of correctedstatic pressure. This output is used by thebarometric altimeter, airspeed, and vertical speedindicators, and by some modules within the airdata computer.

There are four data inputs to the ADC:

1. Total pressure (pitot)2. Indicated static pressure3. Indicated angle of attack4. Total temperature

Command and test signal inputs use the availableraw and corrected primary data for makingfunctional tests of various ADC outputs.

Table 6-1 is a list of symbols and theirrespective definitions. Since these symbols are

Table 6-1.-Symbols Used with an ADC System

used many times in this section, you should referto this table for symbol meanings.

Figure 6-15 shows the major distribution ofsystems that depend on all or part of the ADC.Notice that all inputs, such as pitot and staticpressures, go to the digital processor. Signalsresulting from the processing of the inputs go tothe using systems. Although the diagram infigure 6-15 shows pneumatic inputs only, bothpneumatic and electrical inputs are used.

Major ComponentsFigure 6-16.-Angle of attack transmitter (probe).

The four major components that collect anddistribute information used in the air datacomputer system are listed below. aircraft model. This way you will be sure the

1.2.3.4.

information on the ADC system is current.Angle-of-attack transmitterPitot-static system ANGLE-OF-ATTACK TRANSMITTER. —Total temperature probe Forces vary with the angle of attack. The angle ofAir inlet control system attack is the angle between the relative wind and

As you know, the functions of these com-ponents are essentially the same on all aircraft.Only the processing and distribution of air datainformation varies from aircraft model to model.When performing maintenance on any ADCsystem, you should refer to the latest maintenanceinstructions manual (MIM) for that particular

the chord of the wing. (The chord of the wing isa straight line running from the leading edge tothe trailing edge.) Increasing the angle of attackincreases the pressure felt under the wing and viceversa.

The angle-of-attack transmitter (fig. 6-16)detects changes in the aircraft’s local angle ofattack. It sends these changes, in the form of

Figure 6-15.-ADC (digital processor) inputs and outputs.

mechanical motion, to potentiometers within the drives the wiper arms of the three internallytransmitter. These potentiometers convert the mounted potentiometers.mechanical motion to proportional electrical The angle-of-attack system shows the pilotvoltages. These voltages go to associated angle- aircraft pitch attitude with respect to theof-attack indicating and interface equipment. surrounding air mass.

The transmitter has a detector probe thatsenses changes in airflow. Changes in airflow PITOT-STATIC SYSTEM. —Figure 6-17cause the probe paddle to rotate. This, in turn, shows the airstream sensors of the pitot-static

Figure 6-17.-ADC system airstream senors.

system. These sensors sense the air surroundingthe aircraft and provide impact (pitot/Pt) pressureand atmospheric (static/Ps) pressure. Thesepressures go to the flight instruments, the air datacomputer, and the engine air inlet control system(AICS) programmers. The pitot-static system isactually two separate systems with individualpitot-static probes (fig. 6-18), one on each sideof the forward fuselage. The ADC receives staticpressure (Ps) from both probes. However, itreceives total pressure (Pt) from only one probe.

Indicated Static Pressure. —This pressure (P)is the atmospheric pressure as sensed at a pointon the aircraft that is relatively free from airflowdisturbances. At subsonic speeds, static pressureerror is small and of little significance. However,at transonic and supersonic speeds, the static portssense extreme static pressure errors. Both Machnumber and angle of attack can cause significanterrors in the static pressure system. Indicated staticpressure (P), as detected by the aircraft staticports, deviates from true static pressure. Thesedeviations have a definite relationship to Mach

number and angle of attack. The size of the erroris the ratio of true static pressure to indicated staticpressure, as related to Mach number and angleof attack.

Impact Pressure. —As implied, impactpressure (Qc) is the force of the air against theaircraft. Qc is measured directly by use of apitot-static probe (fig. 6-18) or calculated fromthe outputs of the static and total pressuretransducers. The ADC calculates actual impactpressure (QA) as a function of Mach numbersquared and static pressure.

Indicated Total Pressure. —This pressure (Pti)is the sum of static air pressure and the pressurecreated by aircraft motion through the air. Thepitot tube senses Total pressure, which you alsoknow by the familiar term pitot pressure.

Corrected Static and Corrected Total Pres-sures. —These pressures, Ps and Pt, contain errorsthat must be corrected to get true static and truetotal pressures. These errors are a result of slope

Figure 6-18.-Pitot-static probe.

and offset errors related to Mach speeds. The computercalculates the specified slope and intercept errors asfunctions of the indicated pressure ratio (Pti/P) and of theindicated angle of attack (ai).

CAUTION

Be sure to disable the pitot-static heaterbefore working on this system. You may beseriously burned by touching the probes.

TOTAL TEMPERATURE PROBE.—Totaltemperature (Tt) is the temperature of ambient air plusthe temperature increase created by the motion of theaircraft. Total temperature is sensed by a probe. Thisprobe includes a platinum resistance element inside anaerodynamic housing placed in the airstream. Theresistive element, whose resistance varies withtemperature, acts as the variable portion of a bridgecircuit.

The total temperature probe provides the ADC withaccurate outside air temperature information. The rawinformation is the indicated total temperature (Tti). Thecomputer smoothes and limits computations on the Ttibefore using the resultant output to calculate true totaltemperature (Tt). Figure 6-19 shows a typicaltemperature-sensitive bridge circuit that providestemperature data to the air data computer.

AIR INLET CONTROL SYSTEM.—The air inletcontrol system (AICS) decelerates supersonic air. It alsoprovides even, subsonic airflow to the engine throughoutthe flight envelope. The AICS is similar to the variableinlet duct ramp system mentioned in chapter 5. Threeautomatically controlled hinged ramps on the upper sideof the intake vary inlet geometry. The ramps position todecelerate supersonic air by creating a compression field

Figure 6-19.-Total temperature circuit.

outside the inlet. They also regulate the amount andquality of air going to the engine. Electrohydraulicactuators, which respond to fixed schedules in the AICSprogrammers, position the ramps.

An electrical total pressure (Pt) input from the rightAICS programmer goes to the ADC backup channel asairspeed indications for wing sweep.

AUTOMATIC ALTITUDE SYSTEM

In the past, the air traffic control system radarpresented azimuth and distance information to thecontroller on a horizontal radarscopes. Aircraftidentification was primarily by voice radio, the use ofposition reports over definite fixes, making identifyingturns of the aircraft to headings requested by thecontroller, or by a beacon identification signal. Altitudeinformation was given over the voice radio. After thisinformation-gathering process, the information wasrecorded on a flight strip by the controller and updated asrequired. When the aircraft moved into anothercontroller’s area, the handoff of the aircraft and theassociated information was basically a manual process.Although the system was adequate, it becamecumbersome during heavy traffic.

The increase in air traffic since 1950 has causedserious problems of vertical separation, terrain clearance,and collision avoidance. Because of these problems,improved air traffic control techniques were developed.These techniques included the use of altitude-codedtransponders for automatic altitude and positionreporting.

Automatic altitude reporting equipment thatprovides continuous automatic identification of aircraft onthe ground controller’s radarscopes has been developed.This equipment cuts out many of the manual stepsrequired in the old air traffic control system. An air datacomputer corrects static pressure errors and providessynchro-driven altitude information to the pilot’saltimeter. It also provides altitude in digital form to theaircraft transponder in high-performance aircraft. In low-performance aircraft, the equipment provides a directreadout of altitude to the pilot and digital altitudeinformation to the aircraft transponder. The digitalinformation then goes to the ground interrogator andshows on the radarscopes in alphanumeric form.

The automatic altitude system operation is discussedin the following paragraphs. An interrogation pulse groupgoes from the interrogator-transmitter unit through adirectional interrogator antenna assembly. The pulsegroup triggers an airborne transponder, causing amultiple pulse reply group to be transmitted. The

transponder transmission goes to the groundinterrogator-receiver, which is processed through acomputer. It is then displayed in alphanumeric formon the controller’s radar screen. The length of theround-trip transit time determines the range of thereplying aircraft. The mean direction of the mainbeam of the interrogator antenna during the replydetermines the azimuth. The encoded signal from thetransponder provides, via mode C, the aircraft’saltitude in 100-foot increments.

Look at figure 6-20, which shows the automaticaltitude reporting system. As you can see, asemiautomated air traffic control system includes thefollowing improvements over the past system:

• It automatically provides the air traffic controllerwith a radar presentation. It identifies, in threedimensions, every properly equipped aircraft withinhis/her control area.

• As a result of the three-dimensional presentation,

it greatly reduces the use of voice radio. It also easesthe workload of the air traffic controller, thusincreasing air traffic control efficiency.

• A transponder signal reinforces the radar signalnormally seen on the radarscopes. It makes thesignal stronger and much less susceptible toatmospheric interference.

• The beacon system altitude reporting feature may

reduce vertical separation in the higher flight levels. • It continuously updates aircraft altitude records

in 100-foot increments. This permits more ac-curate traffic control when aircraft are chang-ing altitude rapidly, as they do in terminal areas.

The AIMS Program

The AIMS Program, implemented by theDepartment of Defense, derives its name from thefollowing acronyms:

ATCRBS (air traffic control radar beacon system)

IFF (identification friend or foe)

MARK XII Identification System

Systems (reflecting the many AIMS configurations)

Figure 6-20.—The automatic altitude reporting system.

The primary goal of the AIMS program wasto improve air traffic control within the UnitedStates. DoD participated in the first phase of theFAA plan. The plan was to upgrade peacetimeoperational use of the airspace. It included theprovision of additional identification and altitudesignals to ground air traffic control stations. Useof this plan resulted in a major improvementin military air traffic control. The secondaryobjective is to classify equipment configurationsand capabilities.

The AIMS system consists of transponders,interrogators, computers, servoed altimeters,altimeter-encoders, control panels, and associatedequipment. Interrogators are found in mostAIMS-equipped ground and surface sites, tacticalground and surface systems, and certain specialtask airborne vehicles. To avoid misreadingsand the friction problems associated withcurrent aircraft altimeters, counter-drum-pointeraltimetry displays are used.

Concurrent with the Department of Defensedecision to use the AIMS, the following toleranceswere established:

• Maximum difference between the pilot’sand ground displays, ±125 feet.

• Maximum difference between the pilot’sdisplay and true altitude, ±250 feet.

Tolerances during takeoff, instrument approach,and landing remain within existing military/FAAstandards.

Static-Pressure Errors

The altimeter system accuracy limitation of±250 feet required the elimination of, correctionof, or compensation for system errors. To achieveaccuracy, each aircraft type was treated separatelyto determine what corrections its pressure systemrequired within its flight envelope. Aircraft withsmall or negligible measured static-pressuredefects required only an instrument. Thisinstrument transforms a static-pressure input intoa digital electrical output of pressure altitude.

High-performance, subsonic aircraft withlarge position errors require an instrument tocorrect the static-pressure error as a function ofMach number. To correct the pressure error, itmust be known and be repeatable in aircraft ofthe same type. Since repeatability is a problemwith flush static ports, the ports were replacedwith a pitot-static tube.

High-performance, supersonic aircraft alsorequire an instrument capable of correcting forpressure errors as a function of Mach number.Because of the greater size of errors metover a wider range of altitude and speed, theinstrumentation in this type of high-altitudeaircraft is more complex than that of subsonicaircraft.

You can place errors produced in the measure-ment and transmission of true pressure altitudeinto three groups—measured altitude, altitudetransmission, and instruments (transducer orcomputer and digitizer).

ERRORS IN MEASURED STATIC PRES-SURE. —Errors in measured static pressure resultfrom many variables. Variables include the designof the pressure pickup, location of the pressurepickup on the aircraft, and the Mach number. Theangle of attack and the aircraft configuration alsocontribute to errors. The static-pressure sensor canbe either a flush fuselage vent or a static-pressure,L-shaped tube. When using a flush vent, theslightest surface irregularity or structure in thevicinity of the vent may deflect the perpendicularflow of air past the vent. With static-pressuretubes, you must give concern to other variables.Theses are the shape of the tube, the orificeconfiguration, and the type of structural support.In either case, there is a local pressure surroundingthe pressure-sensing vent that differs from the freestream pressure.

ALTITUDE ERRORS. —Altitude errors areintroduced by pressure lag, system leaks, andinstrument error. Lag can cause errors ofhundreds of feet, depending upon the rate ofpressure change and the system volume. In levelflight or during very low rates of ascent anddescent (500 to 600 ft/min), the errors are smalland almost negligible. Errors produced by systemleaks can be significant if the leak is in apressurized area of the aircraft.

INSTRUMENT ERRORS. —Instrumenterrors are the result of friction, hysteresis,temperature, acceleration, and design limitations.The overall accuracy of the instrument dependson the range of the instrument and complexity ofthe mechanism.

Equipment Requirements

Today, all naval aircraft have automatic alti-tude reporting. In most of the high-performance

aircraft containing an ADC, the ADC auto-matically reports altitude. The modification alsoprovides an output for the AAU-19/A servoedaltimeter. Slower, low-flying aircraft, notrequiring an ADC, have AAU-21A altimeter-encoders. The signal is based on the standardatmospheric pressure of 29.92 inches of mercury.

Air Data Requirements

The basic functions of the air data computerin the Air Traffic Control Radar Beacon System(ATCRBS) are to correct the sensed static pressureand the provision of appropriate outputs ofaltitude. The altitude outputs are usable in twoforms—a digital output in 100-foot incrementsfor wire transmission to transponders, andan analog synchro output for providing pressurealtitude information to an AAU-19/A servoedaltimeter. The system also includes a failuremonitoring circuit. Table 6-2 gives the generalcomputer requirements for the CPU-46/A and theCPU-66/A.

Air data computers consist of a pressure-sensing means, a computing mechanism, anerror signal generator, a servo loop, and outputdevices. The output devices provide signals forboth altitude indication and altitude reporting.The unit computes corrected pressure altitudefrom the inputs of indicated static pressure,total pressure, and information on the aircraftstatic-pressure error. All of this informationin the computer is on a two-dimensional cam. Thecam is readily replaceable if it should fail or ifyou must calibrate the unit for another type ofaircraft.

Look at figure 6-21 as you read about systemoperation. A small step change in static pressurecauses the aneroid capsule to deflect, rotating therocking shaft assembly. The rotation makes theGeneva sector and locking disk rotate themicrosyn rotor from its null position. Alternately,a correction for static-pressure defect can rotatethe microsyn rotor. In either case, the errorvoltage generated is amplified and directed to thetwo-phase servomotor. The servomotor drives the

Table 6-2.-Computer Requirements

Figure 6-21.-CPU-66/A computer components.

output devices, lead screw, and frame until themicrosyn is again in a null position. Thus, theangular rotation of the frame caused by theaneroid capsule displacement provides a changein the synchro and encoder outputs.

CPU-46/A SUPERSONIC ALTITUDECOMPUTER. —The CPU-46/A altitude com-puter (fig. 6-22) is used in high-performance,

Figure 6-22.-CPU-46/A supersonic altitude computer.

supersonic aircraft. These aircraft haveoperational limitations of 80,000 feet of altitudeand Mach 2.5. The CPU-46/A consists ofpressure sensors, a computing system, an elec-tronic package, and output devices. The unitreceives pressure from the aircraft pitot-staticsystem and requires a 115-volt, 400-hertz powersource. It provides altitude information, correctedfor static-pressure error, as follows:

• Two separate analog signals forpositioning the remote AAU-19/A altimeters.

• An encoded binary signal to the airbornetransponder for altitude reporting. The out-put of the encoder must agree with the out-put to the AAU-19/A to within ±20 feet. Amechanical cam compensates for static-pressure position error. Corrected pressure altitudeis computed and provided as output signals.

If the computer fails, the AAU-19/A altimeterautomatically reverts to a pneumatic standbymode, and the altitude reporting encoderdeactivates. A standby flagAAU-19/A display to advisefailure.

appears on thethe pilot of the

CPU-66/A SUBSONIC ALTITUDE COM-PUTER. —The CPU-66/A altitude computer(fig. 6-23) is used to altitudes of 50,000 feet andat speeds up to 0.95 Mach. Its operation is similarto that of the CPU-46/A.

The ADCs have been modified to have thesame altitude output capabilities as the CPU-46/Aor the CPU-66/A. The modification depends onthe speed and altitude capabilities of the aircraftin which the unit is installed.

Altimetry

The three altimeters that work with theautomatic altitude reporting system are theAAU-19/A, AAU-21/A, and AAU-24/A (fig.6-24).

SERVOED BAROMETER ALTIMETERAAU-19/A. —The counter-drum-pointer servoedbarometric altimeter (fig. 6-25) consists of a

Figure 6-23.-CPU-66/A subsonic altitude computer. pressure altimeter combined with an at-poweredservomechanism. The altitude display is in digitalform, using a 10,000-foot counter, a 1,000-footcounter, and a 100-foot drum. Also, a single

Figure 6-24.-AAU-19/A, AAU-21/A, AAU-24/A altimeters.

Figure 6-25.-AAU-19/A altimeter.

pointer shows hundreds of feet on a circular scale.The barometric pressure setting (baroset) knob isused to insert the local pressure in inches ofmercury. The baroset knob has no effect on thedigital output (mode C) of the ADC. This digitaloutput is always referenced to 29.92 inches ofmercury.

The altimeter has a servoed mode and apressure mode of operation. The mode ofoperation is controlled by a spring-loaded,self-centering mode switch, placarded RESET andSTBY. In the servoed mode, the altimeter displaysaltitude, corrected for position error, from thesynchro output of the air data computer. In thestandby mode, the altimeter operates as astandard altimeter. In this mode, it uses staticpressure from the static system that is uncorrectedfor position error.

The servoed mode is selected by placing themode switch to RESET for 3 seconds. The ac

power must be on. During standby operation, ared STBY flag appears on the dial face. Thealtimeter automatically switches to standbyoperation during an electrical power loss or whenthe altimeter or altitude computer fails. Thestandby operation is selected by placing the modeswitch to STBY. An ac-powered internal vibratorautomatically energizes in the standby mode tolessen friction in the display mechanism.

With the local barometric pressure set, thealtimeter should agree to ±75 feet of fieldelevation in both modes.

AAU-21/A ALTIMETER. —AAU-21/Aaltimeter is used in low/slow aircraft. It has acounter-drum-pointer display similar in appear-ance to the AAU-19/A. The altimeter containsa servo-driven encoder. The encoder provides analtitude signal to the aircraft transponder fortransmission to a ground station.

AAU-24/A ALTIMETER. —The AAU-24/Aaltimeter (fig. 6-26) contains a precision pressure-sensing device, counter, and pointer drivemechanisms. It also contains a combinationcounter drum and pointer for altitude display. Thecounter displays two digits, showing multiples of10,000 feet and 1,000 feet respectively, and movesintermittently. The drum shows multiples of 100feet and moves continuously. The pointer travelsone revolution for each increment of 1,000 feetof altitude. The pointer scale is from 0 to 9; eachstep representing an increment of 100 feet. Each100-foot step is split into two increments of 50feet each.

The barometric setting (baroset) knob islocated in the lower left corner of the bezel. Itprotrudes a maximum of 0.73 inch in front of thebezel. The baroset knob works with a four-digitcounter, designated IN Hg, to set the altitudeindication to the prevailing barometric pressure.It is adjustable from 28 to 31 inches of mercury.There is a locking screw next to the baroset knob.This screw is used only during calibrationprocedures to align the barometric pressure(IN Hg) indication with altitude indication.

Two sets of internal lights, one red and onewhite, provide dial lighting. Each set consists of

four lights. Controls for dial lighting are externalto the altimeter.

To overcome the effects of stop and jumpfriction in altimeter mechanisms, the altimeter hasan internal, electrically operated mechanicalvibrator.

The block diagram, shown in figure 6-27,shows altimeter operation. Outside air pressurefrom the aircraft static-pressure system goes tothe altimeter case. Two evacuated pressure-sensingelement assemblies within the case expandas the case pressure decreases with increasingaltitude. Expansion or contraction of eachelement sends movement to a rocker shaftassembly via a link-arm configuration. Pivotsfor the rocker shaft assemblies couple to atemperature compensating ring that changes themovement of the rocker shafts. This movementcompensates for temperature changes. A dif-ferential gear assembly, coupled by segment gearsto the rocker shafts, provides an output move-ment. This movement is proportional to theaverage expansion or contraction of the twosensing element assemblies.

The output gear of the differential gearassembly drives a gear train mechanism. This gear

Figure 6-26.-AAU-24/A altimeter dial face.

train operates a counter-drum assembly and apointer assembly that traverses the instrumentdial. The counter/drum provides a numericalreadout of altitude. The position of the pointerprovides greater accuracy of readout.

The baroset knob is coupled, via a barometricsetting shaft, to two spur gears that run on an idlershaft. One spur gear meshes with a gear that drivesthe barometric counter. The other spur gearmeshes with the gear train mechanism thatprovides altitude indication. Thus, the pressurethe barometric counter shows is tied to a particularindication of altitude.

To permit calibration of the pressure andaltitude relationship, a device can disengage thenormal barometric operation of the altimeter.Using a locking screw, you lock the device in thelocked position. In this position, both spur gearson the idler shaft are actuated by the barosetknob. When calibrating, slacken the locking screwand move it toward the edge of the bezel. Youcan now withdraw the baroset knob by a smallamount away from the bezel. This actiondisengages the knob from one spur gear soturning the knob now operates only the altitudeindication.

The large increases in the speed and volumeof air traffic caused the need for a more efficient,automated air traffic control system. TheATCRBS, which automatically provides thecontroller with a three-dimensional identifiedpresentation, is a step toward the solutionof the problems of vertical separation, terrainclearance, and collision avoidance. Automaticaltitude reporting relieves both the pilot andthe ground controller of much radio work.It also provides continuous monitoring of thealtitude of properly equipped aircraft. Thissystem results in increased efficiency in the airtraffic control system. The system also gives pilotsa more accurate altimetry system, reducingaltimetry errors of 0.8 percent of altitude to 0.35percent of altitude.

ANGLE-OF-ATTACK (AOA)INDICATING SYSTEM

The angle-of-attack indicating system detectsaircraft angle of attack from a point on theside of the fuselage. It furnishes referenceinformation for the control and actuationof other units and aircraft systems. It providessignals to operate an angle-of-attack indicator

Figure 6-27.-Altimeter block diagram.

(fig. 6-28) on the pilot’s instrument panel. Thisindicator displays a continuous visual indicationof the local angle of attack. A typical angle-of-attack system provides electrical signals foroperating the rudder pedal shaker. The shakerwarns the pilot of an impending stall when theaircraft is approaching the critical stall angle ofattack. Electrical switches in the AOA indicatoroperating at various preset angles of attackenergize colored lights in the approach lightsystem and an approach index light in the cockpit.

Figure 6-28.-AOA indicators: (A) radial; (B) vertical scale.

These lights furnish the landing signal officerand the pilot with an accurate indication ofapproach angle of attack during landing. Anangle-of-sideslip system, consisting of anairstream direction detector, angle-of-attack, andangle-of-sideslip compensator, is installed onsome aircraft. The outputs from these are usedfor controlled rocket firing.

The angle-of-attack indicating system consistsof an airstream direction detector transmitter(fig. 6-29) and an indicator. The airstreamdirection detector measures local airflow directionrelative to the true angle of attack. It does thisby determining the angular difference betweenlocal airflow and the fuselage reference plane. Thesensing element works with a servo-drivenbalanced bridge circuit, which converts probepositions into electrical signals.

The angle-of-attack indicating systemoperation is based on detection of differentialpressure at a point where the airstream is flowingin a direction that is not parallel to the true angleof attack oft he aircraft. This differential pressureis caused by changes in airflow around the probe.

Figure 6-29.-AOA transmitter.

The probe extends through the skin of the aircraftinto the airstream.

The exposed end of the probe contains twoparallel slots (ports). These slots detect thedifferential airflow pressure (fig. 6-30). Air fromthe slots passes through two separate air passagesto separate compartments in a paddle chamber.Any differential pressure, caused by misalignmentof the probe to the direction of airflow, causesthe paddles to rotate. The moving paddles rotatethe probe, through mechanical linkage, until thepressure differential is zero. This occurs when theslots are symmetrical with the airstream direction.

Two potentiometer wipers, rotating with theprobe, provide signals for remote indications.Probe position, or rotation, is converted into anelectrical signal by one of the potentiometers thatis the transmitter component of a self-balancingbridge circuit. When the angle of attack of theaircraft changes, the position of the transmitterpotentiometer alters. This causes an error voltageto exist between the transmitter potentiometer andthe receiver potentiometer in the indicator.

Current flows through a sensitive polarizedrelay to rotate a servomotor located in theindicator. The servomotor drives a receiverpotentiometer in the direction required to reducethe error voltage. This action restores the circuitto a null or electrically balanced condition. Thepolarity of the error voltage determines theresultant direction of rotation of the servomotor.The indicating pointer is attached to, and moveswith, the receiver potentiometer wiper arm toshow on the dial the relative angle of attack.

Adjustable cam switches (fig. 6-31) control theapproach light, approach indexer, and stallwarning. The cams (switches) operate as the

indicator responds to movement of the transmitterpotentiometer. Figure 6-32 shows the relationshipof the indicator to the lights and stall warning.The AOA indexer on the pilot’s glare shield hastwo arrows and a circle illuminated by coloredlamps to provide approach information. The cam-actuated switch contacts in the angie-of-attackindicator also control the angle-of-attack indexer.The upper arrow is for high angle of attack(green). The lower arrow is for low angle of attack(red). The circle is for optimum angle-of-attack(amber). An arrow and a circle together show anintermediate position.

The indexer lights function only when thelanding gear is down. A flasher unit causes theindexer lights to pulsate when the arresting hookis up with the HOOK BYPASS switch in theCARRIER position.

STALL WARNING SYSTEM

Many aircraft have stall warning indicators towarn the pilot of an impending aerodynamic stall.In the past, stall warning indicators were of apneumatic control type. These devices activatedeither warning horns or flashing lights. Later,research found that a stall relates directly to theangle of attack, regardless of airspeed, powersetting, or aircraft loading. The stall warningdevices of most aircraft now in the fleet operateat a specified angle of attack. The devices operatethrough cams in the angle-of-attack indicator. Thecam-driven switch activates a vibrator motorconnected to either a rudder pedal or the control

Figure 6-30.-Mechanical schematic of airstream directiondetector.

Figure 6-31.-Angle-of-attack indicator cams.

Figure 6-32.-Angle-of-attack (AOA) indications.

stick. Figure 6-33 shows a simplified schematicof the rudder shaker system. When the aircraftreaches stall angle of attack, the AOA indicatorcam-actuated switch completes the rudder shakermotor circuit to ground. When the angle of attackreturns below stall conditions, the cam deactuatesthe switch. The switch action removes the groundto the rudder shaker motor.

GYROSCOPIC INSTRUMENTS

Early aircraft were flown by visually aligningthe aircraft with the horizon. With poor visibility,it was not possible to fly the aircraft safely. Theneed for flight instruments to correct thiscondition lead to the development of fyroscopicinstruments. The gyroscopic properties of aspinning wheel made precision instrument flying,precise navigation, and pinpoint bombingpractical and reliable. Some of the instrumentsthat use this principle are the turn-and-bankindicator, directional gyro, gyro horizon, and thedrift meter. Systems that use the gyroscopicprinciple include the AFCS, gyrostabilizedflux-gate compass, and the inertial navigationsystem. The following paragraphs contain a briefreview of gyroscopic principles.

Figure 6-33.-Rudder shaker schematic (simplified).

NOTE: You should review Navy Electricityand Electronics Training Series (NEETS), module15, Principles of Synchros, Servos, and Gyros,before continuing.

A gyroscope is a spinning wheel or rotor withuniversal mounting. This mounting allows thegyroscope to assume any position in space. Anyspinning object exhibits gyroscopic properties.The wheel, with specific design and mounts touse these properties, is a gyroscope. The twoimportant design characteristics for instrumentgyros are

1. high density weight for small size, and2. high speeds rotation with low-friction.

The mountings of the gyro wheels are gimbals.They can be circular rings or rectangular frames.However, some flight instruments use part of theinstrument case itself as a gimbal. A simplegyroscope is shown in figure 6-34.

The two general types of mountings for gyrosare the free or universal mounting and therestricted or semirigid mounting. The type ofmounting the gyro uses depends on the gyro’spurpose.

A gyro can have different degrees of freedom.The degree of freedom depends on the numberof gimbals supporting the gyro and the arrange-ment of the gimbals. Do not confuse the termdegrees of freedom, as used here, with an angularvalue as in degrees of a circle. The term degrees

Figure 6-34.-Simple gyroscope.

of freedom, as used with gyros, shows the numberof directions in which the rotor is free to move.(Some authorities consider the spin of the rotoras one degree of freedom, but most do not.)

A gyro enclosed in one gimbal, such as the oneshown in figure 6-34, has only one degree offreedom. This is a freedom of movement back andforth at a right angle to the axis of spin. Whenthis gyro is mounted in an aircraft, with its spinaxis parallel to the direction of travel and capableof swinging from left to right, it has one degreeof freedom. The gyro has no other freedom ofmovement. Therefore, if the aircraft should noseup or down, the geometric plane containing thegyro spin axis would move exactly as the aircraftdoes in these directions. If the aircraft turns rightor left, the gyro would not change position, sinceit has a degree of freedom in these directions.

A gyro mounted in two gimbals normally hastwo degrees of freedom. Such a gyro can assumeand maintain any attitude in space. For illustrativepurposes, consider a rubber ball in a bucket ofwater. Even though the water is supporting theball, it does not restrict the ball’s attitude. Theball can lie with its spin axis pointed in anydirection. Such is the case with a two-degree-of-freedom gyro (often called a free gyro).

In a two-degree-of-freedom gyro, the basesurface turns around the outer gimbal axis oraround the inner gimbal axis, while the gyro spinaxis remains fixed. The gimbal system isolates therotor from the base rotation. The universallymounted gyro is an example of this type.Restricted or semirigid mounted gyros are thosemounted so one plane of freedom is fixed inrelation to the base.

Practical applications of the gyro are basedupon two basic properties of gyroscopic action:

1. Rigidity in space

2. Precession

Newton’s first law of motion states “A bodyat rest will remain at rest, or if in motion willcontinue in motion in a straight line, unless actedupon by an outside force.” An example of thislaw is the rotor in a universally mounted gyro.When the wheel is spinning, it stays in its originalplane of rotation regardless of how the basemoves. Figure 6-35 shows this principle. The

Figure 6-35.-Action of a freely mounted gyroscope.

gyroscope holds its position relative to space, eventhough the earth turns around once every 24hours.

The factors that determine how much rigiditya spinning wheel has are in Newton’s second lawof motion. This law states “The deflection of amoving body is directly proportional to thedeflective force applied and is inverselyproportional to its mass and speed.” To obtainas much rigidity as possible in the rotor, it hasgreat weight for size and rotates at high speeds.To keep the deflective force at a minimum, therotor shaft mounts in low friction bearings. Thebasic flight instruments that use the gyroscopicproperty of rigidity are the gyro horizon, thedirectional gyro, and any gyrostabilized compasssystem. Therefore, their rotors must be freely oruniversally mounted.

Precession (fig. 6-36) is the resultant action ordeflection of a spinning wheel when a deflectiveforce is applied to its rim. When a deflective forceis applied to the rim of a rotating wheel, theresultant force is 90 degrees ahead of the directionof rotation and in the direction of the appliedforce. The rate at which the wheel precesses isinversely proportional to rotor speed and directlyproportional to the deflective force. The forcewith which a wheel precesses is the same as thedeflective force applied (minus the friction in thegimbal ring, pivots, and bearings). If too greata deflective force is applied for the amount of

Figure 6-36.-Precession resulting from deflective force.

rigidity in the wheel, the wheel precesses andtopples over at the same time.

Any spinning mass exhibits the gyroscopicproperties of rigidity in space and precession. Therigidity of a spinning rotor is directly proportionalto the weight and speed of the rotor speed andinversely proportional to the deflective force.

Attitude Indicator

Pilots determine aircraft attitude by referringto the horizon when they can see it. Often,however, the horizon is not visible. When it isdark or when there are obstructions to visibilitysuch as overcast, smoke, or dust, the pilot cannotuse the earth’s horizon as a reference. When thiscondition exists, they refer to an instrument calledthe attitude indicator. This instrument is alsoknown as a vertical gyro indicator (VGI), artificialhorizon, or gyro horizon. From these instruments,pilots learn the relative position of the aircraftwith reference to the earth’s horizon.

The attitude indicator gyro rotor revolveswith its spin axis in a vertical position tothe earth’s surface. This vertical position isrigidly maintained as the aircraft pitches androlls about the space-rigid gyro (fig. 6-37).

The case of the gyro identically duplicatesaircraft movement. The case is free to revolvearound the stable gyro because of the mountingof the gyro rotor in gimbals. It follows, there-fore, that the aircraft itself actually revolvesaround the rotor and is the complementingfactor in establishing the indications of theinstrument.

Figure 6-37.-Aircraft reference and gyro stability.

Although attitude indicators (fig. 6-38)differ in size and appearance, they all havethe same basic components and present thesame basic information. There will alwaysbe a miniature aircraft on the face of theindicator that represents the nose (pitch) andwing (bank) attitude of the aircraft. The bankpointer on the indicator face shows the degree ofbank (in 10-degree increments up to 30 degrees,then in 30-degree increments to 90 degrees). Thesphere is always light on the upper half and darkon the lower half to show the difference betweensky and ground. Calibration marks on thesphere show degrees of pitch in 5- or 10-degreeincrements. Each indicator has a pitch trimadjustment for the pilot to center the horizon asnecessary.

Some attitude indicators have a self-containedgyro. Other more modern indicators use pitch androll information from the inertial system or theattitude heading reference system. These systemsare accurate and reliable. They gain their relia-bility and accuracy from being larger since sizeis not limited by the space of an instrument panel.Electrical signals from the remote gyro travel viasynchros. The signal is amplified in the indicatorto drive servomotors and position the indicatorsphere. This positioning is exactly as the verticalgyro position in the gyro case. In the newerattitude indicators, the sphere is gimbal-mountedand capable of 360-degree rotation. In contrast,the older gyros could only travel 60 degrees to 70degrees of pitch and 100 degrees to 110 degreesof roll. Some aircraft incorporate an all-attitude

Figure 6-38.-Roll and pitch indications on the attitude indicator.

indicator (fig. 6-39). In addition to pitch and roll,this indicator shows compass information alongthe horizon bar. It also shows turn-and-bankinformation on the bottom. An even moresophisticated instrument, the flight director,displays the above information plus radionavigation information, all on one instrument.

NOTE: Because of close association andinterface with inertial navigation systems, chapter7 of this TRAMAN contains information aboutattitude indicating systems.

Turn-and-Bank Indicator

The turn-and-bank indicator (fig. 6-40), alsocalled the turn-and-slip indicator, shows the lateralattitude of an aircraft in straight flight. It alsoprovides a reference for the proper executions ofa coordinated bank and turn. It shows when theaircraft is flying on a straight course and thedirection and rate of a turn. It was one of the firstmodern instruments for controlling an aircraftwithout visual reference to the ground or horizon.

Figure 6-40.-Turn-and-bank indicator.

The indicator is a combination of twoinstruments—a ball and a turn pointer. The ballpart of the instrument operates by natural forces(centrifugal and gravitational). The turn pointerdepends on the gyroscopic property of precession

Figure 6-39.-All-attitude indicator (AAI).

for its indications. The power for the turnindicator gyro is either electrical or vacuum.

BALL. —The ball portion of a turn-and-bankindicator (fig. 6-40) consists of a sealed, curved,glass tube. The tube contains water-white keroseneand a black or white agate or common steel ballbearing. The ball bearing is free to move insidethe tube. The fluid provides a damping action andensures smooth and easy movement of the ball.The curved tube allows the ball to seek the lowestpoint when in level flight. This is the tube center.A small projection on the left end of the tubecontains a bubble of air. The bubble lets the fluidexpand during changes in temperature. There aretwo markings or wires around the center of theglass tube. They serve as reference markers toshow the correct position of the ball in the tube.The plate that holds the tube and the referencesare painted with luminous paint.

The only force acting on the ball duringstraight flight (no turning) with the wings levelis gravity. The ball seeks its lowest point and stayswithin the reference marks. In a turn, centrifugalforce also acts on the ball in a horizontal planeopposite to the direction of the turn.

The ball assumes a position between thereference markers when the resultant of centri-fugal force and gravity acts directly opposite toa point midway between the reference markers.When the force acting on the ball becomeunbalanced, the ball moves away from the centerof the tube.

In a skid, the rate of turn is too great for theangle of bank. The excessive centrifugal forcemoves the ball to the outside of the turn. Theresultant of centrifugal force and gravity is notopposite the midpoint between the referencemarkers. The ball moves in the direction of theforce, toward the outside of the turn. Returningthe ball to center (coordinated turn) calls forincreasing bank or decreasing rate of turn, or acombination of both.

In a slip, the rate of turn is too slow for theangle of bank. The resultant of centrifugal forceand gravity moves the ball to the inside of theturn. Returning the ball to the center (coordinatedturn) requires decreasing the bank or increasingthe rate of turn, or a combination of both.

The ball instrument is actually a balanceindicator because it shows the relationshipbetween angle of bank and rate of turn. It letsthe pilot know when the aircraft has the correctrate of turn for its angle of bank.

TURN POINTER. —The turn pointeroperates on a gyro. The gimbal ring encircles thegyro in a horizontal plane and pivots fore and aftin the instrument case. The major parts of the turnportion of a turn-and-bank indicator are asfollows.

• A frame assembly used for assembling theinstrument.

• A motor assembly consisting basically ofthe stator, rotor, and motor bearings.The electrical motor serves as the gyro forthe turn indicator.

• A plate assembly for mounting theelectrical receptacle, pivot assembly, andchoke coil and capacitors for eliminat-ing radio interference. It also mounts thepower supply of transistorized indicators.

• A damping unit that absorbs vibrationsand prevents excessive oscillations of theneedle. The unit consists of a piston andcylinder mechanism. The adjustment screwcontrols the amount of damping.

• An indicating assembly composed of a dialand pointer.

• The cover assembly.

The carefully balanced gyro rotates about thelateral axis of the aircraft in a frame that pivotsabout the longitudinal axis. When mounted in thisway, the gyro responds only to motion arounda vertical axis. It is unaffected by rolling orpitching.

The turn indicator takes advantage of one ofthe basic principles of gyroscopes—precession.Precession, as already explained, is a gyroscopes’snatural reaction 90 degrees in the direction ofrotation from an applied force. It is visible asresistance of the spinning gyro to a change indirection when a force is applied. As a result,when the aircraft makes a turn, the gyro positionremains constant. However, the frame in whichthe gyro hangs dips to the side opposite thedirection of turn. Because of the design of thelinkage between the gyro frame and the pointer,the pointer shows the correct direction of turn.The pointer displacement is proportional to the. .aircraft rate of turn. If the pointer remains on

center, it shows the aircraft is flying straight. Ifit moves off center, it shows the aircraft is turningin the direction of the pointer deflection. The turnneedle shows the rate (number of degrees perminute) at which the aircraft is turning.

By using the turn-and-bank indicator, the pilotchecks for coordination and balance in straightflight and in turns. By cross-checking thisinstrument against the airspeed indicator, therelation between the aircraft lateral axis and thehorizon can be determined. For any givenairspeed, there is a definite angle of banknecessary to maintain a coordinated turn at agiven rate.

MISCELLANEOUS FLIGHTINSTRUMENTS

The pilot uses several other indicators tocontrol the aircraft. These indicators are notalways useful, but they come into play underspecial flight conditions. As you read this section,look at figure 6-41.

Accelerometer Indicators

The pilot must limit aircraft maneuvers sovarious combinations of acceleration, airspeed,gross weight, and altitude remain within specified

Figure 6-41.-Miscellaneous flight instruments.

values. This cuts out the possibility of damagingaircraft as a result of excessive stresses. Theaccelerometer shows the load on the aircraftstructure in terms of gravitation (g) units. Itpresents information that lets the aircraft bemaneuvered within its operational limits.

The forces sensed by the accelerometer actalong the vertical axis of the aircraft. The mainhand moves clockwise as the aircraft acceleratesupward and counterclockwise as the aircraftaccelerates downward.

The accelerometer indications are in g units.The main indicating hand turns to +1 g when the

lift of the aircraft wing equals the weight of theaircraft. Such a condition prevails in level flight.The hand turns to +3 g when the lift is three timesthe weight. The hand turns to minus readingswhen the forces acting on the aircraft surfacescause the aircraft to accelerate downward.

The accelerometer operates independently ofall other aircraft instruments and installations.The activating element of the mechanism is a massthat is movable in a vertical direction on a pairof shafts (fig. 6-42). A spiral-wound main springdampens the vertical movement of the mass. Theforce of the mass travels by a string-and-pulley

Figure 6-42.-Accelerometer mechanical schematic.

system to the main spring and main shaft. Fromhere, it goes to the plus and minus assemblies. Thehand assemblies mount on the plus and minusassemblies. Changes in vertical acceleration causemovement of the mass on the shafts, whichtranslates into a turning motion of the main shaft.The turning motion pivots the indicating handsaround the dial. The hand travels a distanceequivalent to the value, in g units, of the upwardor downward acceleration of the aircraft.

The accelerometer operates on the principleof Newton’s Law of Motion. During level flight,no forces act to displace the mass from a positionmidway from the top and bottom of the shafts.Therefore, the accelerometer pulley systemperforms no work, and the indicating handsremain stationary at +1 g. When the aircraftchanges from level flight, forces act on the mass.This action causes the mass to move either aboveor below its midway position. These movementscause the accelerometer indicating hands tochange position. When the aircraft goes nosedown, the hands move to the minus section of thedial. When the nose goes up, they move to theplus section.

The main hand continuously shows changesin loading. The two other hands on theaccelerometer show the highest plus accelerationand highest minus acceleration of the aircraftduring any maneuver. The indicator uses a ratchetmechanism to maintain these readings. A knobin the lower left of the instrument face is used toreset the maximum- and minimum-reading handsto normal. Thus, the accelerometer keeps anindication of the highest accelerations during aparticular flight phase or during a series of flights.

Clocks

The standard Navy clock is a 12-hour, elapsed-time, stem-wound clock with an 8-day movement.This type of clock is in the cockpit for use by thepilot or copilot. Clocks maybe located elsewherefor use by other crew members as well. The pull-to-set winding stem is at the lower left of the dial.The dial has 60 divisions, which you read asminutes or seconds, as appropriate. The face hasstandard minute and hour hands, a sweep-secondhand, and an elapsed-time minute hand. You maystart, stop, or reset the elapsed-time minute handby pressing a single button at the upper right ofthe dial.

Direct-Reading Magnetic Compasses

During the early days of aviation, directionof flight was determined chiefly by direct-readingmagnetic compasses. Today, the direct-readingmagnetic compass (fig. 6-41) is used as a standbycompass. Direct-reading magnetic compasses usedin Navy aircraft mount on the instrument panelfor use by the pilot. It is read like the dial of agauge.

A nonmagnetic metal bowl, filled with liquid,contains the compass indicating card. Thecard provides the means of reading compassindications. The card mounts on a float assemblyand is actually a disk with numbers painted onits edge. A set of small magnetized bars or needlesfasten to this card. The card-magnet assembly sitson a jeweled pivot, which lets the magnets alignthemselves freely with the north-south componentof the earth’s magnetic field. The compass cardand a fixed-position reference marker (lubber’sline) are visible through a glass window on theside of the bowl.

An expansion chamber in the compass pro-vides for expansion and contraction of the liquidcaused by altitude and temperature changes. Theliquid dampens, or slows down, the oscillation ofthe card. Aircraft vibration and changes inheading cause oscillation. If suspended in air, thecard would keep swinging back and forth and bedifficult to read. The liquid also buoys up the floatassembly, reducing the weight and friction on thepivot bearing.

Instrument-panel compasses for naval aircraftare available with cards marked in steps of either2 degrees or 5 degrees. Such a compass indicatescontinuously without electrical or informationinputs. You can read the aircraft heading bylooking at the card in reference to the lubber linethrough the bowl window.

Standby Attitude Indicator

The standby attitude indicator (fig. 6-41 onthe pilot instrument panel) consists of a miniatureaircraft symbol, a bank angle dial, and a bankindex. It also includes a two-colored drumbackground with a horizon line dividing the two.

The indicator roll index is graduated in10-degree increments to 30 degrees, with

graduation marks at 60 degrees and 90 degrees. Q5. What instrument uses both pitot and staticThe indicator is capable of displaying 360 degreesof roll, 92 degrees of climb, and 79 degrees ofdive. Because of the high spin rate of the gyro,the indicator displays accurate pitch-and-roll datafor 9 minutes after electrical failure. The attitudeindicator incorporates a pitch trim knob toposition the miniature aircraft symbol above orbelow the horizon reference line. The pitch trimknob also cages the gyro. When the pitch trimknob is pulled out, the gyro will cage. Rotatingthe knob clockwise while extended will lock in theextended position. The attitude indicator alsoincorporates an OFF flag. The flag appearsif electrical power fails, or if you cage thegyro.

Outside Air Temperature Indicator

An indicator displaying uncorrected outsideair temperature is located on the pilot’s instrumentpanel (fig. 6-41). A temperature-sensitive resistor(temperature bulb) is exposed to the slipstream.This resistor measures changes in temperature.The temperature of the air measurement is in theform of changing resistance. The outside airtemperature indicator displays this change inresistance. The graduated indicator dial is markedin Celsius, from +50 degrees to –50 degrees.

air pressure?

Q6. What is the Mach number for an aircraftflying at twice the speed of sound?

Q7. Define the terms “AGL” and “MSL.”

Q8. Define the term “absolute altitude.”

Q9. What is the name for the aneroid mech-anism used in most altimeters?

Q10. Name the four inputs to the air datacomputer system.

Q11. Define the term “impact pressure.”

Q1. List the two ways of grouping aircraftinstruments.

Q12. What is the operational limit of aircraftusing the CPU-46/A supersonic altitudecomputer?

Q2. What determines the air pressure at anyQ13. What system actuates the rudder pedal

given altitude?shaker to warn the pilot of an impendingstall?

Q3. What is the standard atmospheric pressure Q14. What determines the degrees of freedom forat 5,000 feet of altitude? a gyro?

Q4. Name the three indicators that use pressures Q15. Name the two basic properties of gyro-from the pitot-static system. scopic action.

Q16. What indicator shows the lateral attitudeof an aircraft?

ENGINE INSTRUMENT SYSTEMS

Learning Objective: Recognize the operat-ing principles and characteristics of variousengine instrument systems, includingtachometer, temperature indicating, fuelflow, oil pressure, fuel pressure, oiltemperature, exhaust-nozzle indicating,and torquemeter systems.

Engine instruments provide indications of tailpipe temperature, oil and fuel pressure, engineRPM, oil temperature, and fuel flow rate. Thepilot must be aware of engine operation at alltimes. If oil pressure falls below the normaloperating limit or tail pipe temperature becomesexcessively high, the engine instruments providethese indications to the pilot.

TACHOMETER SYSTEMS

The tachometer indicator is an instrument thatshows the speed of a gas turbine engine (jet) mainrotor assembly. Figure 6-43 shows tachometerindicators for various types of engines. The dialsof tachometer indicators used with jet engines areshown in percentage of revolutions per minute(RPM), based on takeoff RPM.

Several types and sizes of generators andindicators are used in the tachometer systems ofnaval aircraft. As a rule, they all operate on thesame basic principle. This section introduces youto information on tachometer systems. A typicalgenerator and a typical indicator are describedbecause it is not practical to describe all thegenerators and indicators. For detailed informa-tion on a particular system, you should refer tothe manufacturer’s manuals.

Essentially, the tachometer system consists ofan ac generator coupled to the aircraft engine andan indicator consisting of a magnetic-drag elementon the instrument panel. The generator transmitselectric power to a synchronous motor, a part ofthe indicator. The frequency of this power isproportional to the engine speed. An accurateindication of engine speed is obtained by applyingthe magnetic-drag principle to the indicating

Figure 6-43.-Tachometer indicators: (A) jet engine (radial);(B) jet engine (vertical scale).

element. The problem of changes in generatoroutput voltage is cut out by the generator andsynchronous-motor combination. These unitsmake a frequency-sensitive system for sending anindication of engine speed to the indicator withabsolute accuracy.

For many installations, it is desirable to senda single engine-speed indication to two differentstations in the aircraft. The frequency-sensitivesystem is ideal for this application because thereis no change in indication when a second indicatorconnects in parallel with the first. Synchronous-motor operation in each indicator depends onlyon the availability of enough power in thegenerator to operate both indicator motors.

Tachometer Generator

Tachometer generator units are small andcompact (about 4 inches by 6 inches). Thegenerator is constructed with an end shielddesigned so the generator can attach to a flat plateon the engine frame or reduction gearbox, withfour bolts.

Figure 6-44 shows a cutaway view of atachometer generator. You should refer to it whileyou read this section. The generator consistsessentially of a permanent-magnet rotor (callout1) and a stator (callout 8) that develop three-phasepower as the rotor turns.

The armature of the generator consists of amagnetized rotor. The rotor is cast directly ontothe generator shaft. The generator may be ofeither two- or four-pole construction. The two-and four-pole rotors are identical in appearanceand construction. They differ in that the two-polerotor is magnetized north and south diametricallyacross the rotor, while the four-pole rotor ismagnetized alternately north and south at eachof the four pole faces.

The key (callout 2) that drives the rotor is along, slender shaft. It has enough flexibility toprevent failure under the torsional oscillations

originating in the aircraft drive shaft. It will alsoaccommodate small misalignments between thegenerator and its mounting surfaces. This key goesinto the hollow rotor shaft. A pin (callout 3) atthe end opposite the drive shaft secures the keyin place. An oil-seal ring (callout 4) is locatedinside the hollow shaft and over this key. This sealprevents oil from leaking into the generatorthrough the hollow shaft. The shaft runs in twoball bearings (callout 5) set in stainless steel inserts.The inserts are cast directly into the generator endshields (callout 6). An adjusting spring (callout7) at the receptacle end of the shaft maintains theproper amount of end play.

The stator consists of a steel ring with alaminated core of ferromagnetic material. Athree-phase winding goes around this core and isinsulated from it. The winding is adapted fortwo- or four-pole construction, depending on the

Figure 6-44.-Cutaway view of a tachometer generator.

generator in which it is used. The two end shieldsare of die-cast aluminum alloy. They serve tosupport the generator stator and rotor by meansof a receptacle (callout 9). The receptacle attachesto the junction box (callout 10) of the generator.

Tachometer Indicators

Tachometer indicators mount on the cockpitinstrument panel. They are relatively small in size.The type of unit varies. Depending on theparticular installation, some are single elementand others are dual element, The operatingprinciples of the two types are basically the same.

Figure 6-45 shows a cutaway view of a single-element tachometer (radial) indicator. The unitconsists essentially of two parts a synchronousmotor and an indicating element. The motor runsin synchronism with the tachometer generator. Italso drives the indicating element through amagnetic-drag coupling. The indicating elementindicates the speed of the synchronous motor,and, therefore, the speed of the aircraft engine.

The synchronous motor (callout 3) consists ofa three-phase stator winding that goes in, and isinsulated from, a laminated circular core. Withinthe circular core is a shaft. The rotating partsattach to this shaft. A cotter pin secures ahysteresis disk (callout 1) to the shaft. Apermanent magnet rotor (callout 2) is free to moveon the shaft. The hysteresis disk at one shaft endand a spring at the other restrain longitudinal

motion of the permanent magnet rotor. Thespring is secured to the shaft to transmit torquefrom the rotor to the shaft. Ball bearings in themotor end shields support the shaft. These endshields also serve to locate the stator. This lets allparts of the motor maintain their proper positionwith respect to each other.

The armature of the synchronous motorconsists mainly of the permanent magnet and thehysteresis disk. The purpose of the permanent-magnet material is to provide starting and runningtorque at low speeds. The hysteresis disk providesstarting torque at high speed. This is necessarybecause the magnitude of flux is great, but thepermanent magnet, by itself, cannot pull into step.At the higher speeds, the hysteresis disk movesthe rotor up to near synchronism, and then thepermanent magnet pulls it into exact synchronism.

One end of the motor shaft extends throughthe front end shield and supports the drag-magnetassembly (callout 9). The drag-magnet assembly,which is driven by the synchronous motor,consists of two plates to which small permanentmagnets attach. The arrangement of the magnetsconcentrate the flux near the outside edge of thedrag disk. This arrangement obtains maximumtorque with minimum weight. Between the two,plates carrying the magnets is a disk (callout 4)of conducting material. This material is an alloyhaving a low-temperature coefficient. Thisprevents temperature changes from affecting thematerial’s resistance. The magnet assembly

Figure 6-45.-Cutaway view of a tachometer indicator (radial).

spinning around the disk of conducting materialproduces torque on the disk.

The drag disk connects to the lower end of theindicator assembly shaft. When the disk rotates,the indicator pointer moves to show the speed ofthe aircraft engine. The indicating element issupported by three posts. These post haveadjustable nuts (callout 8) for leveling theassembly as necessary. You can obtain furtherpositioning by the adjusting arm (callout 7).

The scale plate (callout 5) is calibrated eitherin RPM or percentage and shows engine speed.The cover assembly (callout 6) serves as aprotective container for the mechanism. Electricalconnection to the tachometer generator is by thereceptacle (callout 10) at the rear of the indicator.

Dual Indicators

With the increasing requirement for moreinstruments for efficient flight, the combinationof several instruments in one has become verycommon. The dual tachometer is an example ofa combination of instruments. Some multi-engineaircraft use dual tachometers.

The dual tachometer consists of twosynchronous-motor magnetic-drag tachometerindicators housed in a single case. The indicatorsshow the speed of rotation of the enginessimultaneously on a single dial. There is onetachometer indicator for each pair of engines onthe aircraft.

Vertical Scale Indicators

Vertical scale indicators are used on somemodels of naval aircraft. A vertical scale showsengine performance data such as fuel flow, enginespeed, exhaust gas temperature, and accelerometerreadings. Vertical scale indicators are compact,light in weight, and easily read.

Basically, all vertical scale indicators consistof a vertical tape that operates by an amplifier,motor, gears, and sprockets. These systems (fig.6-46) are basically the same as systems used onother aircraft. There may be possible changes innomenclature due to the various aircraft andengine manufacturers.

The engine indicating groups consist of cockpitindicators and associated sensing devices requiredto monitor left and right engine performance.Dual indicators display percentage of engine N2

rotor speed (RPM indicator), turbine inlettemperature (TIT indicator), and engine fuel flow(FF indicator). Individual indicators display

engine power trim (PT indicator) and N 1

overspeed caution (L or R N1OVSP cautionindicator light),

TACHOMETER INDICATOR. —The elec-trical tachometer (RPM) indicator displayspercentage of engine N2 rotor speed on twovertical scales. One each for the left and rightengines. The indicator scales are linear from 0 to6, and from 6 to 11 multiplied by 10 to get percentof RPM. Upper left and right limit range markersand OFF failure flags appear between the 10.4 and11 points of the scales. The absence of the OFFfailure flags confirms the indicator channels arereceiving power. The indicator receives variable-frequency signals proportional to N2 compressorspeed from each engine tachometer generator. Thesignals go to solid-state circuitry to produce aproportional drive signal for a servomotor. Theservomotor drives gears and sprockets to positiona tape on the indicator face, showing percentageof engine N2 rotational speed.

The test selector switch on the MASTERTEST panel is used to test the indicators. Itdisconnects a self-test circuit in the indicators fromthe tachometer-generator input circuits. Theself-test circuit then connects with the appropriatetest signal within the two indicator channels. Thechannel circuitry processes the test signal anddrives the indicator tapes to show 80 percent ofRPM.

N1 , N2 TACHOMETER GENERATORS. —The N1 and N2 tachometer generators are identicaltwo-phase alternator types. They supply electricalsignals directly proportional to engine-compressorrotational speed. The N1 tachometer generator forthe left and right engines is driven by the N2

low-pressure compressor rotor. An output fromthis generator goes to the engine rotor overspeeddetector. The signals represent N1 compressorrevolutions per minute (RPM). The N2

tachometer generator is driven by the N2 high-pressure compressor through the engine accessorygearbox. Signals from this generator go directlyto the electrical tachometer indicator (RPMindicator) at the pilot station. The indicatordisplays percentage of N2 rotor speed.

ENGINE ROTOR OVERSPEED DETEC-TOR. —The engine rotor overspeed detectorreceives electrical signals from the left and rightN1 tachometer generator. The detector consistsof three solid-state circuit boards—a crystaloscillator (clock) board and two identical gated

counter boards, one for each engine, The detector each counter. The counters measure the periodreceives a 0- to 33-volt, 0- to 70-hertz signal from for their respective tachometer-generator signal.each N1 tachometer generator. In the detector, Then, they compare the tachometer-generatorthe voltage is cut to a constant 5 volts. The normal signal period with the established clock referencefrequency signal is reshaped and processed period. If the incoming generator signal periodthrough the counter circuit to establish a signal is shorter than the clock reference period, theperiod. The clock sets up a reference period for counter energizes a switching relay. The relay

Figure 6-46.-F-14 engine instrument indicating group.

completes a circuit to power its respectivecaution indicator light on the pilot CAUTIONADVISORY indicator.

TEMPERATURE INDICATINGSYSTEMS

To properly monitor the operation of anaircraft engine, you must know various tempera-ture indications. Some of the more importantindications include the temperatures of the engineoil, free air, and exhaust systems of jet engines.Various types of thermometers, such as thebimetal and resistance types, collect and presentthis information.

The main parts of resistance thermometers arethe indicating instrument, the temperature-sensitive element (resistance bulb), and theconnecting wires leading from the bulb.

The indicator dials of resistance thermometersare calibrated according to the range of tempera-ture to be measured. Figure 6-47 shows aresistance thermometer indicator. The indicatorsare self-compensating, allowing for the changesin cockpit temperatures.

Wheatstone Bridge System

A schematic diagram of a Whetstone bridgethermometer circuit is shown in figure 6-48. Youshould refer to it as you read this section. Theresistance bulb element is one side of theWheatstone bridge circuit. The other three sidesare resistors in the indicating meter. The circuitreceives voltage from the aircraft dc power supply.

When the temperature bulb senses a tempera-ture of 0°C, its resistance is 100 ohms. The

Figure 6-47.-Indicator of a resistance thermometer.

Figure 6-48.-Wheatstone bridge thermometer.

resistance of arms X, Y, and Z are also 100 ohmseach. At this temperature the Wheatstone bridgeis in balance. This means the sum resistance ofX and Y equals the sum resistance of the bulb andZ. Therefore, the same amount of current flowsin both sides of this parallel circuit. Sinceall four sides are equal in resistance, thevoltage drop across side X equals the drop acrossthe bulb. Since these voltages are equal, thevoltage from A to B is zero, and the indicatorreads zero.

NOTE: If there is an open in the voltagesupply circuit, the galvanometers will also readzero.

When the temperature of the bulb increases,its resistance also increases. This unbalances thebridge circuit causing the needle to deflectto the right. When the temperature of the bulbdecreases, its resistance decreases. Again thebridge circuit goes out of balance. However, thistime the needle swings to the left.

The galvanometers is calibrated so the amountof deflection causes the needle to point tothe number of the meter scale. This numbercorresponds to the temperature at the location ofthe resistance bulb.

This instrument requires a constant and steadysupply of dc voltage. Fluctuations in the powersupply affect total bridge current, which can causean unbalanced bridge. Unless excessive heatdamages the bulb, it will give accurate serviceindefinitely. When a thermometer does notoperate properly, check carefully for loose wiringconnections before replacing the bulb.

Radiometer System

The radiometer is a temperature indicator thatuses two coils in a balanced circuit. In someinstruments, the coils turn between the poles ofa permanent magnet. In other instruments, a smallpermanent-magnet rotor turns between stationarycoils. Radiometer circuits vary in design, but theprinciple of operation is very much the same forall.

Figure 6-49 shows a simplified circuit with apermanent-magnet rotor. The two coils arestationary in the instrument, and the indicatorneedle fastens to the permanent-magnet rotor.The needle position is determined by how thepermanent magnet aligns itself with the resultantflux of the two coils.

For an understanding of how the circuitoperates, let’s trace the current through the circuit.Starting at ground, current flows up through thebulb, centering potentiometer R5 and R6, to pointD. Current through the left leg of the bridge isfrom ground through R1 to point A. Current thengoes from point A through the lower part of theexpansion and contraction potentiometer R2. Italso goes from pin 2 of R2 through R4 to point

Figure 6-49.-Radiometer type temperature indicator.

D. Here, the currents of the two legs combine andflow through R7 to the positive 28 volts.

Note that restoring coil L2, resistor R3, andupper part of potentiometer R2 forms a parallelpath for current flow from point A to pin 2 ofR2. Deflection coil L1 connects between pointsA and B. Therefore, any difference in potentialbetween these two points will cause current to flowthrough L1.

The radiometer temperature indicator uses afixed permanent magnet to pull the pointer to anoff position when the indicator is not operating.Thus, current through restoring coil L2 mustcompensate for the pulloff magnet when theindicator is operating. Variations in the resistanceof the bulb, because of temperature changes,cause a change in voltage at point B. It also causesthe resulting change in current through deflectioncoil L1.

Thermocouple System

Thermocouple temperature indicators showthe air temperatures in the heater duct ofanti-icing systems and in the exhaust systems ofjet engines.

A thermocouple is a junction or connectionof two unlike metals; such a circuit has twojunctions. When one of the junctions becomeshotter than the other, an electromotive forceis produced in the circuit. By including agalvanometers in the circuit, this electromotiveforce can be measured. The hotter the high-temperature junction (hot junction) becomes, thegreater the electromotive force. By calibrating thegalvanometers dial, in degrees of temperature,it becomes a thermometer. The galvanometerscontains the cold junction.

The thermocouple thermometer systems usedin naval aircraft consist of a galvanometersindicator, a thermocouple or thermocouples, andthermocouple leads. Some thermocouples consistof a strip of copper and a strip of constantanpressed tightly together. Constantan is an alloyof copper and nickel. Other thermocouples consistof a strip of iron and a strip of constantan. Othersmay consist of a strip of Chromel and a strip ofAlumel.

The hot junction of the thermocouple variesin shape, depending on its application. Twocommon types—gasket and rivet—are shown in

figure 6-50. In the gasket thermocouple, the ringsof two dissimilar metals are pressed together,forming a spark plug gasket. Each lead thatconnects back to the galvanometers must be of thesame metal as the thermocouple part to which itconnects. For example, a copper wire connects tothe copper ring, and a constantan wire connectsto the constantan ring. Thermocouple leads arecritical in makeup and length because thegalvanometers are calibrated for a specific set ofleads in the circuits.

TURBINE INLET TEMPERATURE (TIT)INDICATOR SYSTEM. —Some aircraft have anengine turbine inlet temperature indicator system(fig. 6-51) to provide a visual indication oftemperatures entering the turbine. The tempera-ture of each engine turbine inlet is measured by18 dual-unit thermocouples in the turbine inletcasing. These dual thermocouples are connectedin parallel. One set sends signals through a harnessand aircraft wiring to an indicator. The otherset of thermocouples provides signals to thetemperature datum control. Each circuit iselectrically independent and provides dual systemdependability.

All parts of the engine temperature measure-ment system are made of Chromel and Alumelmaterial, including welds. Special wiring andwire identification are in the aircraft from

Figure 6-50.-Thermocouples: (A) gasket type; (B) rivet type.

the thermocouple harness terminal block to theindicator. Plugs in the thermocouple circuits arealso of a special type. The thermocouple harnessmounts on the turbine unit aft of the thermo-couple. The harness includes separate leads foreach of the 18 thermocouples, and it maintainstwo electrically separate circuits. The harness islocated inside a rigid metal, channel type ofhousing and cover. The leads and terminalsproject through holes in the front side of thehousing wall. Electrical signals from the 18dual-junction thermocouples are averaged withinthe harness.

The thermocouple assemblies mount on padsprovided around the turbine inlet case. Eachthermocouple incorporates two electricallyindependent junctions within a sampling typeprobe. AL identifies Alumel terminal studs, andCR identifies Chromel terminal studs.

Since the average voltage of the thermocouplesat the thermocouple terminal blocks represents theturbine inlet temperature, it is necessary that nointerference with the signal take place while thesignal goes to the indicator. Therefore, the wiringfrom the thermocouple terminal block to theindicator goes through the harness. The harnesswiring goes separately from other interference-producing wiring.

The indicator contains a bridge circuit withcold junction compensation, a two-phase motorto drive the pointer, and a feedback potenti-ometer. Also included in the indicator is the Zenervoltage reference circuit, a chopper circuit, anamplifier, power supply, a power-off flag, andan overtemp warning light.

Output of the bridge circuit goes to thechopper circuit, so the bridge circuit isn’t loaded.The chopper output goes to the amplifier. Outputof the amplifier feeds the variable field ofa two-phase motor. This field positions theindicator main pointer and the digital indicator,The motor also drives the feedback potentiometerto provide a nulling signal. The signal is relativeto the temperature signal and stops the drivemotor upon reaching the correct pointer position.

The Zener diode circuit protides a closelyregulated reference voltage in the bridge, Thissignal avoids the error caused by voltage variationfrom the indicator power supply. The indicatorpower supply powers the Zener circuit, thechopper, and the amplifier. It also powers thepower off warning flag and the fixed field of thetwo-phase motor.

The overtemperature warning light in theindicator comes on when the turbine inlet

Figure 6-51.-Turbine inlet temperature indicator system.

temperature (TIT) reaches 1,082°C. At this point,a switch in the indicator closes to energize thewarning light. One test switch installed externalto the indicators lets the crew test all the indicatorovertemperature warning lights at once. The testswitch simulates an overtemperature signal in eachindicator’s temperature control bridge circuit.

When power to an indicator fails, a redwarning flag becomes visible. Also, the indicatorpointers maintain their position, and the over-temperature warning light becomes inoperative.

The indicator scale is calibrated in degreesCentigrade from 0 to 12 (times 100°C). The digitalindicator goes from 0°C to 1,200°C in 2-degreeincrements.

The F-14 aircraft also used the thermocoupleprinciple for indicating engine turbine inlettemperatures. Each engine has 10 thermocoupleprobes, distributed at three stations on the engine.They measure and average engine turbine inlettemperature. There are three types of thermo-couple probes—compressor inlet temperature(Tt2), compressor discharge temperature (T~4),and exhaust gas temperature thermocouplepressure (Tt7–Pt7). Each thermocouple probe hasone or more Alumel-Chromel junctions. When

the junctions become hot, a reaction between thedissimilar metal generates a dc voltage. Athermocouple harness connects the thermocouplesin parallel to provide an average heat signal fromeach station.

The thermocouple temperature indicator (fig.6-46) displays turbine inlet gas temperature on twovertical scales, one for each engine. The scales arelinear from 0 to 6; segmented in tens, from 6 to14 and multiplied by 100°C when read. OFFfailure flags appear at the upper left and right ofthe indicator to show loss of signal input orelectrical power.

Internally, the indicator has two channels, onefor each engine. The channels basically consistof a cold junction compensator, rebalancepotentiometer, chopper, servo amplifier, servo-motor, and gear train. Thermocouple signalvoltage from the engines goes to the cold junctioncompensator in each channel. The compensatorprovides corrective voltages to counteract theeffect of secondary thermocouple junctionsin the indicator when Alumel and Chromelleads connect to copper ones. A stable voltagegoes to the old junction compensator andrebalance potentiometer.

The feedback of the potentiometer and outputof the compensator go to the chopper, where itcompares the inputs. The chopper provides a400-hertz error signal to the servo system. Thechopper output (signals relative to temperaturechange and potentiometer versus compensatordifference) goes to a servo amplifier. The servoamplifier modifies the signals to drive theservomotor. The shaft of the motor couples tothe rebalance potentiometer and indicator tapethrough the gear train.

As the amplifier error drives the motor, therebalance potentiometer goes in a direction thatreduces the error signal, nullifying the condition.The tape shows temperature, on the front scaleof the indicator, relative to thermocouple output.When supplied with 28 volts dc, a test circuit inthe indicator energizes a relay, disconnecting thethermocouple input. Then, it substitutes a testsignal of specific value to be processed and todrive the indicator tape.

EXHAUST GAS TEMPERATURE INDI-CATING SYSTEM. —The exhaust gas tempera-ture (EGT) indicating systems provide a visualtemperature indication in the cockpit of the engineexhaust gases. The following is a discussion of atypical exhaust gas temperature indicating system.

The aircraft contains two separate butidentical exhaust gas temperature indicatingsystems (fig. 6-52), one for each engine. Eachsystem has 12 dual thermocouples, a combinationindicator and transistorized amplifier, and the

interconnecting Chromel and Alumel leads.Power for the indicator-amplifier is from theessential 115-volt ac bus.

Both exhaust gas temperature indicators areon the pilot’s main instrument panel. Theyprovide a visual indication of the engine exhausttemperatures. Each instrument is a hermeticallysealed unit with a single receptacle for a matingplug electrical connection. The instrument scaleranges from 0°C to 1,200°C. There is a vernierdial in the upper right corner of the instrumentface. A power-off warning flag is in the lowerportion of the dial.

Internally, the indicator contains a simulatedthermocouple cold junction with compensatingresistors, a reference voltage source, and adc-to-ac modulator. It also contains a transistorpower output stage, miniature ac servomotor, andthe power-off warning flag. The temperatureindicator contains range markings on theinstrument face.

The thermocouples convert engine exhaust gastemperature into millivolts. The voltage from thethermocouples go directly to the indicator-amplifier through the Chromel and Alumel leads.The voltage is amplified and drives a smallservomotor. The motor, in turn, drives theindicator pointer. The thermocouple harnessconsists of two halves, each containing sixdual-loop thermocouples. The assembled halvesmake up two independent thermocouple systems,each consisting of 12 thermocouples connected in

Figure 6-52.-Exhaust gas temperature indicating system.

parallel. The harness mounts on the turbine frameaft of the turbine rotor.

FUEL FLOW SYSTEMS

Fuel flow indicating systems provide a con-tinuous indication of the rate of fuel delivery tothe engine. The rate of flow is in pounds per hour.In some systems, the indicator also shows theamount of fuel remaining in the tanks. A typicalflowmeter consists of two units—a transmitterand an indicator. The measurements are trans-mitted electrically to the panel-mounted indicator.Thus, use of electrical transmission ends the needfor a direct fuel-filled line from the engine to theinstrument panel. This minimizes the chance offire and reduces mechanical failure rate.

The fuel flowmeter system is quite similarto other synchro systems discussed in NavyElectricity and Electronics Training Series(NEETS) course. The following discussiondescribes a typical fuel flow indicating systemto acquaint you with flowmeters in general.However, you should always refer to the manualsfor the particular system you are maintaining.

Fuel Flow Transmitter

Figure 6-53 shows a cutaway view of a fuelflow transmitter. It is a two-in-one unit—afuel-measuring mechanism (or meter) and asynchro transmitter. You can separate these partsfrom one another for maintenance purposes, butthey join as a single assembly for installation.

The fuel enters the inlet port of the transmitterand flows against the vane (callout 1), causing thevane to swing. The spiral fuel chamber designallows the distance between the vane and chamberwall to become increasingly larger as fuel flowincreases. A calibrated hairspring (callout 2)retards the motion of the vane. The vane ceasesmotion when the forces exerted on it by thehairspring and by the fuel are equal.

The rotor shaft of the synchro transmitter(callout 3) connects to a bar magnet (callout 4).Attached to the vane shaft is a ring magnet(callout 5). The ring magnet moves as the vaneshaft moves. The transmitter mounting frame isbetween the bar magnet and the ring magnet,forming a liquid-tight seal. This is the seal betweenthe fuel-metering section of the mechanism andthe synchro. However, the bar magnet moves inunison with the ring magnet because the twomagnets are magnetically coupled. The south poleof the ring magnet is opposite the north pole of

Figure 6-53.-Cutaway view of a fuel flow transmitter.

the bar magnet. The two magnets send vanemovement, caused by the fuel flow, to the synchrorotor. This action results in a correspondingmovement of the rotor. Therefore, the angulardisplacement of the vane in relation to the fuelchamber housing determines the synchro rotormovement with respect to the stator.

The fuel flow transmitter has a relief valve,which automatically opens and bypasses theinstrument when the fuel flow exceeds the capacityof the instrument. At such time, only part of thefuel flows through the metering portion. As thepressure across the instrument falls below thevalue at which the relief valve opens, the valvecloses. This lets the flowmeter again operatenormally. The transmitter unit location is in thefuel line between the fuel pump and fuel nozzle.

Fuel Flow Indicator

The fuel flowmeter indicator is located on theinstrument panel. It is a remote-indicatinginstrument. This indicator consists of a synchro

Figure 6-54.-(A) Fuel flow indicator; (B) fuel flow totalizerindicator.

receiver, a step-up gear train, a magnetic drag cup,and a calibrated spring. When fuel flows throughthe fuel flow transmitter, an electrical signal goesto the indicator receiver. This signal drives thesynchro rotor to the proper position. Thus, theindicator pointer shows the rate of fuel flow.

Figure 6-54, view A, shows the face of thesingle flow indicator. To determine the amountof fuel consummation per hour, multiply the scalereading by 1,000. Figure 6-55 shows a schematicdiagram of the single fuel flow indicator.

The following discussion uses the F-14 systemas an example of a fuel flow system. Other aircraftfuel flow systems operate in a similar way.

The fuel flow transmitter consists of asynchronous motor, drum assembly, impellerassembly, spiral spring, and pickup coils. Thetransmitter housing has fuel inlet and outletattachment flanges. The drum and impeller

assemblies have two miniature permanentmagnets, 180 degrees apart. The motor runs ata constant 120 RPM. The motor connects througha shaft to the drum assembly.

The impeller assembly rotates over the motor-drum shaft and mechanically couples to the drumwith the spiral spring. The pickup coils, one foreach assembly, are in line with the drum andimpeller assembly magnets. As the motor rotatesthe drum, the impeller also rotates. When thereis no fuel flow, the magnets of the drum andimpeller assemblies align. As they pass theirrespective coil, they generate simultaneous outputsignals.

As fuel flow starts and increases through thetransmitter housing, it goes through straighteningvanes. These vanes eliminate the swirling motionof fuel. The fuel then passes through straightdrilled passages of the rotating impeller. Asfuel flow increases through the impeller, aproportional drag factor, or resistance to rotation,is imposed on the impeller assembly. Thisresistance causes the spring to deflect, equalizingthe loading. The impeller magnets then deflect outof alignment. This action produces a later signalthan that of the drum magnets and coil. Thus,an increased time span between signals of therotating assemblies becomes relative to increasedfuel flow.

The fuel rate-of-flow power supply consistsof a power transformer and power supply, twosignal-conditioning channels, and a motor driver.The transformer receives 115 volts ac, 400 Hz,and it feeds the power supply. The powersupply provides low dc voltage to operate the

Figure 6-55.-Single fuel flow indicating system.

motor-driven logic and signal-conditioningchannels. Using a stepping signal to drive controllogic, the motor driver controls positive andnegative 8-Hz ac signals between phases of bothfuel rate-of-flow transmitter motors.

The signal-conditioning channel for eachengine system receives pulses from the coils of itstransmitter. A pulse shaper, for each channel,converts the time between transmitter drum andimpeller coil pulses into a rectangular pulsewidthsignal. An averaging filter processes this convertedsignal, which provides a low-ripple dc signal inputto the fuel rate-of-flow indicator. The size (0 to5 volts dc) of this signal is proportional to flowrate. A test circuit permits testing the powersupply. A tap-off motor, phases A and B fromthe motor driver, routes an 8-hertz signal throughan external test switch to the signal-conditioningchannels for processing. The results of theprocessed signal are displayed on the indicator.

The fuel rate-of-flow indicator, a vertical scaleindicator, displays rate of fuel flow for eachengine on parallel scales. The scales are from 0to 13. The scale reading, multiplied by 1,000,show the rate (pounds per hour) at which theengine is consuming fuel. The upper left and rightOFF failure flags show a loss of power, or signalto the indicator. The indicator has two separatechannels, one for each engine. The channelsinclude a control transformer servo amplifier,servomotor, gears, and sprockets. The indicatorchannels receive 115 volts of ac for signalprocessing. An input of 0 to 5 volts dc fromthe fuel rate-of-flow power supply producesan output. This output is from the controltransformer rotor winding to the servo amplifier.

The servo amplifier modifies the signal to theproper impedance and power level to drive thechannel servomotor. Shaft rotation controlstransformer output, and the rotation reducestransformer output voltage. When the output isnull, the motor, gear train, and sprockets cometo rest at a rate equivalent to the input.

The test selector switch on the MASTERTEST panel tests the indicators. The self-testcircuit in the indicators disconnect the fuel rate-of-flow transmitter input circuits. It then connectsto an appropriate test signal within the twoindicator channels. The channel circuitryprocesses the test signal and drives the indicatortapes to indicate 4,200 to 4,400 pounds per hour.

Fuel Flow Totalizing Systems

Figure 6-54. view B, shows the indicator ofa fuel flow totalizing system. The pointer of thisinstrument usually shows the combined rate offuel flow into two or more engines. Also, if onlyone engine is operating, the pointer gives a trueindication. A continuous reading of the poundsof fuel remaining in the aircraft fuel cells appearsin the small window. Before starting the engines,you set the total amount of fuel in the aircrafton the pounds-fuel-remaining indicator by usingthe reset knob on the front of the instrument. Assoon as the engines are running, the fuel flowpointer shows the rate of fuel consummation.The fuel remaining indicator starts countingtoward zero, giving a continuous reading of fuelremaining in the cells. Numbers rotate past thewindow like those of the mileage indicator of anautomobile speedometer.

The entire fuel flow totalizing system consistsof two or more fuel flow transmitters, anamplifier, and an indicator.

FUEL FLOW TRANSMITTERS. —The fuelflow transmitters are almost identical to thosealready discussed in the single system. In the fuelflow totalizing system, the transmitters connectelectrically, so their combined signals go into thefuel flow amplifier as one.

FUEL FLOW AMPLIFIER. —The fuel flowamplifier is an electronic device that suppliespower of the proper size and phasing to drive theindicator. The speed at which the indicator motorruns depends on the transmitter signal going intothe amplifier.

FUEL FLOW TOTALIZER INDICATOR. —The fuel flow totalizer indicator contains atwo-phase variable speed induction motor. Thismotor travels in one direction only; however, thespeed varies. As the rate of fuel consumptionincreases, more and more power goes to theindicator motor. This causes the speed of themotor to increase proportionally to the rateof fuel consumption. The motor turns a magneticdrum-and-cup linkage (similar to the tachometerindicator hysteresis cup), which causes pointerdeflection. The deflection is proportional to themotor speed, and thus proportional to the rateof fuel consumption. At the same time, a linkagehaving a friction clutch drives the pounds-fuel-remaining indicator. The clutch is disengaged

when using the reset knob to set the reading onthe pounds-fuel-remaining indicator.

OIL PRESSURE SYSTEM

Oil pressure instruments show that oil is or isnot circulating under proper pressure. An oilpressure drop warns of impending engine failuredue to lack of oil, oil pump failure, or brokenlines. Oil pressure shows on an engine gauge unit(fig. 6-56). This unit consists of three separategauges in a single case—oil pressure, fuel pressure,and oil temperature. The gauge has a Bourdontube mechanism for measuring fluid underpressure (fig. 6-57). The instrument’s oil pressurerange is from 0 to 200 PSI. You read the scalein graduations of 10 PSI. There is a singleconnection on the back of the case leading directlyinto the Bourdon tube.

In some aircraft, the oil pressure gauge is aseparate instrument (fig. 6-58). This instrumentoperates on the Bourdon tube principle.

The synchro system is another method ofmeasuring oil under pressure. This type of oilpressure system is used on most modern aircraft.Essentially, it is a method of directly measuringengine oil pressure. After taking the measure-ments, they go electrically from the point ofmeasurement to the synchro indicator on theinstrument panel. The synchro system ends theneed for direct pressure lines from the engine tothe instrument panel. It also reduces the chances

Figure 6-57.-Bourdon tube oil

of fire, loss of oil or fuel,difficulties.

pressure gauge.

and mechanical

The synchro system consists of a synchroindicator and transmitter. The synchro transmitterconsists of a permanent magnet moving within astator. The stator is a circular core of magneticmaterial wrapped with a single, continuoustoroidal winding. Taps divide the winding intothree sections. Voltages in each of the sectionsvary with the position of the permanent magnet.As the magnet moves, the ratio between the threesignal voltages varies accordingly.

Look at figure 6-59. Here, you can see thetransmitter and indicator connect in parallel.When excited by the same fundamental source,the signal voltages in corresponding sections ofthe two stators are equal and balanced. The signalvoltages remain equal and balanced as long as the

Figure 6-56.-Engine gauge unit. Figure 6-58.-Bourdon tube oil pressure instrument.

Figure 6-59.-Schematic of a synchro oil pressure indicating system.

magnets are in the same relative positions.However, if the transmitter magnet moves to anew position, the voltages in the three sections ofthe transmitter are no longer the same. They nowdiffer from the voltages in the correspondingsections of the indicator. Because of thisimbalance, current flows between the two units.This circulating current sets up additionalmagnetic lines of force in each stator, whichestablishes a magnetic force between the statorand the magnet of each unit. Since the indicatormagnet is free to turn, it moves to a position corre-sponding to the position of the transmitter magnet.The indicator magnet connects to the indicatorpointer by a shaft to provide a visual indication.The electrical leads between the transmitter andthe indicator may be any reasonable lengthwithout noticeable effect on the indication.

FUEL PRESSURE SYSTEM

The fuel pressure gauge provides a checkon the operation of the fuel pump and fuelpressure relief valve. The pilot must checkthe gauges often to ensure that the fuel pres-sure is correct. With the fuel pressure correct,the engines have a full range of power atall altitudes. The fuel pressure gauge operateson the same principle as the oil pressuregauge.

Fuel pressure indicators may be located in thecockpit by means of synchro systems. This typeof system is the same for both fuel and oil pressureindications. However, the oil system transmitteris not interchangeable with the fuel systemtransmitter.

Look at figure 6-60. This synchro systemis used to show fuel pressure. A change infuel pressure introduced into the synchrotransmitter causes an electrical signal to gothrough the interconnecting wiring to thesynchro receiver. This signal moves the receiverrotor and the indicator pointer a distanceproportional to the amount of pressure exertedby the fuel.

NOTE: You can see in figure 6-60 that thetransmitter has a vent to the atmosphere. Thisallows the transmitter to accurately measurethe differential between pump pressure andatmospheric pressure.

OIL TEMPERATURE SYSTEM

Two types of oil temperature gauges areavailable for use in the engine gauge unit. Oneunit consists of an electrical resistance type of oilthermometer, supplied with electrical current bythe aircraft dc power system. The other unit, thecapillary oil thermometer, is a vapor pressure

Figure 6-60.-Fuel pressure synchro system.

thermometer. It consists of a bulb connected bya capillary tube to a Bourdon tube and amultiplying mechanism connected to a pointer.The pointer shows the oil temperature on a dial.

EXAUST NOZZLE INDICATINGSYSTEM

The exhaust nozzle position indicating systemshows the pilot engine variable exhaust nozzleposition. This, in turn, provides a measure ofpercentage of afterburning, since constanttemperatures are held throughout the afterburnerrange.

Each engine has a separate but identical nozzleposition indicating system. Each system consistsof a transmitter potentiometer in the nozzle areacontrol unit and an indicator on the maininstrument panel. Power for the system is fromthe essential 28-volt dc bus.

Each indicator is a hermetically sealed unitcontaining a single receptacle for a mating plugelectrical connection. The instrument scale rangesfrom OPEN to CLOSE, with markings at the 1/4,1/2, and 3/4 positions (fig. 6-61).

The transmitter potentiometer consists of aresistance winding with a movable brush. This

brush connects to a linkage within the nozzle areacontrol and moves in relation to the variableexhaust nozzle. Current in the resistance windingis picked up by the movable rush, and it variesaccording to the location of the brush. Then, thissignal current goes to one of the indicator fieldcoils.

The indicator contains two field coils anda rotor. The polarized rotor mounts on afree-moving shaft. The shaft is located inthe center of the magnetic field created by thetwo coils. One coil connects to the transmitterpotentiometer in the nozzle area control. Thesecond receives a constant current to givesmooth indicator operation. The rotor alignsitself with the magnetic field. The magneticfield varies as the signal received from thepotentiometer varies. A pointer mounted onthe rotor shaft shows rotor position in relationto nozzle position.

TORQUEMETER SYSTEMS

The electric torquemeter system in turbopropaircraft measures the torque (horsepower)produced by the engine at the extension shaft.Each system consists of a transmitter (part of theengine extension shaft), a phase detector, and an

Figure 6-61.-Exhaust nozzle position indicating system.

indicator (fig. 6-62). The system measures the outer reference shaft, The displacement betweentorsional deflection (twist) of the extension shaft the shafts is proportional to the torque load onas it sends power from the engine to the propeller. the driving shaft. This displacement causes a phaseMagnetic pickups detect and measure this deflec- displacement between the pickup signals. Thetion electronically. The indicator registers the phase angle of the resultant signal is linearlyamount of deflection in shaft horsepower. proportional to the shaft deflection. The phase

The extension shaft of the engine consists of detector and amplifier in the indicator convertstwo concentric shafts. The inner shaft is the power the signal to current. The current goes to atransmitting shaft. The outer shaft attaches to the servomotor, which drives the indicator pointers.inner shaft at the rear end. Toothed flanges on The servomotor also drives the rotor of a synchrothe front end of each shaft rotate in the field of control transformer in the indicator. The synchromagnetic pickups. When the engine is running, control transformer balances the synchrothe teeth on the flange of the driving (inner) shaft system when the pointer registers the measuredmove in relation to the teeth on the flange of the torque.

Figure 6-62.-Torquemeter system: (A) pickup assembly; (B) indicating system.

What tach generator output is proportionalto the engine speed?

When there is an open in the voltage supplycircuit, what will the galvanometer of aWheatstone bridge read?

In a radiometer-type temperature indicator,what component determines needle position?

What is a thermocouple?

In the fuel flow transmitter, what com-ponent senses vane movement and sends itto the synchro motor?

Name two ways of indicating oil pressure.

In the torquemeter system, what shaft is thereference shaft?

MISCELLANEOUS INSTRUMENTSYSTEMS

Learning Objective: Recognize the operat-ing principles and characteristics ofmiscellaneous aircraft instrument systems,including fuel quantity, hydraulic pressure,and position indicating systems.

There is another group of instruments foundon most Navy aircraft that does not come underflight instruments or engine instruments. It is agroup usually referred to as miscellaneousinstruments. This group consists of instrumentationfor such systems as fuel, hydraulics, flap and gearpositions, and cabin pressure.

CAPACITIVE-TYPE FUEL QUANTITYINDICATING SYSTEM

The capacitive-type fuel quantity systemelectronically measures fuel weight (not gallons)

of the fuel in the tanks of an aircraft. The mainunits of the system are an indicator, tank probes,a bridge unit, and an amplifier. In some systems,the bridge unit and amplifier are one unit mountedin the same box. In the design of newer systems,the bridge and a transistorized amplifier arecontained in the indicator.

Fuel Quantity Indicator

The fuel quantity indicator (fig. 6-63) is ahermetically sealed, self-balancing, motor-driveninstrument. It contains a motor, pointer assembly,

Figure 6-63.—Fuel quantity indicator: (A) Radial; (B) verticalscale.

transistorized amplifier, bridge circuit, andadjustment potentiometers. As the quantity offuel in the tank changes, the capacitance value ofthe tank probe changes proportionately. The tankprobe is one arm of a capacitance bridgecircuit. The change of capacitance of the probeunbalances the bridge circuit of the amplifierpower unit. The unbalance in the circuit causesan error voltage. The amplified error voltage goesto the motor. The motor drives the pointermechanism and the rebalancing potentiometers torestore the bridge to a balanced condition. Thedirection of change in the capacitance of the probeunit determines the phase of the error voltage. Thephase determines the direction of motor rotationand, therefore, the direction of pointer movement.

Tank Probe

A tank probe and a simplified version of atank circuit are shown in figures 6-64 and 6-65.The capacitance of a capacitor depends upon threefactors—the area of the plates (A), the distancebetween the plates (d), and the dielectric constant(K) of the material between the plates, or

where A = the area of the plates,

d = the distance between the plates, and

K = the dielectric constant of the materialsbetween the plates.

The only variable factor in the tank probe isthe dielectric of the material between the plates.When the tank is full, the dielectric material isall fuel. Its dielectric constant is about 2.07 at 0°C,compared to a dielectric constant of 1 for air.When the tank is empty, there is only air betweenthe plates, and capacitance is less. Any change infuel quantity between full and empty produces acorresponding change in capacitance.

System Operation

Look at figure 6-66, which shows a simplifiedcapacitance bridge circuit. The fuel tank capacitorand a fixed reference capacitor connect in seriesacross a transformer secondary winding. Avoltmeter connects from the exact center of thetransformer winding to a point between the twocapacitors. If the two capacitance are equal, thevoltage drops across them are equal. Therefore,

Figure 6-64.-Fuel quantity transmitter.

the voltage between the center tap and point Pis zero. As the fuel quantity increases, thecapacitance of the tank unit increases. Thisincrease causes more current to flow in the tankunit leg of the bridge circuit. A voltage, whichis in phase with the voltage applied to thetransformer, now exists. As the quantity of fuelin the tank decreases, there is less current in thebridge tank unit leg. The voltage across the

Figure 6-65.-Simplified tank circuit.

Figure 6-66.-Simplified capacitance bridge circuit.

voltmeter is now out of phase with the voltageapplied to the transformer.

In an actual capacitive-type fuel gauge, theinput to a two-stage amplifier takes the place ofthe voltmeter. It amplifies the signal resultingfrom an imbalance in the bridge circuit. Theoutput of the amplifier energizes a winding of thetwo-phase indicator motor. The other motorwinding is the line phase winding. It receivesconstant power from the same voltage that powersthe transformer in the bridge circuit. However,a series capacitor shifts its phase by 90 degrees.As a result, the indicator motor is phase sensitive.This means that the indicator motor will operatein either direction, depending on whether the tankunit capacitance is increasing or decreasing.

Look at figure 6-67. As the tank unitcapacitance increases or decreases, it is necessaryto maintain the bridge circuit in a balancedcondition. This prevents the indicator motor from

Figure 6-67.-Fuel quantity indicator schematic.

continually changing the position of the indicatingneedle. To do this, a balancing potentiometer(R128) connects across part of the transformersecondary. The indicator motor drives thispotentiometer wiper in the direction necessary tomaintain a continuous balance in the bridge.

The circuit shown in figure 6-67 is a self-balancing bridge circuit. An empty-calibratingpotentiometer and a full-calibrating potentiometerconnect across portions of the transformersecondary winding. You can adjust these potenti-ometers so the bridge voltage balances over theempty-to-full capacitance range of a specificsystem.

A test switch (not shown) unbalances thebridge circuit momentarily when checking theoperation of the system. When the switchactuates, pin F connects to ground, unbalancingthe circuit. As a result, the indicator drives towardthe empty end of the dial. Opening the switchshould restore the bridge to balance and returnthe indicator pointer to its original position. Thisproves that the system is operating correctly.

In installations where the indicator shows thecontents of one tank, and the tank is fairlysymmetrical, one probe is sufficient. However,for increased accuracy in peculiarly shaped fueltanks, use two or more tank units in parallel. Thisconfiguration minimizes the effects of changes inaircraft attitude and sloshing of fuel in the tanks.

There are two classes of tank units (fig. 6-68)used in a typical capacitance fuel quantitymeasuring system—noncharacterized and charac-terized.

Noncharacterized tank units are variablecapacitors, vertically mounted in the fuel cell. Asthe fuel level changes, the capacitance of the tankunit changes. This change in capacitance isuniform the entire length of the tank unit.

Characterized tank units are similar in con-struction and mounting to the noncharacterizedtank units. As the fuel level changes, thecapacitance of the tank unit changes. This changein capacitance is NOT uniform the entire lengthof the tank unit.

Since neither electrode of the tank unit goesto ground, and one lead to the amplifier isshielded, capacitance to ground does not enterinto the circuit. Therefore, the length of the tankunit leads does not affect the accuracy of thesystem.

Figure 6-68.-Fuel quantity tank units: (A) Noncharacterized;(B) Characterized.

FUEL CHARACTERISTICS. —The charac-teristics of fuel are such that the dielectric constantand density deviate because of temperaturechange. Also, the variable factor in the fuelcomposition changes the dielectric constant anddensity. The weight by volume of aircraft fueldepends on its density, which, in turn, dependson its temperature. As the temperature of thefuel goes down, the density increases. As thetemperature of the fuel goes up, the densitydecreases. Any change in the dielectric constantor density of the fuel affects the movement of theindicator pointer. For example, assume that theindicating system is at balance with the tank unitimmersed to a given depth. The fuel it is immersedin has the density and dielectric constant of thefuel for which the system is calibrated. Theindicator pointer will then reflect the correctamount of fuel in terms of pounds. Then, thetanks are drained and then refilled to the samelevel. This time the fuel has a greater density andhigher dielectric constant, which causes thepointer to show a greater weight. The new readingis correct only if the effect of the changes indensity and dielectric constant are proportional.

However, the effect of the increase in dielectricconstant is greater than the effect of the increasein density. The results are an incorrect indication.The system reduces this error by varying thecapacitance of the reference capacitor in thebridge leg opposite the immersed tank unit.

COMPENSATION. —The reference capacitoris varied by connecting a compensator unit inparallel with it. The compensator, like the tankprobe, is a variable capacitor. However, thecompensator mounts at the lowest level of fuelso it is completely immersed until the tank isalmost dry. Its capacitance depends on thedielectric content of the fuel rather than thequantity. The compensator connects into thecommon reference leg for both phases of thebridge circuit. This allows it to become a part ofthe reference capacitance. A change in thedielectric constant of the fuel affects both thetank probe and the compensator capacitance.Therefore, the current change in the tank probeleg of the bridge is counteracted by a similarchange in the reference leg of the circuit.

There are various capacitor-type fuel quantitysystems that operate on the principle justdescribed. Indicators, tank probes, and powerunits may differ as to shape, size, and specifi-cations from system to system. For this reason,you should always consult the manufacturer’smanuals for specific information on a particularsystem.

HYDRAULIC PRESSURE INDICATORS

In most naval aircraft, the hydraulic systemoperates the landing gear, flaps, speed brakes,bomb bay doors, and certain other units. Aircrafthydraulic pressure gauges show either the pressureof the complete system or the pressure of anindividual unit in the system. A typical directreading gauge contains a Bourdon tube and agear-and-pinion mechanism. The mechanismamplifies and transfers the tube’s motion to thepointer. The position of the pointer on thecalibrated dial shows the pressure in pounds persquare inch.

The pumps supplying pressure for operatingthe aircraft’s hydraulic units are driven by anaircraft engine, an electric motor, or by both.Some installations employ a pressure tank oraccumulator to maintain a reserve of fluid underhydraulic pressure. In such cases, the pressuregauge registers continuously. With other installa-tions, operating pressure builds up only when

needed, and pressure registers on the gauge onlyduring these periods.

The pressures of hydraulic systems vary fordifferent models of aircraft. In older pressuresystems, the gauges registered from 0 to 2,000PSI. With later model aircraft, the pressure rangeshave increased. Some aircraft have systems withpressure ranges as high as 4,000 PSI. The trendis away from the direct reading pressure gauge andtowards the synchro (electric) type of gauge.

Figure 6-69, view A, shows the hydraulicpressure indicator of a late model naval aircraft.This aircraft has three hydraulic systems. Theindicating system shows the No. 1 and No. 2 flightcontrol systems and the utility hydraulic systempressures. The indicating system consists of threeremote pressure transmitters, a hydraulic pressureselector switch, and a dual pointer indicator. Thesystem uses 26-volt, 400-hertz, single-phasealternating current from the 26-volt, single-phasebus.

Each hydraulic pressure system line containsa Bourdon tube type of pressure transmitter.Expansion and contraction of the Bourdon tubetravels by mechanical linkage to the rotor of the

Figure 6-69.-(A) Hydraulic pressure indicator; (B) hydraulicpressure indicating schematic.

transmitter synchro. The pressure transmittersynchro sends an electrical signal to the receivingsynchro within the indicator. The receivingsynchro’s rotor links mechanically to the indicatorpointer. The pressure indicator contains twosynchros that attach mechanically to two separatepointers. When the HYD PRESS SELECTORswitch is in the No. 1 and No. 2 FLT CONTposition, the pointers indicate the pressure of eachsystem. When the HYD PRESS SELECTORswitch is in the UTILITY position, the synchrosconnect in electrical parallel. This causes thepointers to act as one. This type of arrangementsaves instrument panel space.

PNEUMATIC PRESSURE SYSTEMS

The cabin pressure altitude indicator (fig. 6-70)is a sensitive altimeter that measures cabinpressure. The instrument contains a sensitivediaphragm that expands or contracts with changesin cabin pressure. The altitude equivalent of cabinpressure shows on the dial, in increments of 1,000feet. The range is from 0 to 50,000 feet. Anopening in the back of instrument case allows itto sense cabin pressure. You can also use thisinstrument to reflect pressure suit altitude ratherthan cabin altitude when wearing a pressure suit.

CAUTION

Removeindicator

the cabin pressure altitudeif the aircraft is to undergo a

cabin pressurization test on the ground.Failure to remove the indicator will resultin damage to the instrument from excessivepressure.

Figure 6-70.-Cabin pressure altitude indicator.

POSITION INDICATING SYSTEM(DC SYNCHRO SYSTEM)

A dc synchro system is a method of showinga remote mechanical condition. Specifically, thesesystems show movement and position of wingflaps, cowl flaps, oil cooler doors, or similarmovable parts of the aircraft.

The system consists of a transmitter, anindicator, and connecting wires. A dc voltagefrom the aircraft’s electrical power system suppliesthe voltage to operate the system. The transmittermechanically connects to the movable device thatis actually supplying the positioning data. Theindicator repeats this information on a properlycalibrated scale in the indicator on the instrumentpanel. Figure 6-71 illustrates a complete three-wiresystem.

Transmitter

The three-wire system transmitter consists ofa continuous circular toroidal resistance windingwith two diametrically opposite brushes con-tinuously touching the winding. These brushesapply dc to the winding. The brushes rotate withthe movement of the aircraft part to which theyare mechanically attached.

Indicator

The three-wire system indicating elementconsists of an annular core, a permanent magnetrotor, a damping cylinder, and three field coils.The leads between the coils connect to the threetaps of the transmitter winding. As voltages at thetransmitter taps vary through brush rotation, thedistribution of current in the indicator coils varies.This causes the resulting magnetic field of thethree coils to position the pointer’s permanent-magnet rotor.

A copper cylinder provides a damping effect.The induced eddy currents in this cylinder opposemovement of the rotor. This reduces the tendencyof the pointer to oscillate.

Landing Gear

An example of a position indicating system isa wheel position indicator. The system consists

of three back-mounted indicators. A series of limitswitches in the landing gear control circuit controlthese indicators. The wheel instrument (fig. 6-72,view A) gives a ready indication of the landinggear position. There are three positions on a drumthat move about an axis. One position of the drumhas a landing gear wheel symbol. The centerposition has a diagonal barber pole (black andyellow warning lines). The other position has theword UP. The miniature wheel shows that thewheel is down and locked. The UP indicatorshows the wheel is up and locked. The barber poleshows the wheel is somewhere between up anddown, or not locked in position.

The position indicator operates through thelanding gear limit switches in the wheel wells ofthe aircraft. When the landing gear moves thecontact of S1 (fig. 6-72, view A) to the upposition, the indicator shows the wheel is up and

locked. When the landing gear moves the contactof S2 to the down position, the indicator showsthe landing wheel. This says the wheel is downand locked. Each switch connects to a solenoidin the indicator. As the solenoids energize, theindicator element moves to reveal the position ofthe aircraft landing gear. When the landing gearis moving, both the up lock and down lockswitches release. This opens both circuits to thecorresponding indicator, causing the black andyellow barber pole to show. It also causes thewarning light in the landing gear handle toilluminate. Figure 6-72, view B, shows a schematicfor a complete aircraft landing gear positionindicating system.

You can show the aircraft flap position in amanner similar to the landing gear. These

Figure 6-71.-Three-wire dc synchro system.

Figure 6-72.—Landing gear position indicating system: (A) typical; (B) system schematic.

Figure 6-73.-Flap position indicating system.

indicators show UP, 1/2, DN (down), and barberpole. The indicator (fig. 6-73) energizes throughlimit switches.

Other aircraft show the position of the flapsusing the dc synchro system of remote indication.This system consists of a transmitter and anindicator (fig. 6-74). A change in flap positionmoves the transmitter rotor, and a similarrotor movement occurs in the indicator on theinstrument panel. A pointer attached to the

indicator rotor shows the amount of travel of theflaps in percent of full extension. The transmittermounts on the flap drive control unit and actuatesby the control actuating mechanism.

Integrated Position Indicator

In many jet aircraft, the design of positionindicators combines various systems into oneinstrument. This instrument is the integrated

Figure 6-74.-Dc synchro flap position indicating system.

position indicator (IPI) (fig. 6-75). This IPI is from the A-6A Intruder, all-weather attack aircraft.

Figure 6-75 shows the IPI in four differentconfigurations—power off, transient, clean, and dirty.The power off and transient positions are self-explanatory. The term clean describes an aircraft thathas all of its gear up or in, and the aircraft profile is asfree of parasitic drag as possible. The term dirty meansthe opposite of clean—that units are protruding out ordown, and that resulting drags exist. These are not theonly aircraft configurations possible. Most of theindividual systems (stabilizer, flaps, speed brakes, etc.)are capable of independent action. All the systems do notfunction simultaneously as shown in the figure.

Views A and B of figure 6-75 show a transitionalperiod that appears the same as power off—barber poles.

For specific operation and exact position indicationsof these instruments, refer to the MIM for the specificaircraft. There are many possible sets of indications.Briefly, however, slats usually position in conjunctionwith flap movements. The stabilizer shows pitch trimcontrol (nose up, nose down). The speed brakes, whichare for reducing airspeed, show either as in or out, withno midposition.

Q1. Name the two types of fuel probes.

Q2. Upon what does the density of fuel depend?

Figure 6-75.—Integrated position indicator (IPI).

INSTRUMENT SYSTEMMAINTENANCE

Learning Objective: Relative to aircraftinstrument system maintenance, identifythe various procedures used to maintain,troubleshoot, test, interchange, protect,ship, and handle aircraft instruments.

You need a wide variety of skills to maintainaircraft instruments and their associated equip-ment. It is important for you to check, inspect,and maintain these instruments because theaircraft will not perform properly unless theinstruments present reliable information.

Instruments used in high-speed aircraft mustgive correct indications. The accuracy ofinstruments such as percent-type tachometers, tailpipe temperature indicators, and gyro attitudeindicators require preventive maintenance. Theexistence of excessive errors in instrument systemsdirectly relates to flight safety and efficientaircraft performance. You cannot assume thatborderline instrument errors are acceptable. Thisis particularly important in high-speed aircraft.

As an AE, you will perform functional testson aircraft instruments to make sure they giveaccurate indications. Operational and functionaltests take time. When you perform an inspection,you need to know how the particular aircraftinstrument operates. Also, you need to know whattools and test equipment you will need. Withoutthese knowledge and skills, you can’t properlyperform the tasks assigned to your rating.

GENERAL MAINTENANCE

General maintenance of instruments fallinto two categories—scheduled and unscheduledmaintenance. The material condition of thesystems and reliability of instruments are ensuredby the day-to-day maintenance routine.

Instruments come in one of four differentkinds of cases.

1. One-piece phenolic composition cases2. Two-piece phenolic composition cases3. Nonmagnetic all-metal cases4. Metallic-shielded cases

The cases come in several different sizes soinstruments can be easily removed and mainte-nance simplified. Special instruments that centainmechanisms too large for adaption to a standardcase come in specially designed cases.

Instruments easily mount on the instrumentpanel with locking devices molded into theinstrument flange assembly, by spring locknuts,or mounting clamps. You can easily removeinstruments that use a mounting clamp byunscrewing the tension screw in the instrument’slower right corner. You do not have to removethe tension screw to release the tension on theclamp assembly.

Markings and Graduation

Markings on the glass instrument face coverhelp flight personnel confirm instrument opera-tion to within the prescribed ranges of theequipment. The markings usually consist of awhite arc on the outer edges of the instrumentglass. They show the normal operating range. Ared mark shows the operating limit that shouldnot be exceeded.

WARNING

An index marking of white paint, not overone-sixteenth inch wide by three-sixteenthsinch long, is at the bottom center of allinstruments color-marked for operatingranges. Place this index mark at the pointbetween the glass and the case. This markwill show whether or not the glass covermoves at any time after marking theranges. Obtain the proper range markingsfor the aircraft instruments from the flightmanual (NATOPS) for the particularaircraft.

Panels

Instrument panels are made from sheetaluminum alloy, with sufficient strength to resist

flexing. The panel is nonmagnetic and painteda dull black to eliminate glare and reflection.Some panels are constructed in two layers,and the instrument faces are flush with therear panel. The front panel is a reflector panelthat mounts over the rear panel with sufficientclearance to supply an indirect lighting effect. Theindirect lighting system is not standard for allaircraft. Some aircraft have spotlights, edgelighting, or a combination of these. Someinstruments have their own internal lightingsystem.

Instrument panels are shock-mounted toabsorb low-frequency, high-amplitude shocks.The mounts consist of square-plated absorbers insets of two, each secured to separate brackets.You should inspect the mounts periodically fordeterioration; if the rubber is cracked, replace thepair.

As an AE, your instrument maintenanceduties include certain inspections that youshould conduct at regular intervals. Duringa daily inspection, you inspect the followingItems:

Check pointers for excessive errors.Some indicators should show existingatmospheric pressure, existing tempera-tures, etc. Others should indicate zero.

Check instruments for loose or crackedcover glasses. Replace pitot-static instru-ments if damaged.

Check instrument lights for proper opera-tion.

Check caging and setting knobs forfreedom of movement and correct opera-tion.

Carefully investigate any irregularity thepilot reports.

When performing a phase/calendar inspec-tion, make the following checks:

• Check the mounting of all the instru-ments and their dependent units forsecurity.

Check for leaks in instrument cases, lines,and connections.

Check for dull or marred luminous painton dial markings and pointers.

Check the condition of operation andlimitation markings. Also, check theircondition.

Check for contact and condition ofbonding on instruments.

Check shock mountings for condition ofrubber and security of attachment.

Check for freedom of motion of all linesand tubing behind the instrument panel.Also, check that they are properly clampedor taped to avoid chafing, and free frommoisture, crimps, etc.

After starting the engine, check the instrumentpointers for oscillation. Also, check the readingsfor consistency with engine requirements andspeeds. On multiengined aircraft, check theinstruments for the various engines against eachother. Investigate any inconsistency; it mayindicate a faulty engine, component, or instru-ment.

After you have diagnosed a particular troubleand found the instrument to be faulty, removeand turn it into supply. Remember, the followingprecautions when removing and installinginstruments:

• Handle instruments carefully at all times.Additional damage may result if you abusethe instrument.

• Do not change the location of an indi-ator .

• Do not force the mounting screws. If thescrew is cross-threaded, replace it. Do notdraw the screws up too tight against the panel.This may distort the case enough to affectthe operation of the instrument, crack thecase, or break off the mounting lugs.

�• When removing or installing tubing of apressure-operated instrument, use abackup wrench to avoid twisting the tub-ing or fitting. Do not exert undueforce while tightening the connection.

• Install all electrical plugs hand tight.

• Before connecting an electrical plug to aninstrument, check the plug for bent orbroken pins.

• Cap the open electrical receptacles, plugs,and hose connections to prevent foreign mate-rial from entering the instrument or system.

Lubrication

Most instruments require little or no lubrica-tion in the field. The shafts and bearings orinstruments are lubricated before assembly.No further lubrication is necessary until theinstrument goes to a NADEP for overhaul. Onlyspecially trained personnel perform overhauloperations for aircraft instruments.

Aircraft Plumbing

Rigid and flexible tubing is extensively usedin aircraft. These tubes come in many differentsizes. Sizing is usually determined by the outsidediameter of the tube and ranges from one-eighthinch to 2 inches in diameter. The type of materialand wall thickness determine the amount ofpressure that a tube can safely withstand. Whenreplacing or repairing tubing, you should usecaution to make sure that you use the proper type.You can find detailed information on tubing andtubing repairs in Aircraft Structural Hardware,NAVAIR 01-1A-8.

RIGID TUBING. —Rigid metal tubes arewidely used in aircraft for fuel, oil, coolant,oxygen, instrument, and hydraulic and vent lines.Corrosion-resistant steel (stainless steel) andaluminum alloy tubing are the most commonlyused tubing. You may identify the basic tubematerial by either visual inspection or by the alloydesignation stamped on the tubing.

Tube fittings connect tubes together andconnect tubes to instruments. The shape ofthe fitting is determined by the particular

installation—some are straight, while others havevarious angles. The fittings secure to the tube bya beaded or flared joint. The beaded or upset jointis used in low-pressure lines. Systems that uselow-pressure lines include vacuum, deicer, andoil systems that use rubber hose fittings. Allhigh-pressure and some low-pressure lines useflared joints. Grip dies and flaring or beadingtools are used to form flared and beaded joints.When working rigid tubing, you should consultthe manuals on the beading and flaring tools touse the tools properly.

If piping and lines are damaged, you shouldreplace them with new parts. To repair tubing,you must determine how much tubing to remove.Some things you need to consider when decidinghow much tubing to remove include the follow-ing:

• Location of the tubing

• The extent of damage

• The most convenient location for toolmanipulation

There is a tendency to overtighten tubing nutsto prevent high-pressure fluid from escaping. Suchovertightening may severely damage or completelycut off the tube flare. When you remove a tube,check the flare. If you find a flare with less than50 percent of its original wall thickness, reject thetube.

In bending tubing, you must be careful toprevent collapsing of the tube at the bend. Whenmaking bends for fluid tubing, make sure thatyou use the proper bending radius. Thesespecifications can be found in NA 01-1A-8.

Bands of paint or strips of tape around theline near each fitting identify each rigid line in theaircraft. There is at least one identifying markerin each compartment.

All lines less then 4 inches in diameter haveidentification tape on them. The exceptions to thisare cold lines, hot lines, and lines in an oilyenvironment. Another except ion is any line inengine compartments where there is a chance ofthe tape going into the engine intake. In thesecases, and all others where you do not use tape,use paint to identify the lines.

Identification tape codes show the function,contents, hazards, direction of flow, and pressurein the fluid line. When applying these tapes, referto MIL-STD-1247. MIL-STD-1247 standardizes

rigid line identification throughout the Depart-ment of Defense (fig. 6-76).

The function of the line is identified by a1-inch-wide tape that contains printed word(s),color(s), and geometric symbols. Functionalidentification markings (see MIL-STD-1247) arethe subject of an international standardizationagreement. Three-fourths of the total width onthe left side of the tape is a color or color code.This code shows one function only per color orcolors. The function of the line is printed inEnglish across the colored portion of the tape. Anon-English-speaking person can troubleshoot ormaintain the aircraft if he/she knows the colorcode. The right-hand one-fourth of the functionalidentification tape contains a geometric symbol.This symbol is different for every function. Thisallows a colorblind person to identify the linefunction by means of the geometric design ratherthan by the color. Look at figure 6-77. Here, yousee a listing of the functions and their associatedidentification media as used on the tapes.

The identification-of-hazards tape shows thehazard associated with the contents of the line.Tapes that show hazards are one-half inch wide,with the abbreviation of the hazard printed acrossthe tape. There are four general classes of hazardsfound with fluid lines.

1. Flammable material (FLAM). The hazardmarking FLAM identifies all materials ordinarilyknown as flammable or combustibles.

2. Toxic and poisonous materials (TOXIC).TOXIC identifies lines containing materials thatare extremely hazardous to life or health.

3. Anesthetics and harmful materials (AAHM).AAHM identifies all materials producinganesthetic vapors and all liquid chemicals and

compounds hazardous to life and property.However, they do not normally producedangerous quantities of fumes or vapors.

4. Physically dangerous materials (PHDAN).PHDAN marks a line carrying material that is notdangerous within itself. However, the material isasphyxiating in confined areas or is generallyhandled in a dangerous physical state of pressureor temperature.

Table 6-3 lists some of the fluids that you maywork with and the hazards associated with each.

Table 6-3.-Hazards Associated With Various Fluids

Figure 6-76.-Electrical line identification application.

Figure 6-77.-Functional identification tape data.

FLEXIBLE TUBING (HOSE). —Flexiblehose assemblies consist of lengths of hose coupledwith threaded end fittings. There are two classesof flexible tubing—high pressure and lowpressure.

You can get the specifications for flexible hoseby interpreting the identification code on the hose.This identification, a series of dots and dashes,gives hose size, temperature range, and date ofmanufacture. The date of manufacture is inquarter of year and year. Refer to NA 01-1A-8for a detailed discussion of flexible hose identi-fication.

You cannot construct high-pressure flexiblehose at the organizational level. You must orderit through supply, as organizational maintenancecannot perform high-pressure tests.

You can order the parts for a low-pressureflexible hose through supply and make the hoselocally. You can reuse fittings from a damagedhose if they meet the required specifications.

When you install hose, make sure it will nottwist under any operating condition. This typeinstallation lessens the tendency for connectingfittings to loosen. When you replace hose inhydraulic, fuel, oil, alcohol, and pneumaticsystems, make sure the new hose is an exactduplicate of the old hose. By this, the length,outside diameter, inside diameter, material, type,and shape (except on directed modifications) mustbe the same as the old hose.

If a bend is necessary when installing hose influid systems, the radius must not be smaller thanthe minimum specified in NA-01-1A-8. Thispublication shows the minimum bend radii forflexible hose. When practical, it is desirable to usea radius that is larger than the specified minimum.

When you install hose through holes inbrackets and when using supporting clips, therecan be no reduction in hose diameter. If theseconditions are present, they reduce flow, anddamage to the hose may occur.

The hose must have support every 24 inches.Closer supports are desirable when practical. Theflexible line support should never cause deflectionof the rigid connecting lines under any possiblerelative motion that may occur. Flexible hosebetween two rigid connections may have excessive

motion restrained where necessary, but shouldnever be rigidly supported. To avoid chafing usesuitable bulkhead-type grommets or cushionedclips.

Protect hose installations from excessivetemperature, such as exhaust blasts, superchargerducts, and the like, either by shrouding orrelocation. Use flame-resistant hose forward ofthe firewall on certain aircraft or as NAVAIRinstructions may direct.

Where hoses connect to an engine or toengine-mounted accessories, provide 1 1/2 inchesof slack between the last point of support andengine attachment. This prevents the chance ofthe hose pulling off the nipple due to enginemovement, Whenever possible, install the hose soall hose markings are visible.

All hose materials deteriorate from exposureto heat, sunlight, excessive moisture, and ozone.Therefore, you should stow hoses in a cool dryplace and away from electrical equipment. Youcan obtain age limits of shelf items based on themanufacturer’s code from current accessorybulletins. Stow hose in straight lengths to preventit from setting in a curved position. Replace thehose if cover material is peeling or flaking, or thebraid reinforcement is open to the elements.

PITOT-STATIC INSTRUMENTMAINTENANCE

Maintenance of the pitot-static system consistsof checking the lines for integrity, water, andmiscellaneous obstructions. A pressure check isof the utmost importance as any slight leak willresult in erroneous indications of the instrumentsduring flight. You should also check the pitotheater for proper operation. You can do this bymonitoring a voltage drop when the heat is turnedon, or by checking the pitot tube for the presenceof heat.

CAUTION

Do not touch the pitot tube with your barehand with the heat on. The extreme heatmay cause your skin to stick to the surface.

The following is an outline for water anddebris removal from the pitot-static system.

For specific procedures, you should refer to theMIM.

1. Disconnect all altimeters, airspeed, andrate-of-climb indicators, and any other systemsreceiving information from the pitot-static system.Disconnect the lines from pitot or pitot-statictubes.

2. Remove all drain caps in the system.

3. Circulate a stream of clean, dry, filteredair at medium pressure through the completesystem. Be careful not to include the cabinpressurization system static vent. Be certain thatair is flowing from the exit end of each line.

4. Inspect all static vents and the pitot tubewater removal drain holes for damage andevidence of foreign matter and obstructions.Check all low points in the lines for possible cracksdue to icing in the lines.

5. Replace and secure all system drain caps.

6. Reconnect all instruments. Tighten con-nections properly; do not kink or bend the lines.

7. Using a field test set or other approvedtester, thoroughly check the system for properoperation and leaks.

The maintenance of the pitot-static system isrelatively simple when compared to more complexsystems. However, its maintenance is not a minortask.

INSTRUMENT TESTING

Operation of most aircraft instruments isentirely automatic. Once installed, the unitsrequire no further maintenance or servicing otherthan routine and periodic inspections. If asystem or instrument malfunctions, you mustfirst localize the source of trouble. Develop asystematic troubleshooting procedure. Theprocedure should include the possible servicetroubles and their remedies for each type ofinstrument. You will find most of this informationin applicable aeronautic publications, such asthe MIM for specific aircraft and the serviceinstruction manual for specific instruments.

An instrument that doesn’t function properlyor that is suspected of being unserviceable must

first be checked to determine if the instrument orthe installation is at fault. Usually, instrumentproblems fall into three groups—trouble in thepower supply, trouble in the unit, or trouble inthe connections to units, either electrical ormechanical. If the installation is faulty, linemaintenance can correct the problem. If theinstrument is the fault, you can remove andreplace it with a serviceable unit. The defectiveunit goes to a qualified instrument overhaul depotfor detailed inspection, overhaul, and repair.

Only authorized instrument shops can openinstrument cases and make repairs or adjustments.

When making ground tests of electricalinstruments, you should connect an externalpower supply to the aircraft. Do not u s ethe battery when conducting ground tests ofequipment.

When performing ground testing, you shoulduse portable field test sets, such as theTTU-205C/E pressure-temperature tester and theTTU-27/E tachometer tester. Test sets arediscussed in chapter 2 of this TRAMAN.

Always use a precision voltmeter to checkinstrument power. You can check most electricalinstruments with a test indicator to determinewhere the trouble lies. For example, you can checksynchro indicators using a synchro test indicator.

TROUBLESHOOTING

The manuals you use for instrument mainte-nance contain troubleshooting charts. Thesecharts help you determine what is wrong with aninstrument system. When you know what theproblem of a particular instrument or systemis, refer to the troubleshooting chart for theparticular instrument. The chart gives a listing ofthe common troubles, probable causes, andremedies.

Efficient troubleshooting calls for an orderly,systematic plan of attack and a good under-standing of the theory of equipment operation.A visual inspection should be made first; this willfrequently pinpoint the trouble. When makingthis inspection, look for discolored or burnedwires and terminal boards and corroded switchcontacts. Also look for broken or frayed wires,loose connector plugs, and loose mechanical

assemblies. Tables 6-4 and 6-5 are typical for the correct type of instrument. This manualtroubleshooting charts. lists the part number of the instrument.

If you cannot find an exact replacement forINTERCHANGEABILITY OF an instrument, an interchangeable instrument mayINSTRUMENTS sometimes be used to prevent delay in returning

an aircraft to service. Aviation supply itemsYou need to be careful when substituting with the same stock numbers are operationally

aircraft instruments. Check the illustrated parts interchangeable, regardless of the manufac-breakdown (IPB) manual for the specific aircraft turer or the manufacturer’s part number. An

Table 6-4.-Troubleshooting Chart—Airspeed Indicator

Table 6-5.-Troubleshooting Chart—Fuel Flow Indicating System

instrument that has a stock number has at leastone distinctive feature that lets you identify itindividually. Normally, you will not substitute oneinstrument for another; however, in an emergencyit may be necessary. Before you make a substi-tution, properly identify the instrument in theAviction Supply Catalog, which pertains toinstruments.

EMERGENCY PROTECTIVETREATMENT

Aircraft instruments that were submerged inwater must be given treatment to minimizecorrosion as soon as practicable after immersion.Refer to Avionic Cleaning and Corrosion Pre-ventive Control Manual, NAVAIR 16-1-540, forthe correct procedures.

Many of the parts of instrument systems, suchas amplifiers, transformers, and capacitors, havea protective treatment to inhibit fungus growth.This treatment consists of coating the componentswith a special type of fungus-resistant varnish orlacquer. The manufacturer or an aviation depotusually performs this operation. Some mainte-nance activities may do this job. Since the varnishand lacquer are poisonous, this is a dangerousoperation. It requires special equipment andspecially trained personnel.

SHIPPING AND HANDLING

You should handle non-RFI instruments withthe same care you handle new instruments toavoid additional damage. To prevent damage inhandling and storage, all instruments are storedand transported in individual metal containersand well-packed cartons. Use care in packinginoperative instruments so no additional damageoccurs during shipment. Pack individual cartonsfor shipment in strong wooden boxes, except whenprohibited.

Instruments that have a locking mechanismshould be placed in the locked position beforebeing shipped. Refer to the manufacturer’sinstructions for information regarding the lockingof a particular instrument.

Cleaning

Clean all aircraft instruments after removalfrom the aircraft and before shipping to supply.

Drain instruments that are fluid transmitters orpressure indicators to ensure that no fluid remainsinside the instrument.

Be careful you don’t contaminate instrumentswith perspiration. Thoroughly dry the instrumentbefore wrapping it with moistureproof material.

Containers

All instruments and accessories coming fromsupply are in containers. There are two types ofcontainers—nonmetallic and metallic.

The nonmetallic container is usually of ahigh-grade cardboard construction. It will notwithstand the rough treatment that the metallictype container can endure. However, if they arenot abused, they afford adequate protection forthe instrument.

The metallic type is a heavy gauge metal andhas a lid that secures to the can with a snap-ringfastener. You can open and close this containerwithout damaging it.

Do not damage or throw away packagingmaterial from both the nonmetallic and metalliccontainers. Make provisions for retaining thismaterial in the electric shop since it may be usedfor returning equipment to supply. Many ofthese containers have labels stating REUSABLECONTAINER—DO NOT DESTROY.

Q1. What type of marking is used to identifydial face slippage?

Q2. When installing rigid tube markings, whatis the minimum number per compartment?

Q3. What is the color of the identfication tapeon electrical conduits?

Q4. Name the four classes of hazards identifiedby identification of hazards tape.

CHAPTER 7

COMPASS AND INERTIALNAVIGATION SYSTEMS

The material in this chapter is about aircraftnavigation systems. The basic systems discussedare the aircraft compass system and inertialnavigation system (INS). Also, this chapterpresents a discussion of the calibration of thesetwo systems.

The way electrical signals are detected,amplified, and delivered to various indicators andsystems is highly sophisticated. Before you beginthis chapter, you might need to read NavyElectricity and Electronics Training Series(NEETS), Module 15, Principles of Synchros,Servos, and Gyros.

NAVIGATION TERMSAND DEFINITIONS

Learning Objective: Recognize thenavigation-related terms and definitionsbasic to compass and inertial navigationsystem operation.

Any purposeful movement in the universeinvolves an intention to proceed to a definitepoint. Navigation is the business of proceedingso you will arrive at that point. Air navigationis defined as the process of directing themovement of an aircraft from one point toanother. The function of air navigation is to locatepositions and measure distance and time along theintended direction of flight.

POSITION

Position is a point defined by stated orimplied coordinates. You will frequently qualifythis term by such adjectives as estimated, deadreckoning, no wind, etc. However qualified, theword position always refers to some place that youcan identify. One of the basic problems of thenavigator is that of fixing his position. If hedoes not know where he is, he can’t direct

the movement of the aircraft to its intendeddestination.

DIRECTION

Direction is the position of one point in spacerelative to another, without reference to thedistance between them. Direction may be eitherthree-dimensional or two-dimensional; the hori-zontal being the usual plane of the latter. Forexample, the direction of San Francisco from NewYork is approximately west (two-dimensional).However, the direction of an aircraft from anobserver on the ground may be west and 20°above the horizontal (three-dimensional). Direc-tion is not itself an angle (for example, thedirection east), but it is often measured in termsof its angular distance from a reference direction.

COURSE

Course is the intended horizontal direction oftravel. For example, the direction of NASJacksonville from NAS Pensacola is east. Thisshould be the intended direction of flight.However, because of wind conditions aloft,the aircraft might not head straight towardJacksonville, but somewhat to one side. Nomatter what the aircraft heading is, the course (theintended direction) is still east.

HEADING

Heading is the horizontal direction in whichan aircraft is pointing. In the previous example,you can see the difference between course andheading. Heading is the actual orientation of theaircraft’s longitudinal axis at any instant, whilecourse is the direction of travel intended. Trueheading uses the direction of the geographic NorthPole as the reference. Magnetic heading uses thedirection of the earth’s magnetic field at thatlocation as the reference. Magnetic headingdiffers from true heading by the amount of

magnetic variation (MAGVAR) at that location.Compass heading differs from magnetic headingby the amount of magnetic deviation. Compassheading differs from true heading by the amountof compass error (deviation ± variation).

BEARING

Bearing is the horizontal direction of oneterrestrial point from another. Bearings can beexpressed by two terms—true north or the direc-tion in which the aircraft is pointing. If true northis the reference direction, the bearing is a truebearing. If the reference direction is the headingof the aircraft, the bearing is a relative bearing.If you get a bearing by radio, it is a radiobearing; if visual, it is a visual bearing. Thus, thedirection between two objects on (or near) thesurface of the earth can be described concisely bysaying: THE (RADIO, VISUAL) BEARING OFA FROM B IS X ± (RELATIVE, TRUE).

DISTANCE

Distance is the separation between two points.To measure distance, you measure the length ofa line joining the two points. This seemsunderstandable enough. However, suppose thatthe two points are on opposite sides of a baseball.How do you draw the line? Does it run throughthe center of the ball, or around the surface? If

Figure 7-1.-Schematic representation of earth showing axisof rotation and equator.

around the surface, what path does the linefollow? You must qualify the term distance usedin navigation to show how to measure thedistance. The shortest distance on the earth’ssurface from NAS San Diego to Sydney,Australia, is 6,530 miles. However, via Honoluluand Guam, a frequently used route, it is 8,602miles. You can express the length of a chosen linein various units, such as miles, kilometers, oryards.

Time has many definitions. The two defini-tions used with navigation are as follows: (1) thehour of the day and (2) an elapsed interval. Thefirst appoints a definite instant, as takeoff timeis 0214. The second definition appoints aninterval, such as time of flight, 2 hours 15 minutes.

The earth’s geographic poles are the extremitiesof the earth’s axis of rotation. Look at figure 7-1.Here, Pn, E, Ps and W represent the surface ofthe earth at sea level. Line PnP s is the axis ofrotation. The earth’s rotation is such that allpoints in the hemisphere, PnWPs, approach thereader. Those points in the opposite hemispherewill recede from the reader. The extremities of theaxis, points Pn and PSare the north and southpoles, respectively. A man on the surface of theearth, facing in the direction of rotation, has theNorth Pole on his left. East will be in front ofhim, the South Pole on his right, and west behindhim.

The earth has some of the properties of a barmagnet. The magnetic poles are the regions near

Figure 7-2.-The equator is a great circle whose plane isperpendicular to the polar axis.

the ends of the magnet. This is where the highestconcentration of magnetic lines of force exist.However, the earth’s magnetic poles are not atthe geographic poles, nor are they antipodal(opposite) to each other.

GREAT CIRCLES ANDSMALL CIRCLES

The intersection of a sphere and a plane is acircle. The intersection is a great circle if the planepasses through the center of the sphere. It will bea small circle if it does not.

PARALLELS AND MERIDIANS

Look at figure 7-2. Here, the earth’s equatoris a great circle. If a second plane (fig. 7-3) passesthrough the earth parallel to the equator, itsintersection is a small circle. Small circles don’talways have planes perpendicular to the polar axis.However, if they are perpendicular, then all pointson the small circle are equidistant from theequator; that is, the circles are parallel tothe equator. Such small circles, together withthe equator, are PARALLELS. They provideone component of a system of geographicalcoordinates.

Now, suppose that planes pass through theearth’s poles (fig. 7-4). Such planes contain theaxis, and since they also contain the center, theyform great circles at the surface. Great circlesthrough the poles of the earth are MERIDIANS.All meridians are perpendicular to the equator.Meridians form the second part of a system ofgeographical coordinates commonly used bynavigators.

LATITUDE AND LONGITUDE

You can identify any point on earth by theintersection of a parallel and a meridian. It is the

Figure 7-3.-The plane of a parallel is parallel to the equator.

Figure 7-4.-Great circle through the poles form meridians.

same as an address at the corner of FourteenthStreet and Seventh Avenue. You are just usingdifferent names for identifying the parallels andmeridians.

The circumference of a circle is divided into360 units. This unit is the degree. It is the sameunit you use to measure an angle. In figure 7-5,

Figure 7-5.-Latitude of M is angle QOM or arc QM.

circumference QPnQPs represents a meridian.QQ´ represents the equator, whose plane passesthrough the axis of rotation. Let M be someposition north of the equator on a meridian. Thenumber of degrees in arc QM is the measure ofangle QOM. If arc QM is 30°, then angle QOMis 30°. Thus, you measure a central angle bymeasuring its subtended arc.

Let MM´ be the plane of a small circle parallelto QQ´, the equator. Then arc QM measures thedistance of any point on MM´ from the equator.You can describe the whole parallel MM´ bysaying that it is 30° north of the equator.Similarly, you can say any point on NN´ is 45° southof the equator. The angular distance of aposition north or south of the equator is theposition’s latitude. You measure latitude north-ward or southward through 90° and label it N orS to show the direction of measurement. Youexpress latitude in terms of the angle at the center(see angle QOM in figure 7-5). Latitude, then, isthe north-south geographical coordinate.

The east-west geographical coordinate islongitude. You can define longitude in threeways:

1. as an arc of the equator or a parallel,2. as the angle at the pole or the angle at the

center between the planes of the prime meridian,and

3. as the meridian of a point on earth. Youmeasure this point eastward or westward from theprime meridian through 180°. Label it E or Wto show the direction of measurement.

You measure latitude from a standard greatcircle (the equator). You also use a standard greatcircle when measuring longitude. This greatcircle is the meridian. The standard meridian isthe prime meridian. By international agreementin 1884, the meridian adopted as the prime merid-ian was the one on which Greenwich Observatory(near London, England) was located. This is wasthe 0° longitude.

Figure 7-6 represents the earth. QQ´ is theequator, with its plane passing through the center,O. G is the position of Greenwich, and M is aposition in north latitude west of Greenwich.PnM´ and PnG´ are portions of the meridiansthrough M and G intersecting the equator at M´and G´. The longitude of M includes

1. the arc of the equator G´M´,2. the angle GPnM formed by the meridional

planes through G and M, and3. the angle G´OM´ formed at the center

between G´ and M´.

You should recognize that each of these expres-sions measures the same coordinate—longitude.

The longitude of a position is also describedas being east or west of Greenwich. In figure 7-6,the longitude of M is west; the longitude of N iseast. You can see that longitude east or westcannot exceed 180°.

You can subdivide the degree into smallerunits, as in the decimal system. However, themore common method of subdivision is to divideeach degree into 60 minutes (´) of 60 seconds (´´)each. Another method is to divide the degree into60 minutes and tenths of minutes.

Figure 7-6.—Longitude is measured between meridians.

Figure 7-7.-Easterly magnetic variation.

To convert minutes into decimals of degrees,or to convert seconds into decimals of minutes,divide by 6. Thus: 15°30´ = 15.5´, and 15°30´24´´1 5 ° 3 0 . 4 ´ .

VARIATION

As stated under the definitions of poles, theearth’s true (geographic) poles and its magneticpoles are not at the same locations. Also, thelocation of the magnetic poles changes slightlyover the years. In 1960, the north magnetic polewas at latitude 74.9°N and longitude 101.0°W.The southern pole was at latitude 67.1°S andlongitude 142.7°E. Thus, a given line will havea different direction to the true North Pole thanto the magnetic North Pole. In addition, lines ofmagnetic force are not generally straight linesbecause of irregular iron deposits near the earth’ssurface. Since a compass needle aligns to the linesof force at its location, it may not point to trueor magnetic north. The locations on the earthwhere the compass does point to true north, whenconnected together, form an irregular line. Thisis the agonic line. At other locations, theangle between the direction of true north and thedirection of the earth’s magnetic field is thelocations variation. The earth’s magnetic fielddirection may not be the same as the direction ofthe magnetic poles. This same angle is also oftencalled the angle of declination. You label varia-tion (or declination) east or west as the magneticfield direction is east or west, respectively, of truenorth. (See figures 7-7 and 7-8. ) Lines connectinglocations having the same variation are isogoniclines.

Figure 7-8.-Westerly magnetic variation.

Figure 7-9.-Deviation changes with heading.

DEVIATION

Deviation is the error in a magnetic compasscaused by nearby magnetic influences. Theseinfluences may relate to magnetic material in thestructure of the aircraft and to electrical(electronic) circuits. These magnetic forces deflecta compass needle from its normal alignment withthe earth’s magnetic field. You express theamounts of such deflections in degrees. Thedeflection will be east or west as the compasspoints east or west, respectively, of the earth’smagnetic lines of force. Deviation varies with theheading of the aircraft. Figure 7-9 shows onereason for this deviation.

For example, suppose that you represent thenet result of all magnetic forces inherent in anaircraft by an arrowhead in the aircraft’slongitudinal axis and aft of the compass. If theaircraft is heading toward magnetic north, themagnetic forces (arrowhead) attract the south-seeking end of the compass needle. However, theydon’t change the needle’s direction because theinherent magnetism has the same polarity as theearth’s field. Now, suppose that the aircraft takesan east magnetic heading. The aircraft’s magneticforces now repel the north end of the compassneedle and attract the south end, causing easterlydeviation. The figure also shows that thedeviation when heading south is zero and whenheading west is westerly. You can reduce deviation

by changing the position of small compensatingmagnets in the compass case. However, it isusually not possible to remove all the deviationon all headings. You must determine the residualdeviation for each compass installation and recordit on a deviation card. The card shows theactual deviation on various headings or, morefrequently, the compass headings for variousmagnetic headings. You can accomplish thisusing a process known as compass swinging.

COMPASS ERROR

The net result of both variation and deviationis the compass error. If variation and deviationhave the same name (east or west), you add toget compass error. If they have different names,subtract the smaller from the larger. Give thedifference given the name of the larger. (See fig.7-10. ) You can label variation and deviation plus(+) if east, and minus (–) if west. In this case thecompass error is the algebraic sum of the two.

Example 1.

Example 2.

MAGNETIC DIP

At the magnetic poles, the direction of theearth’s magnetic field is vertical (perpendicularto the earth’s surface). Along the aclinic line(sometimes called the magnetic equator), roughlyhalf way between the poles, the field’s directionis horizontal (parallel to the earth’s surface). Thedifference between the direction of the earth’sfield and the horizontal at any location is themagnetic dip. The magnetic dip varies from very

Figure 7-10.-Effect of compass error.

small angles near the equator to very large anglesnear the poles. You can measure the angles witha dip needle, which is a magnetic needle free toturn about a horizontal axis. At San Franciscothe dip angle is about 62°. A line connecting alllocations having equal dip angles is an isoclinicline.

The total intensity of the magnetic field isalong the dip angle. However you can show it astwo components—vertical and horizontal. (Seefig. 7-11. )

Only the horizontal component is effective asa directive force for a magnetic compass (wetcompass). It loses its effectiveness near themagnetic poles because of the weak horizontalcomponent there. The vertical component causeserrors in a magnetic compass during aircraftmaneuvers that tilt the compass card east or west.If an aircraft heading east increases its speed, orwhen heading west decreases its speed, thecompass card tilts. Also, when turning east froma north or south heading, the floating compasscard will tilt. In both cases, the east side of thecard sinks and the west side rises. The verticalcomponent of the earth’s field causes thecompass card to rotate to the east when in thenorthern hemisphere. It will cause the card torotate west when in the southern hemisphere. Theamount of error is zero at the aclinic line, and it

Figure 7-11.-The earth’s magnetism.

increases toward the magnetic poles. It should beclear that precise turns are difficult if referencedto such a compass.

PILOTAGE

Pilotage is the directing of aircraft from pointto point by visual or radar observation of land-marks. These landmarks are either previouslyknown or recognized from a chart. It is similarto taking a trip by automobile where the highwayis the course taken and the towns are the checkpoints. This method has obvious limitations if theflight is made over a large body of water or poorlycharted area. Also, if you make the flight using

only visual observations, darkness, rain, orfog further limit its use. Therefore, wheneverpossible, use pilotage in conjunction with othermethods of navigation.

DEAD RECKONING

Dead reckoning is the process of determininga position from the record of a previously knownposition, course, speed, and time traveled. To beaccurate, you must consider every change ofcourse and speed during the flight. It does notmatter whether the pilot or the air mass (wind)through which the aircraft is flying makes thechanges.

RADAR NAVIGATION

Modern radar can be a valuable aid to naviga-tion. Some radars present a maplike display ofthe terrain around the aircraft on the screen ofa cathode-ray tube (CRT). This allows pilotageto go beyond some of the limitations of visualobservations.

Radar transponders are devices that do notoperate until interrogated or triggered intoaction by a suitable signal from another radartransmitter. Then, they transmit their own signal,which the interrogating radar receives. These areused both for fixed navigational aids, such asradar beacon stations and for airborne identifica-tion friend or foe (IFF) systems.

Doppler radar can detect and show actualground speed and drift of an aircraft, regardlessof wind speed or direction.

Radar altimeters give the actual distance fromthe aircraft to the surface below. The surfacebelow can be a body of water or land masses farabove sea level.

RADIO NAVIGATION

Radio navigational aids vary from a fairlysimple direction finding receiver to complexsystems using special transmitting stations. Thesespecial stations make it possible to fix theposition of an aircraft with considerable accuracy.The usable range varies according to its intendeduse and also with weather and ionosphericconditions. Beacon stations associated with aninstrument landing system (ILS) are usually of lowpower. Long-range air navigation (loran) stationshave a range extending to 1,400 miles underfavorable conditions. Aviation ElectronicsTechnicians (ATs) maintain the airborne portionsof radio and radar systems.

CELESTIAL NAVIGATION

Celestial navigation is the method of fixing theposition of the aircraft relative to celestial bodies.Since the earth is constantly revolving, an accuratetime device is necessary. You may use a sextantto measure the angle of the celestial bodies withrespect to the horizon. In marine navigation, thevisible horizon is the reference. In air navigation,you use an artificial horizon as the referencepoint. Also, the navigator needs an almanac todetermine the celestial equator system coordinatesat the time of observation.

The usual method to show a line ofposition from celestial observation consists of(1) observation, (2) coordinate conversion, and(3) plotting. The navigator tries, wheneverpossible, to select three bodies about 1200 apartin azimuth. This not only results in lines ofposition that cross cleanly, it also minimizes theeffects of a constant error in the observations.

INERTIAL NAVIGATION

Inertial navigation systems (INS) are dead-reckoning devices that are completely self-contained. They are independent of theiroperating environment, such as wind, visibility,or aircraft attitude. They do not radiate or receiveRF energy; therefore, they are impervious tocountermeasures. They make use of the physicallaws of motion that Newton described threecenturies ago. Of course, you must enter thestarting position into the system. When knownpositions are available, you may correct orupdate the system if an error exists.

Another input to the system is from accelera-tion detectors that measure the rate of change inthe motion of the aircraft. The first integral ofacceleration is velocity. Velocity results whenacceleration is integrated with respect to time. Forexample, a body starts from rest and constantlyaccelerates at 8 feet per second per second for 11seconds. The velocity at the end of this time wouldbe 88 feet per second (60 miles per hour).However, in actual practice, acceleration is notalways this constant. The integration of accelera-tion is the process of summing all minuteacceleration-time increments over a given amountof time.

By integrating velocity with respect to time,the result is displacement (distance). Therefore,the second integral of acceleration is displacement.The inertial navigator’s purpose is to keep trackof position and not the total distance traveled.This causes the system to integrate all values ofacceleration (positive and negative) detected overthe time involved.

If the earth were flat and vehicles traveled onlyon the earth’s surface, a two-axes inertial naviga-tion system could plot the position using twoaccelerometers. One accelerometer would besensitive along the x-axis (E-W) and the othersensitive along the y-axis (N-S). The importantpoint to note about detecting acceleration of abody is that each accelerometer detects only thecomponent of the resultant acceleration along itssensitive axis. They have no way of telling whether

the detected velocity change is due to a speedchange or a direction change or both. It does notmatter what forces cause the velocity change.Neither can the accelerometer distinguish betweenthe acceleration of the vehicle and the pull ofgravity. Therefore, if the accelerometer tilts offits level, its output will include a component ofgravity as well as vehicle acceleration. To get thecorrect vehicle acceleration in the horizontalplane, the sensitive axis of the accelerometer mustbe perpendicular to the gravitational field.

Of course, the earth is not flat, and notexactly round either. Its radius at the poles is lessthan its radius at the equator. It also spins aboutits polar axis. A spinning gyro in gimbals tries tomaintain a fixed direction in relation to spacerather than to any point on earth. Consider a gyroat the equator with its spin vector direction east,toward the morning sun. After 6 hours of earthrotation, the spin vector would be up in relationto the earth’s surface. After 12 hours it would bewest. After 18 hours it would be down; and after24 hours it would be east again. Also, considera spinning gyro with its spin axis parallel with theearth’s axis of rotation. At the equator it isparallel with the earth’s surface. However, as itmoves to the North Pole, it becomes vertical tothe earth’s surface. You must take all of theseitems into account and correct for them. So, tonavigate on the earth requires a highly complexinertial system, each component of which iscapable of extreme accuracy.

A four-gimbal system allows the platform toretain the original orientation regardless of whatmaneuvers the aircraft makes. This allows theplatform to serve as a level mount for theaccelerometers. The stable platform contains twoidentical floated, two-degree-of-freedom gyros.They mount with their spin axes horizontal andat right angles to each other. Using the gyroscopicprinciple of precession, it is possible to apply acontinuous torque to the appropriate axes. Thisaction reorients the gyros to maintain the stableplatform horizontal to the earth’s surface andpointed north. An electronic analog computerdevelops the signals necessary to properly torquethe gyros. The correct ions for earth rate dependon the aircraft’s position on the earth’s surface.

Some system’s use as many as three acceler-ometers. Two are horizontal with one sensitive tonorth-south acceleration and the other sensitiveto east-west acceleration. The third accelerometermounts to determine vertical acceleration. Acomputer subtracts the gravity component fromthe output of the vertical accelerometer. A more

detailed description of an INS follows later in thechapter.

Q1. A point that is defined by stated or implied.coordinates is known as a

Q2. The intended horizontal direction.of travel

is known as

Q3. In what two reference directions can youexpress bearings?

Q4. The east/west geographical coordinate is.known as

Q5. You measure longitude 180° east or westfrom what point?

Q6. The angle between true north and thedirection of the earth’s magnetic field is

.known as

Q7. How do you label variation?

Q8. Magnetic influences cause what type oferror in magnetic compasses?

Q9. The net result of both variation and.deviation is known as

Q10. When variation and deviation have thesame name, do you add or subtract toobtain compass error?

You can determine a position from therecord of a previously kno~w position,course, speed, and time traveled by whatprocess?

What navigation system makes use of thephysical laws of motion that Newtondescribed three centuries ago?

AIRCRAFT COMPASS SYSTEMS

Learning Objective: Recognize theoperating principles and features of com-pass systems, the attitude reference system,and associated sensors and indicators.

Countless navigational devices and methodshave been invented and devised. In the presentera, with its supersonic speeds, accurate deter-mination of direction has become increasinglyimportant. An error of only a few degrees in a

space of minutes will carry the modern aviatormany miles off course.

During the early days of aviation, directionof flight was determined within the aircraft chieflyby direct-reading magnetic compasses. Today thedirect-reading magnetic compass still finds use asa standby compass should the more sophisticatedcompass systems fail.

COMPASS SYSTEM SENSORSAND INDICATORS

In chapter 6, the heading indicator ismentioned as a flight instrument. This instrumentis part of the primary heading reference system.The heading indicator receives electrical/electronicsignals from various components in the systemand shows the pilot aircraft heading in degrees.Do not confuse this compass system with thestandby compass system (wet compass) coveredin chapter 6.

Sophisticated navigation systems and weaponsdelivery systems require aircraft heading informa-tion in electrical/electronic signal form. Inthis form, the information goes to computers,

Figure 7-12.—Flux valve heading changes.

Figure 7-13.-Magnetic flux line movement.

indicators, and other components. Also, byusing these signals, indicators can include aircraftheading along with other information in a singleinstrument.

Compass Transmitter

The compass transmitter, called a flux valve,detects the direction of the flux lines of themagnetic field of the earth. It electrically sendsthis information to a servo loop. The transmitteris only a few inches in height. It is usuallymounted within the wing or tail of the aircraftas this area has the lowest magnetic disturbancesin the aircraft. It consists of a hermeticallysealed hemispherical bowl containing a penduloussensing element in a damping fluid. A universalmounting permits the sensing element up to30 degrees of freedom in roll and pitch, whileprohibiting rotation about the vertical axis.

The laminated metal core of the flux valve isa good conductor of magnetic lines of force (highpermeability). It is shaped like a three-spokedwheel with the rim split between the spokes, asshown in figure 7-12. The earth’s magnetic linesof force will concentrate in these legs because theyoffer low reluctance to the flux. The amount offlux in any one leg is proportional to the angularposition of the leg relative to the earth’s magneticlines of flux. The signal pickup coils are woundaround these legs, one for each leg. The coils are1200 apart in a horizontal plane. The coilsconnect electrically in a wye configuration.

The exciter coil is wound around the hub ofthe core, corresponding to the axle of a wheel.This coil receives 400-hertz ac power. This coilis the primary, and the signal pickup coil is thesecondary. The design of the core and windingsprevents transformer action between the coils. Thepurpose of the primary winding and its appliedvoltage is to produce a magnetic field. This

magnetic field changes the reluctance of the core.When the 400-hertz ac goes to the primary coil,the current generates a magnetic field in the core.The magnetic field drives the core to saturationat the peak of each positive and each negativeportion of the cycle. Figure 7-13 shows the sideview of the center hub and one leg of the sensingelement.

When the 400-hertz voltage is at zero, thereluctance of the core is low. This allows themaximum number of the earth’s magnetic fluxlines to concentrate in the core. As the primarysine wave builds up toward maximum, reluctanceincreases, making the core less attractive to earth’smagnetic flux lines. As the magnetic flux linesmove out, they cut through the secondary coil andinduce a voltage (emf) in it. Refer to figure 7-14.

Figure 7-14.-FIux valve sine waves.

Figure 7-15.-Compass transmitter and compensator.

The movement of flux lines increases until pointY on the 400-hertz sine wave, and then themovement tapers off and stops moving at pointZ. At this point no voltage is induced into thesecondary winding. As the primary sine wavestarts dropping toward X1, the reluctance of thecore decreases. The core now attracts more andmore of the earth’s flux lines. These lines of fluxcut through the secondary coil in the oppositedirection and induce a voltage of the oppositepolarity. When the primary sine wave reaches X1,the maximum number of the earth’s flux lines

again concentrate in the core. However, they arenot moving, so the induced voltage in thesecondary coils is zero.

The same action takes place during thenegative half of the 400-hertz excitation current.This makes the resulting output of the flux valvean 800-hertz signal. The voltages of each of thethree secondary coils pass through zero at thesame instant. However, their polarity andamplitude depend on the heading of the aircraft.They vary in the same fashion as the stators ofa synchro transmitter. This would allow you touse it to drive an 800-hertz synchro system.The original 400-hertz excitation is effectivelycancelled by the special construction of the metalcore and windings. This construction allows linesof flux produced by the primary coil to inducecanceling voltages in the secondary coils.

Attached to the top of the compass transmitteris a compensator assembly (fig. 7-15). It consistsprimarily of two sets of two small permanent barmagnets. You can change the relative azimuthposition of each set by rotating a screw on theoutside of the unit. These screws position themagnets by a gear train. One adjusting screwadjusts for north-south compensation, and theother screw adjusts for east-west compensation.

Two wiring symbols for the flux valve areshown in figure 7-16. To distinguish them fromsynchro units, the words compass transmitter orflux valve are usually included in the drawing.

Displacement Gyroscope Assembly

The displacement gyroscope assembly (fig.7-17) consists of a vertical gyro (VG) and a

Figure 7-16.-Compass transmitter schematic and functional symbols.

Figure 7-17.—Typical displacement gyro and servo loops.

directional gyro (DG). These gyros mount in acommon outer roll gimbal. The vertical gyroprovides pitch and roll signals and the directionalgyro provides heading (azimuth) signals. Erectionand leveling servo loops erect and maintain thespin axis of the vertical gyro gravity vertical andthe directional gyro spin axis parallel to the earth.Roll, pitch, and azimuth control transmittersconvert aircraft attitude and heading into electricalsignals.

VERTICAL GYROSCOPE OPERATION. —Look at figure 7-17 as you read this section. Thevertical gyro consists of gyro spin motor B101(which is the inner roll gimbal) and the verticalgyro pitch gimbal. It also includes the outer rollgimbal and the frame. The frame mounts to theassembly case and follows all aircraft maneuvers.The outer roll gimbal mounts in the frame. It may

rotate 360° about the roll axis but follows the air-craft in pitch and yaw.

The vertical gyro pitch gimbal mounts in theouter roll gimbal. This gimbal may rotate 360°about the pitch axis but follows the outer rollgimbal movements in roll and yaw. The gyro spinmotor may rotate ±85° in roll but follows thevertical gyro pitch gimbal in pitch and yaw.Mechanical stops (not shown) limit inner rollgimbal movement to prevent B101’s spin axis,aligning with the vertical gyro pitch gimbal axis.Such an alignment would cause the vertical gyropitch gimbal to spin about its pitch axis (gimballock).

Leveling. —At power application, a frictionbrake (snubber) releases the outer roll gimbal fromthe frame. The gyro spin motor starts, and

electrolytic switches sense unlevel conditions inpitch and roll. The output of the electrolyticswitches activates the torquers. The gyro reactsto the applied torque and precesses until theelectrolytic switches are level. The inner rollcontrol transmitter is mounted between thevertical gyro pitch gimbal and the inner rollgimbal. This transmitter applies signals to the rollservo amplifier to drive the roll motor-generator.The roll motor-generator, in turn, drives the outerroll gimbal to the level of the inner roll gimbal.

Pitch Sensing. —As the aircraft pitches, theouter roll gimbal follows, but the vertical gyropitch gimbal remains level. The pitch servocontrol transmitter detects the pitch attitude andapplies pitch signals to the indicators and otheraircraft systems. The pitch control transmitterapplies pitch signals to the automatic flightcontrol system (AFCS) control amplifier.

Roll Sensing. —As the aircraft rolls, B101remains level, but the vertical gyro pitch gimbalrolls (with the outer roll gimbal). The inner rollcontrol transmitter senses the difference. It thencauses the roll amplifier to drive the roll motor-generator until the outer roll gimbal is level withB101. The outer roll control transmitter (on frontend of frame) detects and applies roll signals toindicators and other systems. The roll controltransmitter (on aft end of frame) applies rollsignals to the AFCS control amplifier.

DIRECTIONAL GYROSCOPE OPERA-TION. —The directional gyro consists of gyro spinmotor B201 (including a leveling gimbal), anazimuth gimbal, and a directional gyro pitchgimbal. The directional gyro pitch gimbal mountsin the outer roll gimbal. The pitch gimbal maymove 3600 about the pitch axis, but it follows theouter roll gimbal in roll and yaw. The azimuthgimbal mounts in the directional gyro pitchgimbal. It may move 3600 about the yaw axis;however, it follows the directional gyro pitchgimbal in pitch and roll. B201 mounts in theazimuth gimbal. B201 is limited to ±85° bymechanical stops (not shown) to prevent gimballock.

Leveling. —The leveling control transmitteroutput goes to the leveling amplifier, which drivesthe leveling torquer. When the leveling torquermoves the azimuth gimbal, B201 precesses untilthe leveling control transmitter senses a levelcondition. The directional gyro pitch gimbal isservoed to the vertical gyro pitch gimbal and

maintained perpendicular to the surface of theearth. The pitch servo control transmitter output,through the pitch servo control transformer, isamplified and drives the pitch follow-up motor-generator. The motor-generator positions thedirectional gyro pitch gimbal.

Azimuth Sensing. —The azimuth gimbal maysettle at any random position in yaw. The onlyforces acting on the gimbal are gyro rigidity,apparent (earth rate) precession, and the levelingtorquer. Azimuth sensing in the directional gyrooperating mode is reliable only after setting thecorrect heading into the system with the SETHDG control. Two azimuth control transmitterssense any movement of the directional gyro pitchgimbal about the azimuth gimbal. The yaw signalof one azimuth control transmitter goes to theattitude indicator. The yaw signal of the otherazimuth control transmitter is processed in thecompass adapter-compensator and applied toother aircraft systems.

Horizontal Situation Indicator (HSI)

Aircraft, such as the P-3, use the horizontalsituation indicator to provide the pilot with a

Figure 7-18.-Horizontal situation indicator.

visual indication of the navigational situation ofthe aircraft. The horizontal situation indicator(fig. 7-18) is on the pilot’s main instrument panel.It provides a visual presentation of the horizon-tal or plane view of the aircraft relative to thenavigation situation. It also provides an integrateddisplay of navigation data from various sources.It presents this data to the pilot in a symbolicpictorial display for quick and easy assimilation.The center portion of the display contains anazimuth or compass card (callout 6). The carddisplays aircraft magnetic heading when readagainst the top of the lubber line (callout 8). Youmay also read the bearing pointer (callout 7),course pointer (callout 10), and command headingmarker (callout 9) against the card.

The bearing pointer (outer pointer, shown at2200, provides pictorial bearing information toa ground electronic station. It can also providea bearing to a base or target (as computed by thenavigational computer set). The course pointer(inside compass card, at 3000, shows the selectedcourse to a ground electronic station, or aircraftmagnetic ground track. The course deviation bar(callout 5) (center segment of this pointer) showsthe aircraft’s deviation from a selected course. Itshows this pictorially with respect to the stationaryminiature aircraft symbol (callout 4) at the centerof the display.

A to-from pointer (callout 11) (just above theaircraft symbol) shows if the selected course leadsto or from a ground electronic station. Acommand heading marker (callout 9) just outsidethe compass card (shown at 3000, shows thecommand magnetic heading. It also shows, by itsangular displacement from the lubber line, theheading error angle. The course pointer and thecommand heading marker may be set manuallyby means of the COURSE SET (callout 12) andHEADING SET (callout 2) knobs. They can alsobe set remotely by external signals to the HSIcourse command and heading command servos.The selected course to a ground electronic stationor aircraft magnetic ground track is also indicatedon the course counter (callout 13). A distancecounter (callout 1) at the lower left shows thedistance in nautical miles. The distance may beto the ground electronic station, base, or target(as computed by the navigational computer set).

There are four mode-of-operation lights(callout 3) (TAC, UHF, MAN, NAV) that showaround the display. They illuminate internallyto show the selected operating modes. Theunilluminated words remain practically invisible.The instrument uses integral lighting.

For detailed operation of the horizontalsituation indicator system, refer to the applicablemaintenance instructions manual.

Bearing-Distance-HeadingIndicator (BDHI)

The BDHI is used with various navigationsystems and provides information according tothe mode selected. Some aircraft have more thanone BDHI, with separate select switches for eachinstrument. The distance counter numerals maybe in a vertical row or horizontal, as theID-663A/U in figure 7-19.

The lubber index is a fixed reference mark atthe top of the instrument face. The compass card(read under the lubber index) shows the aircraftheading (either true or magnetic, depending onthe mode used). Two pointers, a single bar anda double bar, can indicate the following:

Bearing to a ground electronic station

Bearing to destination

Aircraft ground track

Aircraft drift angle

Heading error

The BDHI select switch selects the availablecombinations of these indications in a givenaircraft configuration.

Figure 7-19.-Bearing-distance-heading indicator.

Figure 7-20.-Block diagram of the attitude heading reference system.

Electrical inputs for the compass card andpointers come from synchro transmitters locatedin other equipment. In the ID-663A/U, synchrotorque receivers position the compass card andboth pointers, In the ID-663 B/U and ID-663B/V,torque receivers position both pointers. Thesetorque receivers contain a synchro controltransformer, a transistor error signal amplifier,and a two-phase servomotor to position thecompass card. In the ID-663C/U, the controltransformer-amplifier srvomotor positions thecompass card and both pointers.

The distance counter may display distanceto base, target, or ground electronic station,depending on the mode selected. It consists ofthree synchro torque receivers to position theunits, tens, and hundred numerals. It also has a

1,000 flag to place the numeral 1 preceding thehundreds numeral. This enables the counter todisplay distance up to 1,999 nautical miles.

The OFF flag covers the distance counterwhen distance information is not available. Inthe ID-663B/V, the 1,000 flag and OFF flag arepositioned by separate coils, The ID-663A/U,ID-663B/U, and ID-663 C/U have a meter-typemovement, which, when de-energized, showsOFF. Partial rotation moves the OFF flag out ofview, and full rotation brings the 1,000 flag intoview.

ATTITUDE HEADINGREFERENCE SYSTEM (AHRS)

The AN/ASN-50 attitude heading referencesystem (fig. 7-20) generates and provides

continuous roll, pitch, and heading signals. Thesesignals go to the aircraft attitude indicator andother avionics equipment. Error signals developin the displacement gyroscope as a result ofdisplacement of synchro sensing devices from theirnull position. A remote compass transmittersupplies additional heading information to thesystem. For detailed information on theAN/ASN-50 system, you should refer toReference Al t i tude Heading, N A V A I R05-35LAA-1.

In the attitude heading reference system, rolland pitch information goes directly from thevertical gyroscope to the attitude indicator. Thesystem supplies azimuth information in one ofthree modes selected on the compass controller—COMP (compass), SLAVE (slaved), or FREE(free).

The compass mode is an emergency modeused when the displacement gyroscope is mal-functioning. When operating in the compassmode, the magnetic heading signal from the fluxvalve integrates with a compensating signal. Thisintegration corrects for magnetic field distortionaround the aircraft.

The free mode is used at latitudes greater than700. You can also use it in areas where the earth’smagnetic field has appreciable distortion, andwhen the flux valve fails. When operating in thefree mode, you manually set the initial headingreference, using the compass controller PUSH TOTURN control. After establishing the initialreference, heading information accuracy dependson the directional gyroscope’s ability to maintainits position relative to earth.

The slaved mode is the normal operatingmode. This mode uses both the flux valve anddirectional gyroscope signals. The directionalgyroscope signal acts as a stabilizer for the fluxvalve data. Automatic fast synchronization to themagnetic heading occurs during the start cycle andupon switching to the slaved mode. A SYNC INDmeter on the compass controller shows when aloss of synchronization occurs. The loss ofsynchronization occurs between the azimuthheading output and the flux valve. When such acondition exists, you may accomplish fastsynchronization by momentarily pressing thecompass controller PUSH TO SYNC switch.

Autopilot decoupling occurs whenever manualor automatic synchronization occurs. Circuitrycompensates for apparent drift of the gyroscoperesulting from rotation of the earth. Othercircuitry provides means to interrupt slaving and

Figure 7-21.-Displacement gyroscope.

erection voltages during turns of 15 degrees perminute or greater.

Displacement Gyroscope

The displacement gyroscope (fig. 7-21) is ahermetically sealed, two-gyroscope platformproviding pitch, roll, and azimuth signals to thesystem indicating instruments. The verticalgyroscope provides a source of pitch and rollinformation, while the directional gyroscopeprovides azimuth (yaw, heading) information.Control transmitters (synchros) convert attitudechanges into electrical signals that represent pitch,roll, and yaw. (The AN/ASN-50 displacementgyro is similar to the one discussed earlier in thischapter.)

Erection circuits maintain the spin axis of thevertical gyroscope in a gravity vertical position.These circuits consist of a roll electrolytic switch,with associated roll torquer, and a pitchelectrolytic switch, with associated pitch torquer.

A servo loop maintains the spin axis of thedirectional gyroscope level. The loop consists ofa leveling pickoff, an external leveling amplifier,and a leveling torquer.

The vertical gyroscope spin motor mounts inthe inner roll gimbal, which is free to move in aroll direction. The freedom of movement in theroll direction is limited to ±82° by mechanicalstops to prevent gyroscope gimbal lock. Gimballock would occur if the vertical gyroscope spinaxis were to become aligned with the verticalgyroscope pitch gimbal. The directional spinmotor mounts in the directional gyroscope

leveling gimbal. This gimbal is free to move ina roll direction. Mechanical stops also restrictthe freedom of the leveling gimbal to ±82°,preventing gimbal lock. These stops also preventthe inversion of the leveling and inner rollgimbal during gyroscope rotor coastdown aftersystem shutdown.

The outer roll gimbal is the outermost gimbalfor both the vertical and directional gyroscopes.Pitch control transmitters detect pitch movementof the outer roll gimbal about the verticalgyroscope pitch gimbal. The outer roll controltransmitter and directional gyroscope controltransmitter mount between the outer roll gimbaland gyroscope case (frame). However, they mountat opposite ends. They sense roll movement of theaircraft (and frame) about the outer roll gimbal.

The azimuth gimbal may settle at any randomheading during power application. Therefore, youmust slave the directional gyroscope’s initialazimuth signal to the flux valve. Also, you cancorrect it manually to a known magnetic heading.You must do this so the azimuth signal reflectsactual aircraft heading. The two azimuth controltransmitters, between the azimuth gimbal and thedirectional gyroscope pitch gimbal, will thenfurnish information on any further change inaircraft heading.

The displacement gyroscope incorporatessnubbers. They maintain the approximate normalposition of the outer roll gimbal and thedirectional gyroscope pitch gimbal when poweris removed. Upon application of power, thesnubbers energize, removing their snubbingaction. The motor-generator and gear assembliesthat drive the gyroscope gimbals have enoughpower to override the snubbing action should asnubber failure occur.

Amplifier Power Supply

The amplifier power supply (fig. 7-22) containsthe following components:

• An ac and a dc power filter

• Two power supply modules

• A roll driver amplifier

• Roll, pitch, and leveling servo amplifiers

• A leveling modulator

• A servo failure monitor module

• Two thermal relays

• Ten control relays

Figure 7-22.-Amplifier power supply.

The amplifier power supply provides the timing,switching, and voltages required for the startcycle; erection voltage control during systemoperation; monitoring roll and pitch signals andthe neutral power lead of the vertical gyroscopemotor; and the roll, pitch, and leveling amplifiersrequired for control of the displacement gyroscopegimbals.

Aircraft three-phase power goes to theamplifier power supply. It then routes through theac power filter. The filtered three-phase powerthen goes to other system components. A powersupply connected to all three legs of the three-phase power provides 28-volt dc, which goes toa filter. System components use both filtered andunfiltered dc.

The second power supply is a three-sectionsupply. Each section is independent of the otherand supplies 95 volts dc. The dc outputs connectto various control relays within the amplifierpower supply.

The output from the displacement gyroscopephotoelectric pickoff is a dc voltage. A levelingmodulator converts it to a 400-hertz ac voltage.The amplitude and phase of the modulator out-put depends on the amplitude and polarity of thedc voltage from the photoelectric pickoff. The400-hertz ac voltage goes to the leveling amplifier;then the amplified signal goes to the directionalgyroscope leveling torquer control winding.

Roll and pitch error signals from the displace-ment gyroscope are amplified by the roll andpitch amplifiers. After amplification, the signalsgo to the displacement gyroscope for applicationto the roll and pitch motor-generator controlwindings. The roll and pitch motor-generators

Figure 7-23.-Compass adapter compensator.

drive the displacement gyroscope roll and pitchgimbals.

Two thermal relays and nine control relaysperform the timing and switching required for thestart cycle. The start cycle is of 60 secondsduration. However, certain conditions changeafter the first 12 seconds. High pitch erectionvoltage goes to the pitch torquer for the completestart cycle. High roll erection voltage goes to theroll torquer after 12 seconds and continues untilthe completion of the start cycle. After the first12 seconds, motor excitation voltage increases,and the gyroscope motors attain operating speed.After completion of the start cycle, two additionalrelays provide roll and pitch erection cutoutduring specific aircraft maneuvers.

Compass Adapter Compensator

The compass adapter compensator (fig. 7-23)receives heading information from the flux valveand the displacement gyroscope. It processes thisinformation according to the azimuth modeselected by the compass controller. It thenprovides corrected heading signals to the headingindicator and other aircraft systems. The compassadapter compensator can operate in the followingmodes: free, compass, and slaved.

In the free mode of operation, you engage thePUSH TO TURN control on the compasscent roller to set actual aircraft heading. Thisestablishes an initial azimuth reference. As air-craft heading changes, an azimuth synchro on thedirectional gyro measures the relative changebetween the aircraft and the directional gyroscope.The signal goes to the compass adapter, whichmakes corrections for real and apparent drift. Thecompensating signal for apparent drift is derived

from a resolver in the compass controller. TheEARTH RATE CAL variable resistor in thecompass adapter adjusts this signal. Real driftis caused by mechanical imperfections inconstruction of the directional gyroscope. Thecompensating signal for real drift develops acrossthe compass adapter GYRO DRIFT COMPEN-SATION POT variable resistor. This resistor hasa dial calibrated in degrees per hour. The correctedinformation goes to external aircraft system com-ponents through five heading repeater synchros.

When operating in the compass mode, the24-point compensation network corrects the fluxvalve signal for deviations. This signal goes as anerror signal to the azimuth servo loop, resultingin corrected azimuth information (angle datashaft). Again, this information goes to externalaircraft system components through five headingrepeater synchros.

In the slaved mode, the directional gyroscopeazimuth information and a compensated fluxvalve heading correction signal go to a differentialsynchro in the compass adapter. This slaves orsynchronizes the gyroscope azimuth output to theflux valve heading, providing a heading outputrather than a displacement output. Fast syn-chronization starts when the slaved mode isselected and is maintained until close alignmentis achieved. Slow synchronization is appliedcontinuously while operating in the slaved mode.

Compass Controller

The compass controller (fig. 7-24) providesswitching functions, latitude compensationsignals, and slew signals to the system compass

Figure 7-24.-Compass controller.

adapter. The compass controller also monitorsand provides a visual display of the synchroniza-tion between the azimuth heading output and theflux valve.

The compass controller can operate in threemodes—slaved, free, and compass. The compasscontroller contains a PUSH TO SYNC switch andSYNC IND meter. It also has a PUSH TO TURNcontrol, mode switch (with COMP, SLAVE andFREE positions), and LATITUDE DEGREEScontrol (counter assembly).

The mode switch selects the FREE (free),SLAVE (slaved), and COMP (compass) modesof operation. To accomplish this, it controls themode-selecting relays in the system compassadapter. The mode switch also activates the SYNCIND meter during compass and slaved modes.

The PUSH TO TURN control (set heading)provides switching to decouple the autopilot.Also, it controls the direction and rate ofslewing for alignment of the system to an azimuthheading. This switching action occurs whenusing the PUSH TO TURN control to referencethe system output to the aircraft heading in thefree mode.

The PUSH TO SYNC switch provides switch-ing to synchronize azimuth heading outputto the flux valve when operating in the slavedmode.

The LATITUDE DEGREES control, workingwith a resolver, provides a compensation signalfor apparent drift of the directional gyroscope.Apparent drift results from earth rotation.

The hemisphere switch (N,S) selects thelatitude correction signal for the Northern orSouthern Hemisphere.

The SYNC IND meter shows the synchroniza-tion between the azimuth heading output and theflux valve.

The slaved mode is the normal operating modeexcept when in an area where the earth’s magneticfield is distorted. The slaved mode synchronizesthe directional gyroscope output to the flux valveheading. When initiating the slaved mode, fastsynchronization occurs until close alignment withthe flux valve heading is achieved.

Free mode is an alternate mode of operation.It is used in areas where the earth’s magnetic fieldis distorted. When operating in the free mode,only the directional gyroscope output drives thesystem’s azimuth indicators.

The compass mode is an emergency mode foruse when the directional gyroscope fails. Only theflux valve output (compensated) provides headinginformation.

Figure 7-25.-Switching rate gyroscope.

Rate Gyroscope

The switching rate gyroscope (fig. 7-25)provides a means of interrupting the rollerection and slaving voltage. It is a single-degree-of-freedom gyroscope. It providesgyroscope sensitivity to rates of rotation aboutthe yaw axis of the aircraft. When the air-craft turns at rates of 150 per minute orgreater, the gyroscope precesses away fromthe normal condition. This causes the contactsof a magnetic reed switch to close, energizinga single-stage amplifier. Transistor switchingaction pulls in a relay, completing the 28-volt dcpath to the turn cutout relay coil in the compassadapter.

Attitude Indicator

The attitude indicator (fig. 7-26) is a three-axis,servo-driven sphere that shows heading andrelative roll and pitch attitude of aircraft.Vertical and horizontal pointers provide the pilotwith aircraft deviation information from a desiredflight path. Signals for operating servo systemsof hermetically sealed units are from the displace-ment gyroscope. Aircraft rate of turn informa-tion and vertical displacement deviations from adesired glide path are displayed by a rate of turnpointer and displacement pointer, respectively.Loss of ac power is indicated by a display of apower failure warning flag. Inadequate currentto vertical, horizontal, and displacement pointersresults in display of respective warning flags. Apitch trim knob lets you adjust the sphere tovarying aircraft configurations and reduceparallax error.

Figure 7-26.—Attitude direction indicator.

What is the purpose of the four-gimbalsystem in inertial navigation systems?

When operating the AN/ASN-50 attitudeheading reference system, when do you usethe compass mode?

INERTIAL NAVIGATION SYSTEM

Learning Objective: Recognize theoperating principles and characteristics ofthe inertial navigation system, to includeSchuler loops and tuning; and identifynavigation errors and aligning and calibra-tion procedures.

The following paragraphs describe a headingreference system found on the latest model high-speed aircraft and some patrol aircraft. Theinertial navigation system (INS) is sometimesmaintained by the Aviation Electronics Technician(AT) rating. Some squadrons use a concept calledan integrated weapons team (IWT). It is composedof the three avionics/armament division (workcenter 200) ratings—AT, AO, and AE. Regardlessof who maintains the INS, you, must be familiarwith the theory and operating principles of sucha system.

You may define navigation as the process ofdirecting a vehicle from one point to another.Basically, navigation can be divided into twocategories: (1) position fixing and (2) deadreckoning. In the first category, you determineyour position relative to positions of knownobjects such as stars and landmarks. The mostcommon example of navigation by positionfixing is celestial navigation. Use of loran isanother example of navigation by periodicposition fixes. Dead reckoning, the secondcategory, is the process of estimating yourposition from the following known information:

l Previous position

l Course

l Speed

. Time elapsed

Two examples of navigation by dead reckoningare Doppler radar and inertial navigation systems.

All navigation systems, except the inertialtype, rely on some bit of information external tothe vehicle to solve its navigational problem. Inthis respect, the inertial navigation system standsalone; it is completely self-contained within thevehicle. It is independent of its operating environ-ment, such as wind, visibility, or aircraft attitude.It does not radiate RF energy; therefore, it isimpervious to countermeasures. It does notdepend on ground transmission or any otheroutside source to determine its instantaneousposition. The inertial navigation system simplymakes use of the physical laws of motion thatNewton described three centuries ago.

BASIC PRINCIPLES

The operating principle of the inertialnavigation system (INS) is Newton’s first law ofmotion. This law states “Every body continuesin its state of rest, or of uniform motion in astraight line, unless it is compelled to change thatstate by forces impressed on it.” In laymansterms, this law says that a body at rest tends toremain at rest. It also says a body in motion tendsto remain in motion, unless acted upon by an out-side force.

The full meaning of Newton’s first law is noteasy to visualize in the earth’s reference framebecause Newton’s laws apply to an inertialreference system. You may define an inertial

reference system as a nonrotating coordinateframe. It can be either stationary or movinglinearly at a uniform speed, in which there are noinherent forces such as gravity.

You can make a simple test of whether youare in a true inertial system by releasing anobject and observing its motion. If you releasethe object without imparting any acceleration toit, the object remains in its position relative toyou. If you throw the object, it continues on anundeviating path at a constant speed. Such asystem can exist only in empty space, far fromany mass, for all masses contain gravitationalforces. A reference system attached to the earthcan closely approximate an inertial system. Forthis system to work, you must balance thegravitational force on a body by a second force.For example, an object sliding on a flat,frictionless plane on the earth’s surface wouldmove in a NEARLY straight line. The object willhave a NEARLY constant speed, as you saw inthe earth’s coordinate system.

NOTE: The word nearly is emphasizedbecause the object will deviate slightly from itsstraight-line motion. The cause of this deviationis the earth’s rotation about its axis.

Newton’s second law of motion sharesimportance with his first law in the inertialnavigation system because the inertial navigationsystem works on Newton’s second law. Newton’ssecond law of motion states “Acceleration isproportional to the resultant force and is in thesame direction as this force.” Thus, the secondlaw is written

The physical quality in the above equation thatpertains to the inertial navigation system isacceleration. You can derive velocity anddisplacement from acceleration. For example,consider this fact: Before an object can changeits state of rest or state of motion, it must firstexperience an acceleration. Since acceleration isa change in velocity and velocity is a change in

position, acceleration is a change in the changeof position. However, before any change can havemeaning, it must include the unit of time.Therefore, you can define a change per unit oftime as a rate of change. Thus, a rate of changeof displacement is velocity. A rate of change ofvelocity is acceleration. A rate of change of a rateof change of displacement is acceleration.

Differentiation is the process of investigatingor comparing how one physical property varieswith respect to another. Integration, the reverseof differentiation, is the process of summing allrate of changes that occur within the limits underinvestigation.

The inertial navigation system is not adifferentiating system; it is an integrating system.However, before integration can be done, it mustfirst have a rate of change. Therefore, theinertial navigation system, when stripped to itsbarest essentials, is a detector and an integrator.It first detects changes of motion. It thenintegrates these changes of motion with time toarrive at velocity, and again with time to arriveat displacement.

Fundamentals of Integration

Since an INS performs integration, the follow-ing is a review of integrating principles.

The equations for the integrals of accelerationand velocity are —

When acceleration (a) is integrated overa specific period of time (dt), the result is velocity When velocity (v) is integrated over a specific period of time (dt), the result isdisplacement (s). Therefore, when acceleration (a)is integrated twice over a specific period oftime (dt2), the result is displacement (s).

Figure 7-27.-Simple single-axis INS block diagram.

Remember from elementary physics thatacceleration, whose units are ft/sec2, multipliedby time in seconds is velocity in ft/sec. Also, thatvelocity (ft/sec) multiplied by time (sec) isdisplacement (ft). The integration of acceleration,for example, is the mathematical process ofsumming all minute acceleration-time incrementsover a given period. The result of the integrationof acceleration is velocity over the same period.The same integration process performed onvelocity gives displacement or distance traveledover the same period.

Simple Single-Axis InertialNavigation System

An example of how a simple single-axis INSoperates is illustrated as follows: Assume aperson is on an INS-equipped train on railroadtracks at the equator. The tracks run in a straightline east and west only, The INS consists of anacceleration detecting device (accelerometer), anintegrating device, and a displacement readoutdevice. The accelerometer can sense movement inonly one direction, along its sensitive axis. Thesensitive axis is an imaginary line parallel to themovement of the mass within the detecting device.The acceleration detecting device is oriented in thetrain so that it detects accelerations when the trainis moving forward or backward. Figure 7-27 isa block diagram of such a device.

If the train starts moving at point A, you willnote a specific reading on the displacementreadout device. When the train reaches point Band stops, the readout device will show the newposition. The distance traveled from point A topoint B added to the reference value noted at pointA will show on the displacement indicator. The

Figure 7-28.-Integration of acceleration and velocity:(A) acceleration, (B) velocity, (C) displacement.

train returns to point A by traveling backwards.Thus, the simple inertial device is not disoriented.At point A the readout device shows the value thatwas chosen as a reference. This is the displace-ment at point B minus the distance traveled frompoint B to point A.

Figure 7-28, view A, is a graph of the detectedacceleration. View B is the velocity curve obtainedby integrating the acceleration curve shown inview A. View C is the displacement curve obtainedby integrating the velocity curve shown in viewB. All three curves are plotted as a function oftime.

The acceleration curve (fig. 7-28, view A)begins at time to as the train begins to travel frompoint A. Look at view C. The acceleration at timeto has a value of al, and it remains at that valueuntil t1. At t1 the train ceases to accelerate.Therefore, acceleration goes to zero. At this point,the train reaches a steady velocity. The traincontinues traveling at a constant velocity untiltime t2 where the train begins to stop. Theacceleration detector detects an acceleration equalin value to a1, but its direction is opposite. Thisacceleration is constant from time t2 to time t3.At t3 the acceleration goes to zero. The train isnow stationary and standing at its destination—point B.

Look at the velocity curve for the timeinterval to to t3 (fig. 7-28, view B). It is the resultobtained when acceleration is integrated over thesame interval. The velocity curve is the output ofthe first integrator from to to t3. During theinterval t o to t1, velocity is changing in anincreasing or positive direction. This means thata positive acceleration is taking place. Velocity isconstant during interval tl to t2, which meansacceleration is zero. At time t2, velocity beginsto decrease. This says that an acceleration is againtaking place. In this case the acceleration isnegative. At time t3, both acceleration andvelocity are zero.

The purpose of an INS is to keep track ofposition and not total distance traveled. To dothis it integrates all values of acceleration (positiveand negative) detected over the interval. There-fore, it is the net value of acceleration thatinterests the INS. For instance, in the interval to

to t3, all accelerations that occur over theinterval are summed, giving a net value at timet3. In this case, integration of acceleration (fig.7-28, view A) is the process of summing the areabounded by the acceleration curve and the timeaxis. The area above the time axis is positive,and the area below the time axis is negative. Sincethe areas above and below the time axis are equal,the net value for interval to to t3 acceleration iszero. The integral of acceleration for the intervalto to t3 is therefore zero. This means that thevelocity at time t3 is equal to the velocity at timeto, in this case zero.

Integrating velocity from time tO to t3 i sthe job of the second integrator. It givesB units of displacement on the displacement axisat time t3. The displacement readout devicechanges continuously as long as the second

integrator produces an output. The secondintegrator ceases to produce an output when thefirst integrator (velocity) ceases to produce anoutput. The velocity integrator continues toproduce until receiving an acceleration thatbalances out the initial acceleration. At this point,it produces a net acceleration of zero. The readoutdevice stops at the point where the net accelera-tion is zero. Until reaching this condition, thereadout device shows a continuous change indisplacement.

The return trip is described as follows. Thetrain is at point B during time interval t3 to t4;it begins traveling backwards to point A at timet4. The acceleration detector senses an accelera-tion –a2, which is negative and slightly less thanthe previous acceleration, –a1. At time t5 i treaches a steady velocity, and acceleration goesto zero at this point. Note that velocity is nownegative since the direction of travel is reversed.Since the size of acceleration –a2 is less than thatof al, maximum velocity on the return trip is less.Therefore, the time required to return to pointA is greater. Interval to to t7, greater than timeinterval to to t3 reflects this fact. The train beginsto stop within a short distance of point A. Thishappens at time t6, producing an acceleration ofa2 as sensed by the acceleration detector. The traincomes to a full stop at time t7. Here the detectorsenses zero acceleration. Since the net accelera-tion over the interval is again zero, the output ofthe first integrator (velocity) is zero. The secondintegrator (displacement) output stops with thedisplacement readout device showing the referencevalue. This value is the same originally noted atreference point A.

The simple single-axis INS just described willdetect and compute all changes in displacement.However, the acceleration detector (accelerometer)must retain its straight-line orientation. Also, all

Figure 7-29.-Two-axis inertial navigation system, block diagram.

Figure 7-30.-Two-axis inertial platform in a plane coordinate system.

motion must be along a straight line passing axis lies along the x-axis. The other accelerometerthrough the reference or initial point. Obviously;using this simple INS, a person must navigatealong a straight line.

Two-Axis Inertial Navigation System

Suppose, for example, that the earth is flat.If so, you can determine position by using asystem of coordinate axes. This system ofcoordinates uses two sets of parallel lines(x and y). One set of lines is perpendicular to theother set of lines. These lines form a grid networkover the earth’s surface.

If you use two single-axis inertial navigationsystems, you can determine position on the plane(flat surface). You simply maintain properorientation of each accelerometer’s sensitive axisrelative to the coordinate system. One acceler-ometer mounts on a platform so that its sensitive

mounts on the same platform so its sensitive axislies along the y-axis. This will maintain their axesmutually perpendicular. The accelerometers willthen sense any rate of change of velocity alongthe coordinate axes. Figure 7-29 is a blockdiagram of a simplified two-axis inertial naviga-tion system.

Figure 7-30 is an illustration of the inertialplatform mounted on a vehicle moving over aplane coordinate system. Note that the platformand accelerometers remain oriented with thecoordinate axes regardless of vehicle heading. Thevehicle’s ground track represents the vehicle’sdisplacement over the grid system. You can locatethe vehicle at any given time by the x and ycoordinates. You plot the x-displacement left toright, and the y-displacement is top to bottom onthe page. You reference time to the x-axis.

Figure 7-31.-Acceleration and velocity curves.

Figure 7-31 is an illustration of a typical setof acceleration and velocity curves from the INSshown in figure 7-30.

Referring to figures 7-30 and 7-31, theoperation of the plane inertial navigation systemis explained as follows: The vehicle aligns(initializes) on the coordinate system with adisplacement of 3 on the x and y axes. That is,both x and y displacement indicators read 3. Attime t3, the vehicle experiences an acceleration,A, in a direction of 45° from the x-axis. The ac-celerometers detect only that portion of theacceleration that lies along its sensitive axis. Thismeans the x-accelerometer detects the componentof acceleration along the x-axis. This is A cos a.The y-accelerometer detects the component of Aalong the y-axis, which is A sin a. The vehiclecontinues in a direction of 45° until time t15. Atthis time it begins a turn to the right. Since sineand cosine are equal at an angle of 450, thedisplacements along x and y are equal at time t15.This says x = 15 and y = 15. Also, it means theacceleration and velocity along the x-axis is equalto the acceleration and velocity along the y-axis.

At time tl5, the vehicle begins a right turn, itcompletes the turn at time t20. The new directionis parallel to the x-coordinate and perpendicular

to the y-coordinate. By looking at figure 7-31, yousee interval t15 to t20 shows the x-accelerometerdetecting a positive acceleration. It also shows they-accelerometer detecting a negative accelerationduring this interval. If the vehicle maintains aconstant speed throughout the turn, the detectedacceleration results from a velocity change. Thischange is due to a change in direction rather thana change in speed. This acceleration is radial(centripetal) acceleration (AR). The direction ofthe radial acceleration is toward the center of theturn and perpendicular to tangential velocity (VT).Find coordinates (17.5, 17) in figure 7-30. If thespeed hadn’t been constant during the turn, atangential acceleration (AT) would have occurred.This acceleration would parallel the tangentialvelocity (VT) vector and be normal (at a rightangle) to the radial acceleration vector (AR). Thedirection of the tangential acceleration woulddepend upon whether the speed was increasingor decreasing, positive or negative. If theturning vehicle’s acceleration is due to changesin speed and direction, the accelerometers detectthe x and y components of the resultant ofthe two accelerations.

Remember, when detecting acceleration of anaccelerating body, accelerometers detect onlythe component of the resultant acceleration alongtheir sensitive axis. Accelerometers can’t tell if thedetected velocity change is due to a speed changeor a direction change or both, Nor does it matterwhat forces cause the velocity changes. The resultis the same, provided the accelerometers maintaincorrespondence with the coordinate axes.

Refer to the acceleration and velocity curvesin figure 7-31. Notice the integration of thex-component of acceleration for the interval t15

to t20. It shows an increase in the x-componentof velocity and, therefore, a correspondingincrease in displacement along the x-axis.Integration of the y-component of accelerationover the same interval shows that the velocity goesto zero at time t20. Therefore, the displacementalong the y-axis ceases to change. Hence at timet15, the displacement is (15, 15). At time t20, thedisplacement is (20, 15) and at time t25, thedisplacement is (25, 15), etc.

The INS just described navigates very well ona flat surface. However, navigation on the earthrequires a highly complex inertial system. Theearth, of course, is not flat, neither is it exactlyround. Its radius at the poles is less than its radiusat the equator. It also spins about its polar axisand orbits around the sun. You must take all ofthese factors into account and correct them

(except the earth’s motion in orbit around thesun). Once you accomplish these corrections, youmay navigate on the earth by inertial means. Theearth’s motion about the sun does not affect anearth inertial navigation system. This motion istranslational, and it is equal at all points on theearth.

SYSTEM COMPONENTS

inertial navigation system continuouslymeasures aircraft accelerations to computeaircraft velocity and change in present position.These measurements are made by precisioninertial devices mounted on a three-axis stableelement, which is part of a four-gimbal structure.The four-gimbal structure allows the stableelement to move with 360 degrees of freedomabout the three axes.

Two gyros provide gimbal stabilization signalsto maintain the stable element level with theearth’s surface and aligned to true north. Also,the system uses these signals to measure aircraftpitch-and-roll attitudes. The inertial characteristicsof the gyroscopes used in the system define andmaintain the reference axes for relatively longperiods with great accuracy. With a gyrostabilizedplatform as a reference, it is possible toaccurately detect components of motion inany direction. To do this, precision accelerometersand analog or digital computers are used in anINS.

Accelerometers

The primary data source for the inertialnavigation system is the accelerometer. Threeaccelerometers are mounted on the stable elementbetween the gyros. They provide output signalsproportional to total accelerations experiencedalong the three axes of the stable element. Thesystem uses these accelerations to produce aircraftvelocities and changes in position.

An accelerometer consists of a pendulous massthat is free to rotate about a pivot axis in theinstrument. Figure 7-32 shows one form of thisdevice. It has an electrical pickoff that convertsthe rotation of the mass about the pivot axis toan output signal. An acceleration of the deviceto the right causes the pendulum to swing to theleft. This provides an electrical pickoff signal,

Figure 7-32.-Typical torque-balanced accelerometer.

which causes a torquer to restrain the pendulum.The pickoff signal goes to a high gain amplifier.The output of the amplifier connects to thetorquer on the accelerometer. During an accelera-tion, this feedback loop sends a voltage to thetorquer. This voltage holds the pickoff signalat a null under the influence of the measuredacceleration. This voltage is proportional to themeasured acceleration. It also provides theelectrical output acceleration signal that goes tothe computer.

The accelerometer (fig. 7-33, view A) cannotdistinguish between the acceleration of the

Figure 7-33.-Principle of an accelerometer: (A) acceler-ometer at null, (B) true acceleration, (C) spuriousacceleration due to gravity.

vehicle and gravitational acceleration. Therefore,if the accelerometer tilts off level, its outputincludes a component of gravitational accelera-tion as well as vehicle acceleration. Look atfigure 7-33, view C. To get the correct vehicleacceleration in the horizontal plane, hold thesensitive axis of the accelerometer normalto the gravitational field. Refer to figure 7-33,view B.

The accelerometer mounts on a platform(stable element) in a way that it is always level.In this position the accelerometer measures trueaircraft acceleration in a horizontal directionalong its sensitive axis. Mounting another levelaccelerometer perpendicular to the first one givesyou the x and y axes. The system can nowdetermine total true acceleration in a horizontalplane for any movement in any direction.

Integrators

To convert the measured acceleration toaircraft position information, the system processesthe acceleration signals to produce velocityinformation. It must then process the velocityinformation to derive distance traveled. Figure7-34 shows an analog type integrator. It is anelectromechanical device that receives electricalinput (acceleration or velocity) and produces ashaft speed proportional to the input. Theshaft angle is the output of the integrator,and it is the mathematical integral of theinput. If the input is acceleration, the outputis velocity; if the input is velocity, the output isdistance.

If one of the horizontal accelerometers pointsnorth, the other one will always point east.By connecting the accelerometer outputs to

Figure 7-34.-Analog integrating device.

Figure 7-35.-Basic inertial navigation system.

integrators (fig. 7-35), the system can determinedistance traveled in the north-south and east-west directions. It is important to maintainthe proper accelerometer pointed north andmaintain both accelerometers horizontal to theearth’s surface. If the accelerometers tilt offlevel, it measures gravitational components,which results in navigation errors. A thirdaccelerometer sometimes mounts on the stableelement in the vertical plane to determinevertical acceleration. The computer subtractsthe gravity component from the output ofthe accelerometer. The resulting signal representsactual aircraft vertical acceleration. A verticalacceleration signal goes to an integrator inthe attitude computer. This computer computesvertical velocity.

Platform Stable Element

To maintain the proper orientation ofthe accelerometers, they mount on a stableelement together with gyroscopes. The gyro-scopes are the sensing elements for controllingthe orientation of the stable element. Thestable element (fig. 7-36) mounts on gimbals,which isolate it from angular motions of theaircraft.

GYROSCOPES. —The stable element con-tains two identical, floated, two-degree-of-freedomgyroscopes. They mount one on top of the otherin a dumbbell configuration (fig. 7-36). Thegyroscopes have their spin axes horizontal and atright angles to each other. The wheels in thesegyroscopes, which spin at high speed, resist anyeffort to change the orientation of their spinaxes.

Figure 7-36.-Simplified platform stable element.

Figure 7-37.-Single-axis, gyrostablized platform.

Figure 7-37 shows a two-degree-of-freedomgyro and a single-axis stable platform. Thepickoffs on the gimbals within the gyro produceelectrical signals. These signals occur when thegyro case moves from its null position with respectto the gyro motor. The electrical pickoffs willsense any displacement of the stable element fromthe frame of reference. The signals thus createddrive the platform gimbals to realign the stableelement.

PLATFORM GIMBAL STRUCTURE. —Figure 7-36 shows the four-gimbal platformconfiguration actually used in an inertial naviga-tion system. The stable element mounts in thegimbal structure so that, regardless of aircraftmaneuvers, it retains the original orientation. Thestable element serves as a level mount for the

Figure 7-38.-Gimbal flipping action.

Figure 7-39.-Earth rate torquing: (A) without gyro torquing; (B) with gyro torquing.

accelerometers. An azimuth gimbal lets the Consider what happens to a stable element asaircraft change heading without affecting theorientation of the stable element. A pitch gimbalremoves the effect of aircraft pitch, and a rollgimbal stops the effects of roll. An extra rollgimbal prevents the occurrence of gimbal lockduring certain aircraft maneuvers and makes thesystem truly all-attitude. Look at figure 7-38. Notethe inner roll gimbal that prevents gimbal lock,which would cause the stable element to tumble.Gimbal lock occurs when two of the gimbal axesbecome aligned parallel to each other. This causesthe stable element to lose one of its degrees offreedom. When the aircraft exceeds 90° in pitch,the outer roll gimbal rotates through 180°. Thegimbals are oriented so the system may sense air-craft attitude and heading by measuring anglesbetween the gimbals. Synchros send this informa-tion to the attitude indicator and other systemsin the aircraft.

PLATFORM ORIENTATION. —Figure 7-39,view A, shows the apparent rotation of a stabilizedplatform located at the equator. As shown, theplatform will remain fixed with respect toinertial space. However, it appears to rotate aboutthe surface of the earth as the earth spins aboutits polar axis. This is undesirable for navigationsince the accelerometers will not remain horizontalto the earth’s surface. Consequently, this producesgravitational components of acceleration in theoutputs of the accelerometers.

the aircraft flies over the surface of the earth. Asthe aircraft flies straight north from the equatorto the North Pole, the aircraft sees a continuingpitch maneuver. Look at figure 7-40, view A. Atthe pole, instead of the platform being level withthe surface of the earth, it is now 90° off level.

GYRO TORQUING COMPUTATIONS. —To overcome the problems that arise fromplatform tilt, the system uses the gyroscopicprinciple of precession. By using this principle asthe aircraft flies over the rotating earth, it ispossible to apply a continuous torque to theproper gyro axis. This reorients the gyros tomaintain the stable element horizontal to theearth’s surface and pointed north. Figure 7-39,

Figure 7-40.-Aircraft rate torquing: (A) withouttorquing; (B) with gyro torquing.

Figure 7-41.-Simple pendulum.

view B, and figure 7-40, view B, show platformoperation with proper earth-rate and aircraft-ratetorquing corrections.

An analog or a digital computer developsthe signals necessary to properly torque thegyros. The corrections for earth rate dependon the aircraft’s position on the earth’s surface.The analog corrections come from highly accuratepotentiometers that produce trigonometric func-tions of aircraft position. Position integratorshafts drive the potentiometers. To maintain thestable element oriented to the north reference,torquing corrections rotate the platform. Therotation is about the vertical axis compensatingfor vehicle velocity.

Schuler Pendulum

A pendulum is any suspended mass that is freeto rotate about at least one axis. However, itscenter of gravity is NOT on the axis of rotation.Therefore, any pivoted mass that is not perfectlybalanced is, by definition, a pendulum. Theinertial platform is a pendulous device and,therefore, behaves as all pendulums behave. Italigns to the dynamic vertical when at rest. Thepivot axis and the center of gravity align with thegravity vector. The center of gravity will be onthe bottom. Also, it tends to break into its naturalperiod of oscillation whenever the aircraftaccelerates.

Pendulous oscillation is periodic angularmotion with the gravity vector as its midpoint.Periodic motion around the local vertical producesobvious errors from an inertial platform.This happens because misalignment about thehorizontal plane introduces gravity componentson accelerometer inputs. The system will interpretgravity accelerations as horizontal acceleration ofthe aircraft. The Schuler pendulum is a speciallyconstructed pendulum without the unwantedoscillatory motions of non-Schuler pendulums. Itis a special case of both the simple and thecompound pendulums, which are discussed in thefollowing paragraphs.

SIMPLE PENDULUM. —The simple pen-dulum consists of a small body suspended by aweightless string. The motion of the simplependulum is both periodic and oscillatory. Theperiod of the simple pendulum is given by themathematical formula

The formula shows that the period of asimple pendulum is proportional to the squareroot of the length of the suspending string. Thelonger the string, the longer the period.

One property of the simple pendulum that isvery useful in the construction of an inertialstable element is shown in figure 7-41. Twopendulums are suspended by strings of differentlengths. Equal forces horizontally accelerate thesuspension point of each pendulum. The inertiaof the bob resists the change in its state ofmotion. This action causes the bob to lag the pointof suspension. It also produces an angularmotion of the pendulum about the local gravityvector. Figure 7-41 shows that the length ofpendulum (B) is longer than pendulum (A). It alsoshows angular motion of pendulum (B) is less thanpendulum (A) for a corresponding linear motionof the suspension point. Therefore, the longer thesuspending string, the less the angular motion ofthe pendulum for a given linear motion of thesuspension point.

Consider what would happen in the followingcase. The suspending string is long enough tomaintain the bob at the center of the earth. Thesuspension point is transported horizontally alongthe earth’s surface (fig. 7-40, view B). The bobis hypothetically at the center of the earth, the seatof the earth’s gravity field. Accelerating thesuspension point along the earth’s surface merely

realigns the suspending string with the new localgravity vector. Therefore, the angular motion ofthe pendulum about the gravity vector for anyhorizontal acceleration of the suspension point iszero. This particular pendulum is the Schulerpendulum, shown in figure 7-40, view B. Thispendulum gets its name from the Germanengineer, Maximilian Schuler.

Schuler solved the problem of oscillatingshipboard gyrocompasses in the early 1900s. Ofcourse, Schuler couldn’t use the simple pendulumitself to solve this oscillating problem. He usedthe principle of the simple pendulum to constructa pendulum that reacted like a simple pendulum.The length of this pendulum equals the radius ofthe earth, which is about 3,440 nautical miles long.The period of oscillation for this pendulum isabout 84.4 minutes. Remember the period ofoscillation of a pendulum is proportional to thesquare root of its length. Therefore, any pen-dulum constructed to oscillate with a period of84.4 minutes would have an equivalent length ofabout 3,440 nautical miles. Such a pendulum isthe Schuler pendulum, a special case of thecompound or physical pendulum. Figure 7-42shows three examples of compound pendulums.

COMPOUND PENDULUM. —In figure 7-42,view A, the pivot point, P, is farthest away fromthe center of gravity, represented by distance d.

Figure 7-42.-Compound pendulum.

In view B, the pivot point is closer to the centerof gravity than in view A. However, it is fartheraway than the one shown in view C, which pivotsat the center of gravity.

The pivot point of each pendulum in figure7-42 is given the same acceleration. Therefore,each pendulum has the same linear motion at itspivot point. Yet, each pendulum has a differentangular motion. As distance d decreases, theangular motion of the pendulum about the localvertical (gravity vector) decreases and distance Lincreases. Distance L is the distance from pivotpoint P to the center of oscillation, point O. Also,the pivot point and the center of gravity comecloser together, and equivalent length L of thependulum becomes longer. Figure 7-42, view C,shows the pendulum pivoted at the center ofgravity. In this case there is no angular motionof the pendulum and the equivalent length L isinfinite. Therefore, it is not a pendulum; it is aperfectly balanced mass that has an infinite periodof oscillation. Thus, it is possible to construct apendulum of infinite equivalent length and period.It is also possible to construct one that has anequivalent length of 3,440 nautical miles. Such apendulum would be pivoted at some distance dfrom the center of gravity. This distance wouldbe greater than the one in figure 7-42, view C, butless than the one in 7-42, view B. When pivotedat a point where the period of oscillation is foundto be 84.4 minutes, it becomes a Schulerpendulum.

The stable element is essentially a Schulerpendulum. However, it is not entirely mechanicalbecause the earth’s radius varies with latitude. The

Figure 7-43.-Frame of reference.

earth’s radius is greater at the equator than it isat the poles. For this reason the stable elementuses the process of Schuler tuning. Schulertuning torques the platform to a position normalto the gravity vector by signals received from acomputing loop.

Frame of Reference

The frame of reference about which the INSdefines the instantaneous position of the aircraftis the conventional latitude-longitude coordinatesystem (fig. 7-43). The local vertical, establishedand maintained by the inertial navigation system,is the gravity vertical and is coincident with thegeographic vertical. The inertial navigation systemorients to the true north reference by sensing themotion of the earth rotating on the polar axis.The frame of reference defined is horizontallyaligned in a plane parallel to the surface of theearth and oriented to true north.

ESTABLISHING THE REFERENCE. —Referto figures 7-36 and 7-43. By establishing the frameof reference, three perpendicular axes of the stableelement will align to the horizontal coordinatesof the latitude-longitude navigational system.That is, the stable element z-axis aligns withthe local vertical and the y-axis aligns north-south. Therefore, the z-axis is coincident withlines of longitude, and the x-axis aligns east-west coincident with lines of latitude. In allcalculations, x-axis is positive east and y-axis ispositive north. The z-axis is positive away fromthe center of the earth.

A pair of two-degree-of-freedom gyroscopesestablishes and maintains the stable element to theframe of reference. Since a two-degree-of-freedomgyroscope has two sensitive axes, it is necessaryto use two such gyroscopes (fig. 7-36). Theupper gyroscope z-axis is not in use. Theyphysically mount on the stable element so theirspin axes are exactly perpendicular in thehorizontal plane. With this arrangement, align-ment of the upper gyroscope spin axis north-southwill automatically align the lower gyroscope spinaxis east-west.

The stable element containing the gyroscopesis supported by the platform gimbal system. Thus,the gyroscopes control the stable element.However, if a free gyroscope initially orients sothe spin axis aligns east-west in a horizontal plane,the gyro will precess. The precession will be aboutthe earth’s surface because of the earth’s rotationon its polar axis. To maintain an earth reference,

the system must torque the gyro opposite andequal to the apparent precession. The earth’srotation affects the upper (XZ2) and lower (YZ1)gyros. Corrections for earth rotation go to theY and Z1 torquing coils (Z1 and Z2 are cagedtogether and both respond accordingly). Thesystem does not use the X torquing coil for earthrate corrections.

To establish an earth frame of reference, thegyroscopes are controlled by continuouslycomputed signals. These signals introduce forces(torque) to cause the gyro spin axes to precess inthe desired direction. This torque is in the formof direct current signals applied to torquing coilsmounted on the gyro float assembly. It creates amagnetic field that aids or opposes the magneticfields of the permanent magnets mounted on theend bells. This circuit effectively torques the gyro,causing the spin axis to precess to the desiredorientation.

The first step in establishing a frame ofreference is leveling the stable element. To levelthe stable element, you align it to the localvertical (gravity vector). This is done bytorquing the XZ2 and YZ1 gyros. This moves thestable element until the x and y accelerometerscease to sense any acceleration caused by gravity.This says that the outputs from the accelerometersprovide the torquing signals for the gyros.During this time, and while operating, computedearth rate torquing signals continuously go to they and z axes torquing coils of the lower gyro. Thesize of these earth rate torquing signals isresolved by computing the vertical and horizontalcomponents of earth rate as a function of latitude.

After establishing the stable element in a roughlevel position, the x-axis torquing signal drives thestable element in azimuth to null this signal (finealignment). This signal consist of earth rateacceleration only, which is a measure of stableelement unlevelness. At this time, the X gyro’sspin axis is aligned to true north establishing theframe of reference. This alignment condition willremain until you manually sequence the inertialnavigation system to the navigate position.

As previously shown, the earth’s rotation doesnot affect the upper X-axis gyro. Therefore, nocompensating earth rate torquing signal goes tothis gyro.

MAINTAINING THE HORIZONTAL REF-ERENCE. —When the aircraft remains stationaryor moves at a constant velocity, the accelerometeroutputs are zero. If the aircraft attitudechanges while maintaining a constant speed, the

accelerometers on an unstabilized platform sensean acceleration due to gravity. Since theaccelerometers cannot distinguish gravitationalaccelerations from horizontal accelerations, theintegrators develop a fictitious velocity with acorresponding distance error. It is essential,therefore, that the system maintains the accel-erometers in a truly horizontal reference plane.This plane must be independent of aircraftattitude at all times.

This is a basic requirement of the inertialnavigation system. The accuracy with which thehorizontal reference is maintained determines theoverall performance capabilities of the system. Agyrostabilized platform in a gimbal structureserves as an inertial reference. Also, it accuratelydefines directional reference for the coordinatesystem.

With the platform installed in the aircraft andaligned along the y-axis, it will remain levelregardless of aircraft attitude. The instant theaircraft begins to change pitch attitude (fig. 7-44),the platform gyro senses this angular movement.The gyros then begin to precess at a rateproportional to the pitching rate. A pickoff coilon the gyro axis senses this movement and changesit to a voltage. After amplification, the voltagegoes to the pitch gimbal servo drive motor. Themotor rotates the stable element exactly equal andopposite to the aircraft angular motion. As aresult, it continuously precesses the gyro to itsneutral or level position. This action maintainsthe gyro output signal at null.

Regardless of any new pitch attitude theaircraft assumes, the gyro will keep the stableelement level. This is the only position that allowsthe gyro’s output signal to be at null. In actualpractice, the stable element is maintained level.It uses a similar method to maintain its azimuth

Figure 7-44.-Gyro maintaining inertial platform level.

alignment in yaw and roll. This is necessary sothe sensitive axis of the north-south accelerometeraligns true north-south. It is also necessary forthe east-west accelerometer to align true east-west. The stable element is then accurately alignedto the three coordinates—north (y-axis), east(x-axis), and up or true vertical (z-axis).This arrangement allows the accelerometers toaccurately detect aircraft motion.

MAINTAINING THE VERTICAL REF-ERENCE. —The stable element must remainperfectly level or the accelerometers will sense afalse acceleration due to the earth’s gravity. Thegyros try to maintain their inertial position inspace and not with respect to the local vertical.This causes the stable element to drift off levelas the aircraft moves over the curvature of theearth (fig. 7-40). If this situation were allowed tobuild up, very large errors in velocity and distancewould occur. This condition develops whether ornot the aircraft is moving over the earth’s surface.The earth’s rotation alone will develop the sametype of errors.

The stable element must always be perpen-dicular to the local vertical. Therefore, the systemmust make the gyros precess in such a manner asto maintain the stable element level. With thestable element level, the sensitive axes of theaccelerometers are maintained horizontal to theearth at all times. Now they respond only to thehorizontal component of acceleration.

MAINTAINING THE FRAME OF REF-ERENCE. —Accuracy in maintaining the stableelement to the frame of reference determines theoverall performance capabilities of the system.Gyro torquing rate signals are continuouslycomputed to maintain the frame of reference.After alignment, the inertial navigation system ismanually sequenced to its operating condition(navigate). If the aircraft were to remainstationary, the gyro torquing rate signals wouldconsist of earth rate only. However, as theaircraft moves over the curvature of the earth, thestable element earth reference would be lost. Thishappens because vehicle movement causes thegyros to precess. (Refer to figures 7-39 and 7-40.)Therefore, the system continuously computesadditional gyro torquing signals to compensatefor the vehicle’s movement. These signals areaircraft rate torquing signals. They depend on thevelocity of the aircraft wth respect to the frameof reference; that is, east-west velocity and north-south velocity. The angular rate is directly

proportional to the velocity of the aircraft alongthe circumference of the earth. The systemapplies aircraft rate torquing to both gyros so theywill precess about all three axes (x, y, and z). Thus,the system continuously maintains the frame ofreference.

Deriving Velocity and Distance

The inertial navigation system can accuratelydetect aircraft acceleration, and using precisionintegrators, determine aircraft velocity andmeasure distance traveled. Accelerations aremeasured in units of ft/sec2 by the accelerometers.However, for analog computation, the accelera-tions are developed in volts per g of acceleratingforce. The velocity integrator integrates the ac-celerating force to obtain velocity. The distanceintegrator integrates velocity to obtain the distancetraveled.

There are two velocity integrators in the systemto obtain VX and VY along the two horizontal axes.There are two distance integrators to obtaindistance traveled along the two horizontal axes.In addition, some inertial navigation systemsemploy one other velocity integrator to obtain VZ

along the vertical axis.

Accelerometer Output Corrections

The arrangement of the accelerometers isperfectly suited to navigation over a stationaryplane or over flat terrain moving at uniform speedin a straight line. The earth, however, is a rotatingsphere. As far as the inertial platform isconcerned, only points along the equator can beconsidered to possess uniform linear motion.Here, and only here, the accelerometer signals cantranslate directly into position information. Forthis reason, it is necessary to provide an automaticdevice to alter the accelerometer signals, Theautomatic device allows the system to reportmeaningful information. The corrective device ispurely electrical. All or part of the circuitryis active whenever a velocity signal voltage ispresent anywhere north or south of the equator.These circuits use the velocity signals, modifiedaccording to the latitude of the aircraft position,They insert artificial acceleration signals to thosealready in the accelerometer output circuits.

The circuits are divided logically—some arefor centripetal effect, some for Coriolis. Note,however, that the corrections are complementary.

Although it is convenient to assign a separatepurpose to the circuits, as in the following

discussions, the functions overlap, and it is notaccurate for you to consider them separately.

CENTRIPETAL CORRECTION. —Centripetalerrors are false accelerations sensed when theplatform is torqued to maintain its plane ofreference. Centripetal correction differs from thatof the Coriolis correction in that it has norelationship to earth dynamics. Even if the earthwere stationary, it would still be necessary toinsert centripetal correction voltages to theaccelerometer signal. Correction voltages ensurean accurate plot of any course that does notcoincide with one of the earth’s coordinate greatcircle routes; that is, an exact polar or an exactequatorial orbit.On a great circle route, a routethat circles the earth’s center, every linearacceleration initiates motion. However, unless theroute is due north-south or directly along theequator, the system can’t plot the route accuratelyfrom raw accelerometer signals.

The orbit resulting from linear acceleration,and the logic of applying centripetal correctionsto get an accurate plot of track, can be understoodif you consider any simple circumstance involvinga single acceleration and the inertial reaction. Forexample, assume a perfect bowling lane surroundsthe earth at the equator. Theoretically, a bowlercould stand behind the pins and roll the ball inthe lane in the opposite direction (away from thepins). After a few years time, the bowler gets a

perfect strike on the head pin. However, if thelane is on latitude 10°N, the ball would invariablyveer south. After a few thousand yards, the ballwould fall into the right-hand gutter if rolled east.It would fall into the left-hand gutter if rolledwest. In both cases, the ball would roll into thesouth gutter. The reason for this phenomenon isclear if you consider a lane built on a latitude inthe Arctic only a few yards from the pole. Here,the curve of the lane is obvious. It is apparent ifthe ball accelerates due east; it will follow, ininertial space, a straight course intersecting theoutside gutter. For the ball to roll at a uniformspeed and remain in the alley, a uniform northacceleration force must be exerted on the ball intransit. In this case, the north acceleration forceis exerted on the ball during transit. Also,remember that north acceleration is a correctiveforce and does not produce north velocity withrespect to the alley.

As shown in figure 7-45, a north-southaccelerometer aboard an aircraft circling the polefinds the same phenomenon. In a steep-bankedturn, the centrifugal force deflects the north-southaccelerometer. It blindly reports a steady northacceleration. The system will show a growingnorth velocity when, in fact, only an east velocityexists.

The fault is not in the accelerometers but inthe geographic coordinate system’s nonlinearpattern to which the platform is slaved. The

Figure 7-45.-Centripetal correction

along latitude.

Figure 7-46.-Platform on great circle route.

coincidence of earth axis and pole implies earthdynamics in the phenomenon, but actually thisis not a factor. Assume we shift the coordinatesystems to place the pole at New York. Now usethis city as the focal point of one accelerometeraxis of an inertial navigator circling the city. Theaccelerometer sensitive to this axis will reportacceleration toward the city.

In a spherical coordinate system, any linearvehicle acceleration initially affecting bothaccelerometers will result in the vehicle NEVERreaching the pole. The vehicle follows a greatcircle track that first approaches one of the poles.It will fly due east or west for a brief period, andthen depart the pole, as shown in figure 7-46.

Although the velocity resulting from any givenacceleration is initially computed correctly withrespect to space, the direction of the speed mustbe constantly altered. This is necessary if thenavigational system is to accurately report a greatcircle course that crosses both latitude andlongitude. On such a course, the aircraft does notturn as it does following a line of latitude.Therefore, it does not generate uniformly falsenorth acceleration signals. The centripetalcorrection circuit does, however, continue to plantsignals of acceleration toward the equator. Thishas the effect of altering the reported speed (theresult of velocity along both axes). No actual

acceleration has taken place to cause a speedchange. Thus, the centripetal correction circuitmust simultaneously plant positive accelerationsin one axis if it plants negative accelerations inthe other. Therefore, when canceling some of thenorth or south velocity, the system must add asufficient increment of east or west acceleration.This allows the reported total speed to remainunchanged, while the reported direction of flightbends toward the equator. As shown in figure7-46, after a single north-east acceleration hasoccurred, the north vector becomes progressivelyshorter. Note that the resultant speed vector isalways of the same size.

In summary, you can see that an east or westacceleration in either hemisphere contains ahidden element of acceleration toward theequator. Also, you can say centripetal correctionsimply acts to reveal this element. When aircraftvelocity is not due north or due south, in theNorthern Hemisphere the centripetal correctionmanufactures a south velocity component. In theSouthern Hemisphere, it manufactures a northvelocity component.

CORIOLIS CORRECTION. —As mentionedbefore, the scope of the centripetal correctionmakes no allowance for earth dynamics. Since theearth rotates toward the east, all points on the

surface have a constant tangential velocity. Thisvelocity is maximum along the equator and lessensat higher latitudes.

Tangential velocity is a linear quantity. Itrefers to the speed and direction an object wouldtravel in a straight line if freed from the earth’sgravity. An object near the equator travels about1,000 miles per hour in a circular path. Assumethe object is free from earth’s gravity andatmosphere. Now it will travel at that speed ina straight line away from the earth (on a tangentto the earth). Its tangential velocity is 1,000 milesper hour. If the object moves toward the NorthPole, its speed in circular travel will decrease asit approaches the pole. The speed will be zerowhen placed exactly over the pole.

Although the earth has a trajectory inspace, this motion is not important to theinertial navigation system because every pointon the sphere shares this trajectory. The onlyvariable involved is the variation in the earth’stangential velocity at different latitudes. While the

accelerometers don’t automatically makeallowance for this variation, the system must takeit into account.

The need for such an allowance is illustratedin the following discussion. Put an aligned INS,devoid of any Coriolis corrective mechanism,aboard a train in the Northern Hemisphere, andtransport it north. The earth’s tangential velocityat the latitude where the INS aligned is 800 knotseast. It is obvious the train’s eastward velocitymust be constantly reduced as it progresses north.This progressive reduction in velocity representsan acceleration.

Since the tracks constrain the train, a forcefrom the east will be exerted on the wheel flanges.This force is sensed by the east-west accelerometeras acceleration to the west. The system beginscomputing west velocity that will grow. Refer tofigure 7-47. An aircraft flying north directly alonga moving longitude meridian subjects the east-westaccelerometer to this same force. This happensas its course in space alters to the left,

Figure 7-47.-Coriolis effects.

Figure 7-48.-Coriolis and centripetal corrections.

compensating for the decreasing eastward velocityof the earth’s surface.

The train traveling north stops at a latitudewhere the earth’s tangential velocity is 600 knots.This is 200 knots less than the tangential velocityat the point of its initial alignment. The velocitycomputer will continue to report that the vehicleis traveling west at 200 knots even though it isperfectly static. The Coriolis correction preventsthis from happening. It creates enough eastacceleration signal to offset the continual westacceleration generated by the east-west accel-erometer as it travels north. The east and westaccelerations cancel, thus, there is no change inlongitude to report. Of course, on the return trip,the platform is under the delusion that 600 knotseast is zero velocity. So as it travels south, theincreasing earth tangential velocity causes it toconstantly report east acceleration. In this case,the Coriolis correction also reverses. It nowmanufactures a west acceleration that exactlyvoids the east acceleration.

In both of the above cases, the Corioliscorrection supplies an artificial signal of accelera-tion to the right side of the actual track. This isthe nature of the Coriolis correction, regardlessof the direction of travel in the NorthernHemisphere. Look at figure 7-48. When the trackis east along a latitude line, centripetal correctionoffsets the increment of north acceleration

generated by aircraft speed. The Coriolis correc-tion accounts for the additional force generatedby the earth’s rotation. For example, if the earth’stangential velocity at latitude L is 700 knots andaircraft speed 350 knots, total speed is 1,050knots. Centripetal correction offsets the accelera-tion caused by the centrifugal force of a 350-knotturn of this radius. Coriolis correction offsets theforce of a 700-knot turn. In this case, the twocorrections are summed.

If the course is reversed and the plane flieswest, the correct ions become opposite in polarity.The centripetal signal is still south, but Coriolisis north (to the right of the track). Therefore, onlya part of the larger correction is effective.

Schuler-Tuned Loop

The Schuler-tuned loop is a closed loop circuitbetween the accelerometer, velocity integrator,and stable element. It prevents large velocity anddistance errors caused by misalignment of thestable element. Figure 7-49 shows a simplifiedSchuler-tuned loop with the platform aligned.

The output of the accelerometer is integratedto provide a velocity signal. The velocity signalis multiplied by 1/R where R equals the earth’sradius. This operation derives an angular velocityabout the earth’s surface, V/R. The system usesthis angular velocity to torque an integrating gyro.

Figure 7-50.-Simplified Schuler-tuned loop, platformunlevel.

The torque causes the platform to precess aboutthe earth‘s surface. This precession equals the ratethe platform is being transported over the surface.This maintains the platform normal to the localvertical.

There are two such loops in an inertialnavigation system, one for the north and the otherfor the east. The accelerometer in the north loopsenses north-south accelerations, yet the gyro inthe north loop senses east-west angular rates. Thatis, the vehicle’s angular movements about the east-west axis.

By convention, we name accelerometers andgyros according to the direction of their sensitiveor input axis. The inertial or Schuler loop takesthe name of its accelerometer. The north loopcontains the north accelerometer and the eastgyro. The east loop contains the east acceler-ometer and the north gyro.

With the platform initially unlevel, as shownin figure 7-50, the accelerometer senses acomponent of gravity, g sin . This signal isintegrated, resulting in the velocity signal V . Thevelocity signal then causes the gyro to precess ina clockwise direction. When the accelerometer ispositioned to sense zero gravity, the velocityoutput continues to torque the platform in aclockwise direction. This causes the accelerometer

Figure 7-51.-Schuler tuning—acceleration errors versusvelocity errors.

to now sense a gravity component of the oppositepolarity. This signal causes the velocity signal todecrease to zero. The velocity signal now buildsup in the opposite direction and precesses theplatform in a counterclockwise direction. Theoscillation set up by this mechanization has aperiod of 84 minutes, equal to that of the Schulerpendulum. Figure 7-51 shows the buildup anddecay of acceleration and velocity errors as a resultof such errors as just described.

ALIGNMENT

Inertial navigation depends on the integrationof acceleration to obtain velocity and position.In any integration process, the system must firstknow the initial conditions. In this case theinitial conditions are velocity and position. Theaccuracy in solving the navigation problemdepends greatly upon the accuracy of the initialconditions. Therefore, system alignment is ofparamount importance.

System alignment consists of creating acoincidence between the platform axes and thecomputer axes. This can be done by rotating eitheror both systems. There are two general methodsof accomplishing this condition.

The system is slaved to an externalreference source.The system may have the built-in capabilityto sense misalignment and correct itself.

Of course, combinations of both these methodscan be, and often are, used.

External references take three basic forms—terrestrial, celestial, and inertial. The terrestrialsystem uses surveyed lines, bench marks, plumbbobs, and bubble levels. These methods result inlevel accuracies of about 10 seconds of arc andheading accuracies to 3 minutes of arc. Celestialinformation from star trackers and radio sextants

Figure 7-49.-Simplified Schuler-tuned loop, platform level.

has accuracies to 10 seconds of arc. When usingan inertial system as a source, the accuraciesdepend on its initial source and length of timesince last aligned. Such a method is usually formobile alignment where primary sources cannotbe used.

The use of an external reference systemrequires transfer devices to transmit the referenceinformation to the system. The transfer devicesare either optical or electromechanical devices.Optical devices include theodolites and auto-collimators; the electromechanical devices aresynchro-resolver or digital type. Optical methodscan produce accuracies of a few seconds of arc.The electromechanical methods are accurate toabout 30 seconds of arc.

In self-alignment, the inertial sensing instru-ments mounted on the platform sense the deviationfrom the desired position. To determine theorientation of a three-axis orthogonal (perpen-dicular) coordinate system, you must have at leasttwo noncollinear (not parallel) reference vectors.For self-alignment, the earth’s spin vector and themass attraction gravity vector serve this purpose.The self-alignment puts less requirements on thecomputer. In self-alignment, the accelerometeroutputs don’t have to be resolved into componentsof gravity and vehicle acceleration. Whenaccelerometers and gyros mount on the sameelement, their relative position is fixed and doesnot have to be computed.

Self-alignment often divides into two modes—rough or course alignment (sometimes calledcaging) and fine alignment. Fine alignment itselfdivides into two modes—leveling and gyro-compassing.

Rough Alignment

Rough alignment provides a convenientstarting point for the later phases of alignment.In most cases, the gimbals slave to their ownsynchro outputs or to some external source. Thisexternal source will have a particular orientationwith respect to the vehicle. A timing networkcontrols the duration of rough alignment, whichis usually a short period.

Fine Alignment (Leveling)

To accomplish fine alignment or leveling, thesystem rotates the platform axes to the computeraxes. For a locally level system, this is done byplacing the X and Y accelerometer axes mutuallyperpendicular to the gravity vector. Since

accelerometers mount at right angles, a motionabout one axis causes the other to go through anangle with respect to the gravity vector. Therefore,you connect the accelerometer output in a wayto allow it to torque about its perpendicular axis.The accelerometer can now slew itself to a nullposition, where it senses no component of gravity.As pointed out earlier, the device the system cantorque about an axis is a gyro. In this case, thegyro being torqued is the one sensitive about theaxis perpendicular to the accelerometer beingleveled; that is, the Y gyro (north) torques the Xaccelerometer (east) to level, etc.

The accelerometer will provide a dc voltagethat is proportional to the sine of the off levelangle. In the leveling mode, after amplificationthis voltage goes to the gyro, whose sensitive axisis perpendicular to the sensitive axis of theaccelerometer. (See fig. 7-52. ) In other words, theoutput of the X accelerometer goes to the Y gyro.The output of the Y accelerometer goes to theX gyro.

Gyrocompassing

As mentioned, any gyro not having its spinaxis parallel to the earth’s spin axis will apparentlyprecess about the earth. The rate at which itprecesses is proportional to the angle between itsspin axis and the earth’s spin axis. This angle canbe resolved into two components. Since the gyrospin axis lies in the level plane already established,one component can be found in the plane itself.Now, find the angle between the spin axis and avector in the level plane that intersects the earth’sspin vector. Once you find these two angles, youknow the exact position of the platform axis.

You can see the rate at which the gyro inquestion appears to precess is equal to sina where

The spin axis of this gyro can be torqued to a placewhere this term goes to zero Also thecomputer can develop a torquing term equal tothis and apply it to the gyro. In either case, theposition of the platform is known. Figure 7-53shows a typical heading alignment loop.

Figure 7-52.-Typical leveling loop.

Figure 7-53.-Typical heading alignment loop.

In a north seeking platform, the system usesthe earth’s rotation to align the platform to truenorth. It accomplishes this by using the outputof the y accelerometer. The system applies thisoutput to the torquing coils of the z (azimuth) andthe x (east) gyros. At the beginning of the gyro-compass phase (after the platform is leveled), thestable element is torqued in azimuth. This nullsout the residual east gyro torquing rate. If thestable element is not aligned to true north now,it begins to tilt due to precession of the gyros. They accelerometer senses deviation of the stableelement from level because of gravity. Theoutput of the accelerometer then torques the xgyro until the stable element is level. At the sametime, the output signal also torques the z gyro inazimuth. The process continues until the stableelement aligns to true north.

Once the platform is aligned, the operatorswitches the system from the alignment phaseto the navigation phase of operation. In thenavigation phase, the stable element wouldmaintain an orientation about free space if notfor corrections supplied by the computer. Thecomputer maintains the stable element level withrespect to the earth and oriented to true north.If not, the accelerometers sense gravity inaddition to movement of the aircraft. Coriolis,centripetal, and earth rate corrections arecomputed and used to hold the stable element leveland aligned to true north. In the navigation phaseof operation, the orientation to true north isdependent on the original aligned position and thecomputed corrections.

To this point, a north-pointing inertial systemhas been discussed. A disadvantage of the north-pointing system is that it cannot operate in thepolar regions; it must always be physically pointednorth. If the system flies directly over the pole,it must rotate 1800 to again be pointing north.This rotation would not be physically possiblebecause of the extremely high torquing ratesnecessary. Most north-pointing inertial systemscannot operate within several hundred miles ofthe poles due to stress on the system components.

Wander Azimuth

The wander-azimuth inertial system solves theproblems of operating an inertial system at thepoles. The fundamentals of a wander-azimuthsystem are the same as a north-pointing system.During the gyrocompassing mode, the systemallows the platform to take an arbitrary angle(wander angle) with respect to true north.

As previously mentioned, the platform isleveled; but, the accelerometer outputs are nowsupplying torquing signals to both gyros. Thisaction compensates for the earth’s rotation (thissignal was sent to x gyro only in the north-pointingsystem). Eventually, the correct earth rate torqu-ing signals maintain the platform level. The com-puter then uses the ratio of earth rate compensa-tion to compute the wander angle (fig. 7-54).

As the wander-azimuth system navigatesaround the earth, the wander angle (with respect

Figure 7-54.-Wander angle.

to true north) changes as a function of longitude.The system operation is the same as a north-pointing system. However, the wander angle istaken into account by the computer with the northand east sensored accelerations.

Alignment at Sea

Problems that arise in aligning an aircraftinertial navigation system on aircraft carriers atsea are, more complex than aligning the INSashore. This situation exists even though ourcarrier-based inertial reference is another INS ofvery high accuracy. The inertial navigationreference system aboard an aircraft carrier is theship’s inertial navigation system (SINS) and therelative velocity computer (RVC). Outlets on theflight deck make it convenient to pipe the SINSreference information into the aircraft inertialnavigation system. However, one major problemstill remains that makes proper alignmentdifficult. The accelerations experienced by theship’s inertial navigation accelerometers locatedbelow deck. These accelerometers are remote fromthe aircraft accelerometers, and therefore do notsense the same accelerations.

COARSE SEA ALIGNMENT. —Duringcoarse sea alignment, the best available trueheading is from the SINS and the RVC. Aircraftcarrier true heading goes to the RVC. Here amanually selected aircraft heading angle, withrespect to the aircraft carrier, is inserted. Thecombined signals then go to the heading computerin the aircraft’s inertial navigation system, thusconcluding the coarse sea alignment.

FINE SEA ALIGNMENT. —During fine seaalignment, the accelerometers sense the aircraftcarrier movement in addition to gravity. Only thegravity component is used in the leveling. Thisallows accelerometer output caused by aircraftcarrier movement to be canceled. The SINSand RVC accomplish this task by supplyingcontinuously computed corrections. The accel-erometer error signals are integrated to supplya velocity. The reference velocity supplied bythe RVC during sea alignment, actual aircraftvelocity, is subtracted from the accelerometerderived velocity. The difference corresponds tothe gravity component sensed by the acceler-ometer. After amplification it is applied to thetorquing coils of the gyros. The processing of the

gyro pickoff signals cause the gimbals to rotateand cancel the pickoff error signals. The stableelement is torqued until the accelerometers showa null or level condition.

TYPES OF INERTIALNAVIGATION SYSTEMS

Learning Objective: Recognize the twotypes of inertial navigation systems anddiscriminate between systems within thosetwo types.

You can classify inertial navigation systemsunder two broad types—pure and hybrid. Thetypes of pure inertial navigation systems areanalytic, semianalytic, geometric, and strap-down. The types of hybrid inertial navigationsystems are radio inertial, Doppler inertial, andstellar inertial.

PURE INERTIALNAVIGATION SYSTEMS

As the name implies, a pure inertial naviga-tion system is not combined with other equipmentto improve its operating performance.

Analytic Inertial Navigation System

The analytic INS uses a platform with a fixedangular reference to some point in inertial space.The system makes no attempt to force theaccelerometer input axes into a preferredalignment with respect to the earth. This methoddoes not require gyro torquing. As a result,this platform is subjected to errors of gyro driftonly.

Because the platform remains rigid inspace and rotates about the earth, the output

accelerations become complex. They essentiallyconsist of two major accelerations—the actualacceleration of the vehicle and the gravitationalacceleration of the earth. For navigation purposesonly aircraft accelerations are required andwanted; therefore, the gravitational accelerationsmust be canceled. Yet, this cancellation is difficultto obtain because the earth’s gravitationalacceleration is not uniform. Therefore, thecomputer must store an enormous amount of datato effect this cancellation.

The significant disadvantage of analyticinertial navigation is the result of maintaining theaccelerometer referenced to a fixed point ininertial space. As the stable element navigatesabout the earth, the accelerometers must senseaircraft acceleration and the earth’s gravitationalfield component. The accelerometers for thissystem must have a wide dynamic range as wellas a high overall accuracy. The most seriousproblem, however, is the cancellation of thegravitational accelerations. Irregularities in theearth’s shape and mass cause variations in thegravitational field. Therefore, the cancellation ofthese variations requires a complex computer witha very large storage capacity.

Semianalytic Inertial Navigation System

The semianalytic system is the inertialnavigation system most commonly in use today.All naval aircraft that use inertial navigationsystems have this type of system. It may eitherbe a pure system or work with another navigationsystem as a hybrid system. This system’s chiefadvantages are the simple platform gimbalstructure and computer functions are easilyattained by either analog or digital means.

The semianalytic system always maintains thestable element normal to the earth’s gravitationalvector just as in other systems already discussedin this chapter. In this system, the computerconverts the output accelerations of the stableelement to angular velocities. These angularvelocity signals then torque the platform gyrosto maintain the platform normal to the earth’sgravity vector.

The computer also develops signals to preventthe platform from processing off level due to theearth’s rotation about its polar axis. These signalsare equal to the angular velocity of the earth

resolved into the system axes. The system thenapplies these signals to the gyro torquers. Atypical simplified block diagram of a semi-analytic inertial navigation system is shown infigure 7-55.

In a semianalytic inertial system, the platformaligns normal to the gravity vector. It may or maynot align to true north. The output of the northaccelerometer goes to an integrator. Here, theoutput is summed with acceleration correctionterms to derive a true vehicular acceleration overthe earth’s surface. This acceleration signal is thenintegrated with respect to time, deriving the northvelocity component of the vehicle’s track.Through scaling, the INS converts the velocityterm to an angular velocity. It then integrates theangular velocity to provide a position readout inthe form of latitude. In addition, the latitudefunction generator uses the north angular velocitysignal to develop accelerometer correction terms.It also uses this signal to develop a gyro torquingsignal for the east gyro.

The east accelerometer output is summed withaccelerometer correction terms and integrated toprovide an east component of the vehicle’s track.Through scaling, the INS converts the velocitysignal to an angular velocity. This angular velocitygoes to the latitude function generator. Thefunction generator uses it to develop acceler-ometer correction terms and torquing signals forthe north gyro and the azimuth gyro. In addition,the INS integrates the angular velocity to developa position readout in the form of longitude.

Geometric Inertial Navigation System

The geometric INS uses a gyro system that,like the analytic system, is referenced to inertialspace in a nonrotating plane. The accelerometers,however, mount on the gimbal structure in amanner as to remain normal to the earth’sgravitational field.

Figure 7-56.-Transport of the geometric system’s stableelement.

Figure 7-55.-Typical semianalytic inertial navagation system, block diagram.

Figure 7-57.-Strap-down inertial navigation system, block diagram.

Figure 7-56 shows the relationship of theaccelerometers and gyros as the platform movesover the surface of the earth. When the platformis aligned at the equator and then moves north,the gyros maintain their position in inertial space.The accelerometers remain in a plane tangent tothe earth’s surface at all times.

The main advantage of this system is that thegyros are not torqued. Therefore, scaling of thegyros is not critical.

The major disadvantage is economy. Thesystem requires a high degree of accuracy toposition the latitude and longitude gimbals. Thesemianalytic system requires much less precisionto achieve similar accuracy; therefore, it costs less.

Strap-Down Inertial Navigation System

In the strap-down system, the gyros andaccelerometers mount directly to the frame of thevehicle. Its principal use is in ballistic missiles andspacecraft. This type of system can be mechanizedfor use in aircraft. However, the present state oftechnology makes it more feasible to use one ofthe other type of systems for aircraft use.

The strap-down system requires complexdigital computers; analog computers are notaccurate enough for use in this system. Thecomputer in the strap-down system replaces thegimbal structure as the gimbal structure replacedthe physical length of the Schuler pendulum.Figure 7-57 shows a simplified block diagram ofa strap-down inertial navigation system.

In the strap-down system, the gyros provideangular rates, which the system converts todirectional cosines (for example, space vectors).The strap-down uses these signals to determinevehicle attitude about an inertial frame ofreference. The coordinate converter, using inputsfrom the accelerometers and the directional cosineconverter, determines accelerations along theinertial reference axes. The position converter

accepts inertial acceleration and altitude informa-tion to develop Cartesian coordinates representingthe vehicle’s position in inertial space. Thesevectors then go to the vector solver, where theyare summed to provide readouts of latitude andlongitude.

To accomplish strap-down system alignment,you must supply the directional cosines of thevehicle frame to the computer. The vehiclerequires no physical orientation.

HYBRID INERTIALNAVIGATION SYSTEMS

The hybrid system is a combination ofinertial navigation system and some other type ofnavigation system. The other navigation systemis for updating or improving the accuracy of theinertial navigation system. In other words, thehybrid inertial system combines two navigationsystems so that the good characteristics of bothare maintained.

There are two types of updating processes usedin hybrid systems. One type is the damping effect,which compares the inertial ground velocities withthe ground velocities of some other system. Thesystem uses the error, or difference between thetwo velocities, to damp out platform errors. Theother type is the reset method. This methodignores the orientation of the platform and merelyresets the position of the velocity shaftsperiodically.

CALIBRATION

As you learned earlier in this chapter,variation and deviation affect the accuracy of amagnetic compass. Variation is a naturalphenomenon whose magnetic strength varies inintensity throughout the world. Variation ismarked on navigation maps and is corrected forby the pilot. You can consider deviation as

man-made. The magnetic fields of aircraftcomponents cause deviation. These componentsare engines, electric equipment, landing gearstruts, and flight control surfaces and theircontrol cables. You can keep the effects ofdeviation to the very minimum by a process calledcompass swinging.

As an AE, you must be completely familiarwith the two methods of compass swinging. Thesetwo methods are the MC-2 compass calibrationset and the use of a compass rose. For a gyro-stabilized compass or inertial navigation system,the MC-2 is considered the primary means ofswinging. Deviation in a wet or standby compassis corrected for on a compass rose. The followingparagraphs describe the compass rose and theMC-2 calibration set.

Compass Rose

Aircraft magnetic compasses (wet or standby)have devices called compensators, which providea means for correcting deviation errors. Youcannot eliminate all errors, but you can reducethem to a minimum.

Swinging the compass, you first compensatethe N-S and E-W headings. Then set the aircrafton every 15- or 30-degree heading on thecompass rose. Here, you note the differencebetween the aircraft heading and the indicatedheading. You then adjust the compensators toreduce this difference or deviation to a minimum.

There are two types of compensators. Onetype is the universal screw type. It consists of anassembly having a group of small compensatingmagnets permanently installed in it. To changethe compensating effect of the assembly, you usetwo adjusting screws. One screw is for north-southcompensation, the other for east-west. The othertype of compensator has small, loose magnets thatyou place in special chambers on the compass asneeded. The chamber positions allow one to makeeast-west corrections. The other (at right anglesto the east-west chamber) corrects north-southdeviation. Compensation is done only on thecardinal headings on standby compasses. How-ever, on all other compass systems in navalaircraft, compensation is at 15-degree increments.

Before starting the swinging operation, youshould make sure all magnetic equipment is in theposit ion it occupies in normal flight. Also, be surethat no one near the aircraft compasses duringswinging operations has any magnetic materialson their person. Magnetic materials include tools,pocketknives, mechanical pencils, wristwatches,

Figure 7-58.-Compass rose, with aircraft on south heading.

dog tags, bracelets, eyeglasses, jewelry, officercaps, badges, etc. Remember, too, that you usea nonmagnetic screwdriver in adjusting universalcompensator screws.

You actually swing a compass in one of severalways. However, as an AE, your chief interest isin ground swings. You usually accomplish aground swing with the aircraft at rest on acompass rose. Look at figure 7-58.

Most air stations have a compass rose. Thecompass rose looks much like an oversized cardfrom a navigation compass. The directions shownby it are magnetic directions, and the northarrow points toward the earth’s north magneticpole. A compass rose may also have a lineshowing true north.

Jacks, lifts, hoists, or any dolly needed toperform the ground swinging job shouldpreferably be of nonmagnetic material. However,this is not always possible. Devices used in theswinging process must be tested for their effectson the compasses. You do this by moving themabout the aircraft in a circle with normalseparation distance between the device and theinstruments. Do not use devices that cause morethan one-quarter degree change in the compassreading.

Trucks, automobiles, railroad cars, and otheraircraft containing magnetic metals should not bewithin the swinging area. These items could havea magnetic effect on the compasses of the aircraftbeing adjusted. You should be sure that thecompass is in good condition. Examine thecompass for clear liquid and proper level. Checkto see that the card assembly is level. Also checkthat it turns freely when the aircraft’s tail is ina level flying position.

Set the compensator so it has no effect on themain compass magnets. Using a loose-magnetcompensator, remove all loose magnets from theirchambers. Set universal screw-type compensators

for zero effect by turning both adjusting screwsuntil the dots on the screws match with the dotson the compensator case.

Then, place the aircraft on a south magneticheading over the compass rose, with the tail ina level flying position. The aircraft engine(s)should be turning, and as many pieces of avionicsequipment as possible turned on. This will createas many stray magnetic fields as possible andsimulate the condition of the aircraft in flight.Note the compass reading and record it. From thisreading, it is simply a matter of algebraicsubtraction (or subtraction of numbers havingplus and minus signs) to determine the deviationon the south heading. The deviation is thealgebraic difference between the magnetic headingand the compass reading. Deviation is the errorin a magnetic compass caused by electromagneticdisturbances in the aircraft.

After doing this, place the aircraft on a westheading. Again, note the compass reading anddetermine the deviation or difference between themagnetic heading and what the compass reads.Next, turn the aircraft heading to magneticnorth. Take the compass reading on this headingand determine the deviation. Now subtract,algebraically, the south heading deviation fromthe north heading deviation and divide theremainder by two.

For example, if the compass reads 175 1/2°while on the south heading (180°0), record this asa deviation of +4 1/2°(180°– 175 1/2°). If thecompass reading is too low, the deviation is plus;if the reading is too high, the deviation is minus.

Suppose that on the north (000°), heading, thecompass reads 006 1/2°. Such a reading is 6 1/2°too high. You would record this as a deviationof –6 1/2° (000° – 006 1/2°).

The next job is to determine the coefficientof north-south deviation. You accomplish this bysubtracting, algebraically, the deviation on thesouth heading from the deviation on the northheading. You then divide the remainder by two.

The aircraft is still on the north heading and thecompass reads 006 1/2°. Since the coefficient ofthe north-south deviation is – 5 1/2°, you mustadjust the north-south compensator by thisamount. The compass reading on the northheading will now be 001°. This adjustment alsocorrects the south deviation by the same amount

(but in the opposite sense). The south heading onthe compass will now read 181°. The coefficientof north-south deviation, which is – 5 1/2° in thiscase, is called coefficient C.

On the loose-magnet type compensator, youadjust north-south deviation by inserting thenecessary number of magnets into the lateral(athwartship) chamber of the compensator. If thecompass has a universal compensator, you makethe adjustment by turning the north-south (N-S)compensator screw.

The next step is to determine the east-westdeviation. Turn the aircraft heading to magneticeast, according to the compass rose. Record thecompass reading on that heading. Now determinethe coefficient of east-west deviation, otherwiseknown as coefficient B.

Assume, for example, that the compass reads2760 when the aircraft was on the west (270°,heading. Also assume it reads exactly 90° on theeast (90°) heading. You find coefficient B byalgebraically subtracting the deviation on west( – 6°) from the deviation on east (O°) and dividingby two.

While the aircraft is on the east heading,adjust the east-west (E-W) compensator toadd 3° to the compass reading. This readingbecomes 93° on the east heading, and thecompass would read 273° on the west heading.Make this adjustment by turning the E-W screwon a universal compensator. On the loose magnettype compensator, add the necessary magnets inthe longitudinal (fore-and-aft) chamber. Leavingthe aircraft on an east magnetic heading, nextcompute an overall deviation correction based oncoefficient A. This coefficient is equal to thealgebraic sum of the compass deviations on allfour cardinal headings (north, east, south, andwest) divided by four.

You must compensate instrument panelcompasses for coefficient A if it amounts to 2°or more in either direction. When making thiscorrection, leave the magnetic compensatorsalone. To compensate for coefficient A, move theinstrument in its mounting.

Compensate panel-mounted compasses forcoefficient A by slightly realigning the whole

Figure 7-59.-Compass correction card.

instrument panel. You can also turn the compassa little with relation to the front of the panel andplacing washers or spacers under its mountingscrews.

After completing this swing, swing the aircraftagain on at least eight equally spaced headings (forexample, 0°, 45°, 90°, 135°, 180°, 225°, 270°,and 315°). Record the compass readings for eachheading on a compass correction card. Figure 7-59shows an illustration of a compass correctioncard.

Detach the small right-hand portion of thecompass correction card and mount with thecompass. It is thus available for ready reference,telling the pilot or navigator the comparativecompass headings and magnetic headings. Turnthe larger portion of the card into maintenancecontrol for insertion into the aircraft logbook.

MC-2 Magnetic Compass Calibrator Set

The MC-2 (fig. 7-60) provides a controlledmagnetic field (simulated earth’s magnetic field)

about the aircraft flux valve to accurately calibratethe compass system. Use of the MC-2 onlyrequires the aircraft be accurately placed on anorth-south line, thus eliminating the need forrotating the aircraft on a compass rose. Thecompass calibrator provides electrical headinginputs from 0° to 345° in 15-degree incrementswith an accuracy of 0.1°. The compass calibratorcan survey an area for magnetic uniformity. Italso provides the necessary data for layout andmarking of a compass swing site.

The compass calibrator consists of four majorcomponents—the control console, magnetic fieldmonitor, remote transmitter turntable, and fieldtester. The set also includes various cableassemblies, reels, racks, tripods, and some specialalignment equipment.

The control console contains the controls,indicators, and electronic components that allowthe compass calibrator set to operate. It uses 115volts, 400 Hz at a maximum of 1 ampere.MC-2 changes this electrical power into acdc voltages that it requires.

Theand

Figure 7-60.-MC-2 magnetic compass calibrator set.

The magnetic field monitor is an engineer’s detector. All connectors, controls, switches, andtransit that has been modified to operate as a partof the compass calibrator set. The modificationconsists of installing a magnetic sensing elementin place of the magnetic compass. The monitoris of nonferrous and nonmagnetic materials. Ithas a telescope, a horizontal circular scale withan adjustable vernier azimuth scale, levels, andleveling adjustment screws. The telescope is22-power with an interior focusing optical system,and rotates 180 ° in a vertical plane.

The remote transmitter turntable is also anengineer-type transit with the compass, verticalcircle, and telescope removed. Also included withthe turntable is a transmitter mounting bracketand a rain hood.

The field tester is a portable metal-encasedtester. It consists of a test panel, a shield canassembly, and a magnetic azimuth reference

electronic parts mount on the test panel. Theshield can assembly contains a valve assemblywithin two magnetic shield cans. The magneticazimuth reference detector consists of a 6-powertelescope with azimuth adjustment and a fluxvalve assembly on a triangular support plate. Thevalve assembly has an attaching cable assembly.

The alignment equipment consists of atelescope, two plate assemblies, shaft coupling,quick connector, plumb bob and adapter, screw-drivers, magnifier, wrenches, and sunshade.The parts you use depend on the aircraft andtransmitter under calibration. The telescope is afixed-focus type, 8-power, with 360-degreeazimuth rotation. A drum dial fine-adjustsazimuth, and an azimuth lock prevents unwantedrotation.

Figure 7-61.-Electrical swing and

The compass calibrator set is used to conductan area magnetic survey. The survey determinesthe size and direction of the earth’s magnetic fieldat a proposed aircraft swing site. You will alsouse the compass calibrator to conduct the actualcompass swing. The control console providescontrolled dc currents for the transmitter. Themonitor detects the size and direction of theearth’s magnetic field and supplies this informa-tion to the control console. You use the alignmentequipment with the turntable to optically align thecompass system transmitter (flux valve).

A review of the flux valve will be helpful inthe following discussion.

In an electrical compass swing, a dc magneticfield is generated in the transmitter and varied insize and direction. This allows the MC-2, incombination with the horizontal component ofthe earth’s field, to simulate an equivalent earth’sfield in the transmitter at a desired heading.Errors in the compass system are measured as thedifference between the aircraft magnetic headingand the simulated earth’s field magnetic heading.The aircraft heading shows on the aircraftcompass indicator. The simulated earth’s fieldheading shows by selecting the HEADINGSELECTOR switch on the control console.

manual

swing at a 90-degree heading.

Controlled dc currents to the secondary coilsthe transmitter generate an electromagnetic

field (electrical swing). By applying a current tocoil leg A of the transmitter (fig. 7-61), yougenerate a field that aligns to leg A. In theelectrical swing, this field provides the north-southcomponent of the simulated earth’s field. A dccurrent also goes through coil legs B and C togenerate two fields, each aligned to its respectivecoil. These fields are so oriented that north-southcomponents of these two fields cancel, leaving oneeast-west component. By reversing the directionof the current flow, you can rotate the east-westcomponent 180°.

The procedures for an electrical compassswing using the magnetic compass calibrator setare as listed below.

1. Set up the turntable over the spot wherethe remote compass transmitter will be when theaircraft is on the north line.

2. Remove the remote compass transmitter(flux valve) from the aircraft and mount it on theturntable.

3. Determine the alignment of the transmitterto magnetic north and its electrical calibration tothe ambient magnetic field. Calibrate the N-S andE-W adjustments on the transmitter.

4. Mount the necessary optical alignmentequipment to the remote compass transmitter.Align the telescope to a predetermined target one-half mile or more away.

5. Tow the aircraft into position exactly onthe north line by using plumb bobs or some otheraccurate method.

6. Compute the optical alignment correction.Insert the correction into the optical alignmentscope. Replace the compass transmitter in theaircraft (sighting on the same target used in step4 above).

7. With the transmitter fastened down,reconnect the leads.

8. Using the appropriate adapter cables,connect the compass calibrator set into thecompass system.

The aircraft magnetic headings are set in withthe heading selector on the control console.Record the errors as the difference betweenthe indicated heading and that set in with theheading selector. Calibrate the compass systemcomponents to within 0.10 of the headingselector position.

For more detailed information on compassswinging, refer to Military Standards, MIL-STD-765A. Consult this specification for additional

information in connection with swinging, com-pensating, and calibrating compasses.

What INS solved the problem of operatingan inertial system at the poles?

Having to maintain the accelerometerreferenced to a fixed point in inertial spaceis a significant disadvantage o

.f the

What is the main advantage of thegeometric inertial navigation system?

What two types of compensators are usedon flux valves?

When must you compensate instrumentpanel compasses for coefficient A?

CHAPTER 8

AUTOMATIC FLIGHT CONTROL ANDSTABILIZATION SYSTEMS

Aircraft fly under many conditions. Externalconditions can alter the desired flight charac-teristics of the aircraft. To maintain the desiredcharacteristics of the aircraft, the pilot moves thecontrol surfaces either manually or automatically.

You have already learned about indicatingsystems and instruments that supply the pilot withinformation on the performance of the aircraft.The pilot must be able to see and interpret eachof these indicators, and then react to getthe desired performance. In high-performanceaircraft, especially in single-piloted aircraft, otherflight duties require much of the pilot’s time.Navigation, communication, radar, and otherspecial equipment is severely limited if thepilot has to work consistently on the physicalmanipulation of the controls.

In high-performance aircraft capable ofsupersonic flight, aircraft speed is so great thatthe pilot’s normal response time is far too slow.By the time the pilot reacts to an indicator toposition a control surface, the aircraft mayalready be out of control.

Automatic flight control and stabilizationsystems ease the pilot’s workload and provideaircraft stability at all speeds. The informationnow flows directly to a flight control computerrather than to an indicator. This action lessensthe time required to start a control movement tonearly zero (increased stability). The systemalso provides command controls by which thecomputer can control the aircraft in nearly anydesired flight condition. Some automatic flightcontrol systems are capable of flying the aircraftto radio navigation aids, correcting for wind, andmaking pilot-unaided landings.

The terms automatic flight control system(AFCS) or automatic stabilization equipment(ASE) are used instead of the older termautomatic pilot, or the shortened version,autopilot. A reliable AFCS is necessary because

pilots have duties other than moving the flightcontrols. However, regardless of how sophisti-cated the AFCS computer may be, the reasoningpower of the pilot cannot be duplicated.

PRINCIPLES OF FLIGHT

Learning Objective: Recognize principlesof flight for both fixed- and rotary-wingaircraft.

To understand automatic flight control andstabilization systems, you must study the effectsthat the various controls have on the aircraft.Airman, NAVEDTRA 14014, contains a basicintroduction to the principles of flight and flightcontrols. You should review this text beforeproceeding with this chapter.

An airfoil is any part of an aircraft designedto produce lift. Obviously a wing is the primaryairfoil on an aircraft; but, propeller blades, tailsurfaces, and even the fuselage itself are importantairfoils. The design of a specific airfoil isdetermined by the job it is to do. All airfoils havethe basic elements shown in figure 8-1.

Figure 8-1.-An airfoil.

An airfoil consists of two nearly parallelsurfaces, with one surface being more roundedthan the other. As air passes over these twosurfaces, the air passing over the roundedsurface has farther to travel than the air passingover the flat surface. However, two particles ofair leaving the airfoil’s leading edge at the sameinstant, one going over the rounded surface andone over the flat, arrive at the trailing edge at thesame time. Therefore, you can infer that airpassing over the rounded surface travels at ahigher velocity than air passing over the flatsurface.

Bernoulli’s theory concerning the behavior offluids, E = V x P, explains how pressure ischanged and lift is produced. Here, E is the totalenergy produced by the airfoil passing throughthe air, V is velocity energy, and P is pressureenergy. An airfoil that passes through air at avelocity of 50 feet per second and exerts a pressureof 10 pounds per square inch on the flat surfaceproduces a total energy of 500 foot-pounds persquare inch per second.

If the airflow velocity over the rounded surfaceis increased to 60 feet per second, and the totalenergy is unchanged, it exerts a pressure of 8.33foot-pounds per square inch per second on therounded surface.

The difference in the pressure between therounded surface and the flat surface of the airfoilis called lift.

In actual practice, the flat surface is notperfectly flat and causes some decreased pressure.The decreased pressure is negative lift. Negativelift is compensated for by the creation of a highpressure on the flat surface. Air packed beneaththe airfoil (dynamic lift) causes the high pressure.The true measure of lift remains the differencein pressure between the rounded and flat portionsof the airfoil.

Increased lift is the result of a larger pressuredifference between the surfaces. The difference

can be produced in two ways—by increasingthe forward movement of the airfoil throughthe air (fig. 8-2), or by changing the angleof attack. Angle of attack is the acute anglebetween the chord line of an airfoil and itsdirection of motion relative to the air. T h echord of an airfoil is an imaginary straightline drawn from the leading edge to the trailingedge of the airfoil (fig. 8-3). As the angleof attack increases, the air strikes the leadingedge closer to the flat portion of the airfoil.The distance air must flow over the roundedportion becomes even greater in relation to thatflowing over the flat portion. This action causesa larger pressure difference, and develops morelift.

If the angle of attack increases too much,airflow over the airfoil’s rounded portionseparates from the surface and becomes turbulent.Turbulence causes the pressure on both surfacesto become nearly equal, and the airfoil is said tostall.

When producing lift, a secondary effect calleddrag is also produced. Drag produced by a liftingsurface or airfoil is called induced drag. Induceddrag develops in direct proportion to lift—whenlift increases, induced drag also increases. At agiven speed and angle of attack, a thick airfoilproduces more lift and drag than does a thin

Figure 8-2.-Lift increases as velocity increases.

Figure 8-3.-Constant velocity verus increasing angle of attack.

airfoil (fig. 8-4). It follows that large, subsonicaircraft have thick wings to produce a greatamount of lift at slow speeds. Supersonic aircraftmust have very thin wings to decrease drag at highspeeds.

Many airfoils have devices attached to themto increase or decrease lift in various flightconditions or attitudes. These devices may mounton the leading edge, trailing edge, roundedsurface, or flat surface. If the trailing edgeattaches by a hinge and has controls to move thetrailing edge, you control lift by changing theangle of attack (fig. 8-5). When the trailing edgemoves into the higher pressure air on the airfoil’sflat side, the angle of attack is effectivelyincreased. This causes more lift and drag.Conversely, if the trailing edge moves into theairfoil’s low-pressure side, the angle of attackdecreases. Lift and drag decrease accordingly.

In flight, each aircraft has certain forces actingupon it (fig. 8-6). To sustain flight at a constantaltitude, the total lift of all airfoils must equal theaircraft weight. To change altitude you must

change the total lift. If the aircraft weighs 10,000pounds, 10,001 pounds of lift causes the aircraftto climb; 9,999 pounds of lift causes the aircraftto descend.

To fly at a constant airspeed, the forces ofthrust and drag must be equal. When one forceis greater than the other, the aircraft acceleratesor decelerates.

To turn, place the aircraft in a bank angle(fig. 8-7). The lift developed by the airfoilscan then be broken down into components ofhorizontal and vertical lift. The horizontalcomponent of lift pulls the aircraft around in theturn. The vertical component of lift must be equaland opposite to gravity for the aircraft to remainat a constant altitude. (Note that total lift mustbe increased to prevent a loss in altitude.)

When centrifugal force equals horizontal lift,the aircraft is in a constant-rate turn. For a fasterrate of turn, increase horizontal lift by increasingthe bank angle. When all lift is vertical to gravity,any turning motion is called a skid.

Figure 8-4.-Induced drag.

Figure 8-5.-Lift and drag change proportionately with the shape of the airfoil.

Figure 8-6.-Forces on an aircraft.

Figure 8-7.-Forces in a turn.

FIXED-WING AIRCRAFT

A fixed-wing aircraft is one in which the mainlifting surface remains stationary with respect tothe rest of the aircraft. This classification alsoincludes such aircraft designs as the swing wingF-14. Fixed-wing aircraft have certain fixedsurfaces or airfoils—wings and vertical andhorizontal stabilizers—that provide stability(fig. 8-8). In addition, these aircraft have movableairfoils—ailerons, elevators, and a rudder. Thesepermit the pilot to control the aircraft.

Movement about the lateral axis of the aircraft(the axis that extends from wing to wing throughthe center of gravity) is pitch. To control pitchyou use the elevators. If you desire a nose-upattitude, apply an aft motion on the cockpitcontrol (stick or yoke). This causes the elevatorto move up, creating a downward force on thehorizontal stabilizer. Rotation about the lateralaxis then causes the nose to rise. Conversely, ifyou want to lower the aircraft nose, apply aforward motion on the cockpit control. Thiscauses the elevator to lower, creating an upwardforce on the horizontal stabilizer. Rotation aboutthe lateral axis then causes the nose to lower.

Movement of the aircraft about the longi-tudinal axis (from nose to tail) is known as bankor roll. You control this motion by using theailerons. The ailerons mechanically connect toeach other, but move in opposite directions. Forthe aircraft to enter a left bank, the angle of attackof a portion of the right wing must increase. Youaccomplish this by lowering that aileron toincrease the lift on that wing. The aileron on theleft wing is raised to decrease the lift on that wing.The aircraft then rotates about its longitudinalaxis until the ailerons are neutralized in some angleof bank. The aircraft remains in that bank angleuntil you again move the ailerons.

Refer to figure 8-7. Whenever the aircraft isin a bank, lift developed by the wings is displaced

Figure 8-8.-Fixed-wing aircraft controls.

from the vertical position. If you do not increaselift, its vertical component is insufficient tomaintain the aircraft at a constant altitude. Achange in pitch attitude to increase the angle ofattack of the wings is used to prevent a loss inaltitude. A few degrees of bank angle requires animperceptibly small pitch change. A 90-degreebank in level flight (no altitude change) istheoretically impossible because of the absenceof vertical lift. As the bank angle changes,coordination between ailerons and elevators isnecessary to prevent a loss in altitude.

To return to level flight, increase lift on theleft wing by lowering its aileron into the higherpressure area beneath the airfoil. Reduce lift onthe right wing by raising its aileron into the lowerpressure area at the top of the airfoil. As the wingsbecome level, neutralize the ailerons.

Movement about the vertical axis is yaw.Usually this movement is undesirable in anaircraft. Use the rudder to correct any tendencyof the aircraft to yaw. The rudder is NOT usedto turn the aircraft (change heading).

When placing the ailerons into the airstream,the aircraft has a tendency to yaw. In banking tothe right, the aircraft produces more lift and dragon the left wing, and less lift and drag on the rightwing. Even though the intention is to turn to theright by going into a right bank, the initialtendency is for the nose of the aircraft to go tothe left. This happens because of the increaseddrag on the left wing and decreased drag on theright wing. This is adverse yaw; you compensatefor it by displacing the rudder in the samedirection as the intended turn.

If an aircraft in a turn tends to slip into theinside of the turn or skid to the outside of the turn,this is also yaw. You also compensate for it byusing the rudder. Many other things may causeyaw, such as the engines on one wing of amultiengined aircraft producing more power thanthe engines on the other wing.

You can see, then, when you place a fixed-wing aircraft in a bank angle, coordinationbetween all three controls—ailerons, elevators,and rudder—is necessary. The pilot accomplishescontrol of elevators, ailerons, and rudder throughthe use of a control stick and rudder pedals(fig. 8-9). To operate the ailerons, move thecontrol stick right or left in the direction of theintended turn (fig. 8-9, view A). Aft force on thecontrol stick raises the elevator and causes thenose to pitch up. Forward pressure on the controlstick lowers the elevator and causes the nose topitch down (fig. 8-9, view B). You use your feet

Figure 8-9.-Flight controls. (A) Elevator and aileron;(B) Rudder.

to operate the rudder pedals. Pressure on eitherrudder pedal causes rudder deflection in thatdirection.

Weight distribution in an aircraft varies formany reasons. For example, fuel may be usedfaster from one wing tank than from the other,allowing that wing to become lighter. In largeaircraft where crew members or passengers walkaround, the balance point, called the center ofgravity (CG), shifts whenever someone changesposition in the aircraft. As fuel is used, the aircraftgross weight reduces. The pilot must reduce theangle of attack of the wings to lessen lift andprevent a gain in altitude.

The pilot must use control pressures tocompensate for these unbalanced flight condi-tions. Several methods are used to reduce these

control pressures and to ease the pilot’s workload.The most common method is the trim tab. Lookat figure 8-8. The figure shows only rudder andaileron trim tabs. Trim tabs on the elevators onthis particular aircraft are undesirable becausethey produce excessive drag. Another method ofelevator trim involves linkage pressure.

When pilots must exert a force on the cockpitcontrol, they can use the trim control to relievethat force. For instance, when they must hold leftrudder pressure to prevent yaw movement to theright, they can move the rudder trim tab to theright. The airflow on the vertical stabilizer strikesthe trim tab and moves the complete rudder a littleto the left. Since the rudder trim tab now suppliesthe required rudder pressure, the pilot no longerhas to hold pressure on the rudder pedals.

ROTARY-WING AIRCRAFT

An aircraft that derives its main lifting forcefrom a horizontally driven propeller device (rotor)is a rotary-wing aircraft. The most commonrotary-wing aircraft is the helicopter. There mustbe relative motion between an airfoil and an airmass. Therefore, the major advantage of a rotary-wing aircraft is its ability to maintain zero or verylow airspeed while the wings (rotors) are stillcreating lift.

Forces acting on a rotary-wing aircraft areidentical to those acting on a fixed-wing aircraft(fig. 8-6). You must also control the rotary-wingaircraft about the vertical, longitudinal, andlateral axes, as in figure 8-8.

In the conventional helicopter, the main andtail rotor are engine driven. Remember the earlierdiscussion on airfoils. You increase lift by eitherincreasing the speed of the airfoil through the airor by increasing the angle of attack of the air foil.In helicopters, the air foil’s angle-of-attack isknown as blade pitch. When the rotor speed isconstant, the pilot maintains complete control ofthe aircraft by varying the pitch of the rotorblades.

Figure 8-10 shows helicopter flight controls.The pilot operates collective control with the lefthand, cyclic control with the right hand, andrudder control with the feet. The collectiveand cyclic controls command the main rotor.Operation of the rudder control changes the bladeangle of the tail rotor.

In helicopter flight (except hovering flight),the main rotor provides altitude, bank, anddirectional control through use of the collectiveand cyclic controls. The tail rotor prevents the

Figure 8-10.-Helicopter flight controls.

main body of the helicopter from spinning(yawing) with the torque of the main rotor. Itprevents yaw in much the same way as thefixed-wing aircraft rudder.

Collective control maintains or changesaltitude. Moving the collective control causes anequal change in pitch (angle of attack) of all mainrotor blades. Also, through a mechanical mixer,it automatically changes tail rotor pitch. Thecollective control changes tail rotor pitch tocompensate for increases or decreases in mainrotor torque. Since the rotor blades are somewhatflexible, the more collective control applied, themore an action called coning takes place. As theblades rotate, they take the shape of a cone(fig. 8-11). The speed and pitch of the blade tipsdetermines the coning angle. With a constant

Figure 8-11.-Coning angle increases as load increases.

pitch, the faster the rotor blades turn, the morehorizontal the blades become because of centri-fugal force. As the blade pitch increases, lift alsoincreases, and the coning angle increases becauseof the load on the blades.

In hovering flight, cyclic controls pitch androll, which create forward and sideward motion,respectively. The collective controls altitude andthe rudder pedals control heading.

The cyclic stick provides pitch and directionalcontrol of the helicopter. When the pilot appliespressure to the cyclic stick, each blade moves toa specific pitch angle as it passes a certain point

in its rotation (fig. 8-12). To accomplish forwardflight, the blade pitch is greatest as it passes the90-degree position. The blade pitch will be leastat the 270-degree position, and equal at the0-degree and 180-degree positions. To turn theaircraft, lateral motion of the cyclic stick causesblade pitch to be greatest at 0 and 180 degrees,and least at 90 and 270 degrees. Since the bladesform a spinning mass, the gyroscopic principle ofprecession occurs 90 degrees in the direction ofrotation from where the lifting force is applied.The coning angle remains the same. However, thecone tilts (called flapping angle) in the direction

Figure 8-12.-Flapping angle creates horizontal lift. (A) Hovering flight; (B) Forward flight.

of the desired flight path. You can again breaklift down into its vertical and horizontalcomponents. Look at figure 8-12, view B. Tomaintain altitude, the vertical lift is increased untilit is equal to gravity. With the cone at a flappingangle, the helicopter acceleratesdirection until drag is equallift.

To accelerate the helicopterdirection, move the cyclic control

in the desiredto horizontal

in a forwardstick forward.

You also must make a corresponding increase incollective to maintain altitude. As the collectiveincreases, torque on the main rotor bladeincreases. This makes the helicopter tend to rotatein the direction opposite to the rotor bladerotation (nose right). A mechanical mixerautomatically changes the pitch of the tailrotor to overcome the right turning tendency(skid).

To turn the helicopter (change heading), placethe cyclic control stick to the right or left.Flapping action of the main rotor blades causesthe cone to tilt in the direction of the desired turn.As in the fixed-wing aircraft, the pilot mustmaintain coordination in a turn by using therudder pedals to prevent skid or slip. Also, thepilot must adjust collective to prevent a loss inaltitude. In hovering flight, the pilot uses therudder pedals only to turn the helicopter, thusproducing a skid.

On a fixed-wing aircraft, what controlsurface corrects for yaw?

As fuel is used, what must the pilot reduceto lessen lift and prevent a gain in altitude?

How does the pilot increase lift in ahelicopter with a constant speed rotor?

When does the action called coning takeplace?

AUTOMATIC FLIGHT CONTROLSYSTEMS (AFCS)

Learning Objective: Relative to the auto-matic flight control system, recognizefunctions and operating principles andmodes, including air data and flap positioninformation and coordination inputs, andidentify AFCS components.

In the human body, signals to move us fromplace to place start with our five senses as theyreference outside conditions. The brain processesthese signals and sends them through the nervesto the muscles. The body then does its requiredmovement by muscle power. Similarly, mostautomatic flight control systems (AFCS) havetheir component parts divided into three majorgroups—sensors (information inputs), amplifier/computer, and output units.

The sensors originate the signals as they areacted upon by outside references. They only sensechanges and do not have sufficient power to makecorrections.

The amplifiers and computers are the brainsfor the AFCS. They receive the weak signals fromthe sensors, which in most cases are synchros, anddetermine how much and in which directioncorrection is necessary. The synchro signals areusually in millivolts, but the correct strengthneeded is in volts. Therefore, the amplifierincreases the weak signal to a workable voltage.The value of the synchro signal depends on theamount of rotor displacement with respect to thestator from the null position. The direction ofrotor displacement from the stator determines thedirection of the correction. Most amplifiers haveat least two stages of voltage amplification—onestage of phase discrimination, and an amplifierwhere power amplification takes place, Othertypes of amplifiers control the voltage to controlvalves in hydraulic servos.

The output unit is the muscle of the AFCS.It consists of an electro/hydraulic boosterpackage. There is a booster package for eachcontrol surface—rudder, aileron, and elevator.The boosters also assist the pilot in manual controlof the aircraft.

Summing up the major groups, the sensorssend a small signal to the amplifier/computerwhen a displacement occurs. The amplifier/computer amplifies the weak signal to a workablevoltage and sends it to the output unit. The outputunit changes the electrical energy to mechanical

displacement. It then moves the control surfacesby an amount commanded by the sensor signal.

AFCS COMPONENTS

The automatic flight control system (AFCS)consists of many controls, sensors, and electro-mechanical components. To understand the entiresystem, you need to know what each componentof the system does. In this section, you willread about the components that make up theAFCS.

Electrical/Electronic Components

The electrical/electronic components makeautomatic flight control systems work. Thecontrol panel is used to program any pilot-desiredmaneuver that is within the capability of thesystem. Originally, AFCS systems were verylimited. They supplied only one-channel operationto the ailerons to keep the wings level. Neweraircraft receive signals from other aircraft systems.Radar and barometric altimeter signals couplewith the AFCS to maintain the aircraft at aconstant altitude. Some aircraft use signals fromdata-link systems to fly the aircraft duringapproaches and landings. Some fighters have thefire control system tied in so the aircraft can flyautomatically to an enemy aircraft. Fighterbombers with a weapons control system tie-in canfly automatically to the target and release theirweapons at the proper time. Long-range patrolaircraft have their ASW systems tied into theirAFCS. The AFCS operates the rudder, theelevator, and the ailerons by using various sensorsand electrically controlled hydraulic servos. Beforeengagement, the AFCS is synchronized with theflight control surfaces to prevent sudden or violentmaneuvers.

The system senses deviation from the referenceflight condition and causes the aileron control tomaintain either a reference bank angle or aheading. It also causes the elevator control tomaintain either a reference pitch angle or analtitude. Also, it makes the rudder controlcoordinate turns and provides automatic yawdamping.

CONTROL PANEL. —The AFCS controlpanel contains all the switches and controlsnecessary for the pilot to select/control the

autopilot modes. Control panels are designed forthe particular type and mission of the aircraft.Some control panels are simple, while others arecomplex. Figure 8-13 is an example of a AFCScontrol panel. Here, the switches serve asmanually operated interlocks in setting up thecircuitry to engage the various AFCS modes ofoperation. This control panel has six switches.They are labeled as follows: ACL/OFF/PDC,ALT/OFF/MACH, HDG OFF/NORM/ROLLCMD, AUTO/STAB-AUG, ON/OFF, andATTITUDE REF.

The ACL/OFF/PCD, ALT/OFF/MACH,and HDG OFF/NORM/ROLL CMD switchesare solenoid-held toggle switches. Each has alever-lock toggle feature that prevents accidentalengagement in the operate position. The AUTO/STAB-AUG and ON/OFF switches are solenoid-held, spring-loaded switches. When not engagedor with no power applied, the switches return tothe STAB-AUG and OFF positions, respectively.The ATTITUDE REF switch is a miniature,positive-break, aircraft-type toggle switch.

AIR NAVIGATION COMPUTER (ANC). —All operating functions of the automatic flightcontrol systems channel through the air navigationcomputer. It is sometimes called the amplifiercomputer, and it is the heart of the entire system.

The ANC modifies the combined signalssupplied by various sensors and command con-trols to develop output signals. The output signalscontrol the aircraft’s ailerons, rudder, andelevators. By use of these flight controls, theaircraft automatically maintains a referenceattitude, heading, and altitude. Also, it canmaneuver in a coordinated manner in responseto turn and pitch control settings on the controlpanel.

Figure 8-13.-AFCS control panel.

Figure 8-14.-(A) Air navigation computer; (B) One-channel amplifier/computer.

The physical appearance of the ANC dependson the type and mission of the aircraft. Figure8-14 shows two different types of ANC. View Aof figure 8-14 shows an ANC that consists of anequipment rack and seven amplifier modules,which are listed below.

1. Roll servo amplifier2. Pitch servo amplifier3. Yaw servo amplifier4. Roll computer amplifier5. Pitch computer amplifier6. Heading computer amplifier7. Command coupler

Each of the seven modules contains sub-assemblies and sub-subassemblies. Some of theseare interchangeable between modules. The roll,pitch, and yaw servo amplifiers are identical.The other modules have individual differences.The computer, through an interlocking relayarrangement in conjunction with the control panelmode selection switches, controls signal switchingoperations. A calibration board on the front ofthe ANC provides gain adjustments of the majorsystem parameters.

View B of figure 8-14 shows a one-channelamplifier/computer. Normally, this particulartype of autopilot computer consists of threeindividual amplifier/computer modules—one foreach control surface—aileron channel (roll),rudder channel (yaw), elevator channel (pitch).

This one-channel computer accomplishesanalog computations by using servomechanisms.These servomechanisms consist of electro-mechanical computer cards and electronicamplifier cards mounted in the amplifier/computer. In addition, a transformer board anda resistor board provide summing networks. Thenetworks combine the various signals supplied toand generated within the unit. An interlockingrelay arrangement is included to perform most ofthe switching control in the automatic flightcontrol system.

CONTROL STICK. —Control stick or controlwheel steering is used on some aircraft to controlthe aircraft electronically through the AFCS,using the regular control stick or control wheel.On fighters, the signals are generated in a unitsuch as the one labeled “motional pickuptransducer, ” in figure 8-15.

When the AFCS is on and the control stickis moved left or right, pressure on the roll forceswitch momentarily disengages the roll channelof the AFCS. The pilot then controls the roll

Figure 8-15.-Control stick steering components.

attitude of the aircraft through regular stickcontrol. When the pilot releases stick pressure, theforce switch opens, allowing the AFCS toreengage roll. If the bank angle is above a givenangle (for example, 5 degrees), the AFCS willmaintain the bank angle. If the bank angle isbelow the given angle, the AFCS automaticallyreturns to wings level.

The pitch force switches close when a foreor aft pressure is on the control stick. Thismomentarily disengages the AFCS. The stickpressure also couples a signal through the Epickoff transformer that is labeled “forcesensor,” as shown in figure 8-15. The signalcouples with the AFCS pitch channel. Dependingon the direction of the stick pressure (fore or aft),the aircraft either climbs or dives.

Electrical/Electronic Sensors

Many electrical and electronic sensors provideinput to the AFCS. This section of the TRAMAN

includes a review of the sensors already discussed,and it introduces you to other sensors in theAFCS.

SIGNAL GENERATOR PICKOFF (SYN-CHRO). —Figure 8-16 illustrates the principal ofoperation for a signal generator pickoff. The pick-off consists of a stator and rotor. The stator isring-shaped and has four poles. Each pole has aprimary and secondary winding. The rotor has nowindings. It serves to change the reluctance of themagnetic flux path between the stator poles. Theprimary and secondary windings are connectedso the voltages induced into the secondaries areof opposite polarity on adjacent poles. However,opposite poles have the same polarity.

The voltage output of the secondary is zeroif the rotor is in its neutral position (fig. 8-16, viewA). Repositioning the rotor (fig. 8-16, views B andC) makes a stronger magnetic field on a single pairof poles. This results in a voltage output on thesecondary winding. The amplitude of the outputvoltage is proportional to the amount of rotordisplacement—the greater the displacement, thegreater the amplitude. The direction of rotormovement determines the polarity of the outputvoltage. The polarity will either be in phase withthe input voltage or 180 degrees out of phase withit.

Synchro construction allows accurate voltageproduction versus angle signal, effective through360 degrees of rotation. For more information

Figure 8-16.-Signal generator pickoff operation.

about synchros, servos, and gyros, you shouldrefer to Navy Electricity and Electronics TrainingSeries (NEETS), module 15, Principles ofSynchros, Servos, and Gyros.

GYROS. —There are several different gyrosused with the AFCS. You have already learnedabout most of them earlier in the TRAMAN. Thefollowing paragraphs provide a review of gyrosand how they specifically affect the AFCS.

Vertical Gyro. —The vertical gyroscope(fig. 8-17) is an electrically driven gyro that

provides pitch and bank attitude references forthe AFCS. It can also provide pitch and bankattitude references for servo indicators and othersystems of the aircraft. It has enough signal loadcapacity to sustain several systems at the sametime.

The vertical gyro is a two-degree-of-freedomgyro. This gyro is termed the vertical gyro becauseit is continuously erect with its spin axis verticalto the surface of the earth. The spin axis providesa vertical reference for measurement of aircraftbank angle and pitch angle. Pitch and roll gimbals

Figure 8-17.-Vertical gyro components.

isolate the gyro from its housing and the aircraft.Thus, the aircraft can bank or pitch while the gyroremains vertical because of gyroscopic action(fig. 8-18). A pitch synchro, mounted on thevertical gyro’s pitch pivot, centinuously senses therelative pitch angle between gyro and aircraft.Similarly, a bank synchro, mounted on thevertical gyro’s bank pivot, continuously senses therelative bank angle between gyro and aircraft.

The gyro motor rotates at a speed of about20,000 RPM. A solenoid-operated friction brakeprevents tilting and tumbling when the gyro motoris idle and during the initial starting torque.

Synchros mounted on the gyro detect motionbetween the gyro, the gyro gimbal, and the gyrocase. A synchro generates a weak signal when itsrotor is displaced from its stator. The pitchsynchro mounts with its rotor on the gyro pivotand its stator on the gyro gimbal. Thus, thesynchro measures the displacement angle betweenthe gyro and the gimbal. This is the pitchdisplacement angle from the vertical reference.

The bank synchro mounts with its rotor on thegyro gimbal pivot and its stator on the gyro case.Thus, the synchro measures the displacementangle between the gimbal and the case. This isthe bank displacement angle from the verticalreference.

As the pitch attitude of the aircraft changes,the gyro case and gimbal turn about the gyrorotor. The pitch synchro rotor is held rigid inspace by the gyro. This generates voltages in thepitch synchro proportional to the aircraft’s pitchangle with respect to the surface of the earth.During changes in pitch attitude, the bank synchroremains at null since the gyro gimbal is not freeto rotate with respect to the case in pitch.Therefore, it tilts with the case.

As the bank attitude of the aircraft changes,the bank synchro stator turns about the banksynchro rotor. The gimbal is held rigid in spaceby the gyro because it is not free to rotate withrespect to the gyro in bank. Thus, the banksynchro generates voltages proportional to the

Figure 8-18.-Vertical pitch and roll reference.

bank angle between the aircraft and the surfaceof the earth.

When the aircraft yaws, the case, the gimbal,and the gyro stator turn about the rotating gyromotor. This has no significant effect upon therelative positions between the gyro and thecase. As a result, there are no pitch and banksynchro output voltages in response to changesin yaw.

Three-Axis Rate Gyroscopes. —Rate gyrossense the rate of movement of an aircraft about

its vertical, lateral, or longitudinal axis. Theyprovide synchro signal outputs representing yawrate, pitch rate, or roll rate to the air navigationcomputer. These units are sometimes very similarin appearance to the rate switching gyro, but theyprovide entirely different information to thesystem. Physically, rate gyroscopes are the same.They differ only in respect to calibration, align-ment, range, sensitivity, and natural frequency.Each gyro measures angular rate. It uses theproportional precessional torque generated by therate of movement about the gyro-sensitive axis(fig. 8-19) to make these measurements.

Figure 8-19.-Three-axis rate gyro orientation diagram.

Internally, each rate gyro consists of a smallviscous-damped, single-degree-of-freedom gyrowith a differential transformer pickoff (fig. 8-20).The gyroscopic element of each gyro is the rotorof a synchronous motor. The rotor mounts in agimbal frame and spins at high speed about itsspin axis. The gimbal is flexible and free to rotateabout an output axis. This axis is perpendicularto both the spin axis and the input axis. Atorsion-restoring spring that couples the gimbalto the case limits the rotational freedom about theoutput axis. The gyro gimbal carries the pickoffrotor on an extension along its output axis. Thepickoff rotor senses the relative angulardisplacement of the gimbal and case.

With the pickoff rotor in its zero or neutralposition, the mutual inductance is zero. Thecurrent flowing in the pickoff primary causesessentially no voltage in the secondary (output)winding. As the pickoff rotor is turned one wayor the other about its output axis by gyrogimbal deflection, a proportional mutualinductance is introduced. The polarity of theinductance (positive or negative) depends uponthe direction of deflection from the neutralposition. Hence, the current flowing in theprimary produces a voltage proportional to thismutual inductance in the pickoff secondary. Theoutput voltage is proportional to the aircraft’sangular velocity input to the gyro in the particularaxis.

COMPASS INFORMATION. —Normally,compass information for the AFCS is supplied bythe aircraft compass system or the inertialnavigation system (INS). However, some compassinformation is developed for the AFCS.

The compass system/INS incorporates a geartrain to drive several synchros. The gear train isdriven by a motor generator unit that aligns toaircraft heading. Of the several synchros, oneprovides heading information to the pilot’s compassindicator. One synchro attached to the gear trainthrough a clutch provides a clutched heading.When the AFCS is not engaged, the clutchremains de-energized, wit h its rotor spring loadedto an electrical null condition. When the pilotengages the AFCS, the clutch engages the engagedheading to establish a reference heading for thesystem. If the aircraft drifts off heading, the geartrain drives against the spring tension on the rotorgenerating an electrical signal. This signal goes tothe aileron channel, much like the signal generatorpickoff operation (fig. 8-16). Limiters in theAFCS prevent the bank angle from becomingexcessive when large heading errors are detected.

Another type of compass information isderived from the heading indicator in the cockpit.The pilot selects a desired heading on the face ofthe indicator. The difference between the selectedheading and the actual heading becomes an errorsignal to the AFCS. This error signal causes theaircraft to turn to the desired heading.

Figure 8-20.-Rate gyro axis orientation.

A radio navigation aid can also supply headinginformation to the AFCS. If the pilot desires tofly to a selected ground station, the radio recieverdevelops a signal to produce the desired groundtrack directly to the station.

AIR DATA INFORMATION. —Changes inthe speed of an aircraft also affect the effective-ness of the control surfaces. At a given altitude,slow speeds require more control surface move-ment than high speeds to accomplish the samemaneuver. The pilot maintains (or changes) thealtitude by reference to the altimeter. The AFCScan also maintain a constant altitude. Toaccomplish this task the AFCS uses altitude datasupplied by an air data sensor or air datacomputer (ADC) as the reference altitude.

Airspeed. —Control surface signals aremodified by a gain control unit to compensate forchanges in airspeed. This unit uses the differencebetween ram pressure and static pressure. Amechanical schematic of the gain control unit isshown in figure 8-21. Here, you can see that asairspeed increases (ram air pressure increases), thebellows causes the spring to become morecompressed. This allows the armature to moveeach potentiometer’s sliding arm to modify thecontrol surface signals an amount representativeof the change in airspeed of the aircraft. Whenairspeed decreases, the armature moves to theright, selecting a different amplifier gain. Theopposite occurs for increases in airspeed.

Altitude. —The AFCS includes an altitudecontrol feature to maintain the aircraft at a fixedaltitude. The altitude controller consists of ananeroid, a mechanism for transmitting andmagnifying the motion of the aneroid, and asolenoid-operated clutch. It also includes asynchro transmitter and a centering device forreturning the synchro transmitter rotor to the null

Figure 8-21.-Mechanical schematic of a gain control unit.

or no-signal position. Some aircraft do not usean altitude controller. In place of an altitudecontroller, the AFCS uses signals from the ADC.

Figure 8-22, view A, shows a three-quarterview of a barometric altitude control. The outside

Figure 8-22.-(A) Barometric altitude control; (B) Internalparts; (C) Simplied schematic.

appearance of various controls of this type varies,depending upon the manufacturer; however, theworking parts are similar. Figure 8-22, view B,shows the internal parts of a barometric altitudecontrol, and view C shows a simplified mechanicalschematic.

The aneroid consists of two diaphragms sealedinternally at standard (sea level) barometricpressure. The two diaphragms connect in tandemto a single pushrod, and are mounted in an airtightcase (fig. 8-22, view B). A tube connects the caseto a source of static air pressure. The diaphragmpushrod mechanically links to one of the clutchplates. This linkage consisting of a lever, a pivotedshaft to which a sector gear is attached, and apinion gear.

When the aircraft deviates from the baro-metric pressure altitude to which the altitudecontrol switch is set, the aneroid diaphragmsmove. This motion is transmitted through thelinkage to displace the rotor of the synchrotransmitter. This displacement generates a signalin the synchro transmitter stator. The signal isapplied to the elevator channel to return theaircraft to the pressure altitude indicated by theaneroid. When the aircraft reaches the correctaltitude, the synchro transmitter signal becomeszero, and normal AFCS operation resumes.

When the altitude control switch is off, themagnetic clutch and the centering device actuatingcoil de-energizes. This opens the clutch todisengage the synchro transmitter rotor from theaneroid mechanism. The centering yoke, by springaction, closes on the synchro transmitter rotorshaft lever to return the rotor to the nullor no-signal position. In this way, the synchrotransmitter rotor is always at the no-signalposition when altitude control is not selected.Since the clutch is disengaged, the aneroid is freeto move. This allows the pilot to engage thealtitude control at any time. Regardless of theaneroid position, the altitude that it sensesis the one used as the reference altitude. Itis not necessary to wait for synchronization oralignment.

FLAP POSITION INFORMATION. —Whenthe flaps are lowered on some aircraft, theincreased lift causes the aircraft to gain altitude(normally called ballooning). Ballooning isundesirable, and it is counteracted by usingnosedown pressure on the flight control. Whenthe AFCS is engaged and the flaps are lowered,automatic nosedown force is applied to the

elevator. Flap position is detected by the use ofa flap position transmitter.

The flap position transmitter consists of twosynchro transmitters with a single input shaft(fig. 8-23). The synchro transmitters supply flapposition information to the AFCS elevatorchannel and the external flap position indicator.

ACCELEROMETER TRANSMITTER. —Thenormal accelerometer (fig. 8-24, view A) generatesa signal proportional to normal vertical accelera-tion. This signal is used for altitude or Mach holdvertical path damping and as the g-commandreference. The unit consists of a cast housingassembly, a sensitive element assembly, bellows,and calibration resistors (R51, R52, and R53). Thesensitive element assembly has an E pickoff, anarmature and armature support, flexure springs,and a backplate.

When assembled, the sensitive elementassembly and bellows are sealed inside thehousing, which is filled with damping fluid. Themetal bellows allow the volume of the dampingfluid to change with temperature and pressurevariations. The damping fluid provides viscousdamping during motion of the armature.

The sensitive element is mechanically biasedto produce zero output when mounted in thecorrect position and subjected to the normalgravity force of 1 g.

Figure 8-23.-Flap position transmitter and schematic.

While you read this section, refer to figure8-24. As the aircraft accelerates in the sensitivevertical direction, the suspended armature tendsto remain behind due to its inertia. This reactionvaries the reluctance of the magnetic circuit setup by the E pickoff windings and armature (viewB). The armature completes the magnetic circuitthrough a small air gap. The relative motionbetween the armature and E pickoff varies thereluctance through the signal output windings.This variation results in a signal that isproportional to acceleration. When operating, theoutput voltage is either in phase or 180 degreesout of phase with the excitation, depending onthe direction of acceleration.

COORDINATION INPUT. —In some air-craft, a dynamic vertical sensor detects lateralaccelerations (slip or skid) of the aircraft. Thesensor supplies a signal to position the rudder tocorrect the slip or skid, coordinating the turn.The signal is proportional to the amount of theaircraft deviation from the vertical axis of theaircraft. In other aircraft, a horizontally mountedaccelerometer aligned with the lateral axis of theaircraft provides the same information. Onlythe dynamic vertical sensor is covered in thisTRAMAN.

The dynamic vertical sensor consists of aviscous-damped pendulum mechanically con-nected to the rotor shaft of a transmitter synchro.The cutaway view of the sensor (fig. 8-25) showsthe mechanism assembly, which includes asynchro transmitter with a pendulum and vaneassembly attached to the rotor. The vane movesin an oil-filled chamber. The damping effect of

Figure 8-24.-Accelerometer transmitter.

Figure 8-25.-Cutaway view of a dynamic vertical sensor.

Figure 8-26.-Dynamic sensor pendulum positions: (A) Coordinated turn; (B) Slip; (C)Skid.

the fluid gives a long-term sensing characteristic A pin in the housing fits into a slot inthat makes the unit relatively insensitive to the mechanism shell. This positions the mecha-transient oscillations. The damping chamber also nism to give proper alignment of the synchrolimits displacement of the pendulum to 10 degrees rotor with the longitudinal axis of the aircrafteither side of the center position. when installed.

The sensor functions in the same manner asthe ball in a turn-and-bank indicator. The ballgives a visual indication of slip or skid resultingfrom lateral acceleration. The sensor provides asignal output of this condition. Figure 8-26 showsa diagram of the forces acting on the aircraft ina turn. Refer to this figure as you read this section.In a coordinated turn, the vertical and lateralforces resolve into a vector perpendicular to thespan of the aircraft. When the aircraft is turningwith the two forces in balance, the ball is centered;the pendulum in the sensor gives a null output.

When the aircraft bank angle is too large forthe turn rate, the balance is upset (fig. 8-26, viewB). The ball moves away from the center towardthe inside of the turn, and the pendulum movesthe synchro rotor from the center null position.The rotor displacement produces a signal withmagnitude proportional to the displacement angleand signal polarity corresponding to the directionof displacement. The unbalanced condition resultsfrom a sideways accelerating force, causing theaircraft to slip toward the inside of the turn.

When the aircraft is insufficiently banked forthe turn, an acceleration acts toward the outsideof the turn (fig. 8-26, view C). The ball in theturn-and-bank indicator moves from the center,and the pendulum in the dynamic vertical sensoris displaced from null in the direction cor-responding to the ball. This gives a signal whosepolarity is opposite to that of the signal when theaircraft was in a slip. The signal is fed to therudder channel for right or left rudder tocoordinate the turn. Since the pendulum isunaffected by transients, the rudder adjustmentis on a comparatively long-term basis.

Hydraulic Components

Hydraulic systems provide the physical powerto move the flight control surfaces. All electro-hydraulic (AFCS) actuators work in the samemanner. A brief discussion of electrohydraulicsis presented in the following paragraphs.

ELECTROHYDRAULIC SERVO ACTUA-TORS. —Electrohydraulic servo actuators(hydraulic booster packages) are discussed inchapter 4 of this TRAMAN. Modern aircraft useseveral types of actuators, with each boosteractuator having at least two modes of operation—manual mode and electrical signals.

The first mode is the manual mode. Itsprimary purpose is to aid the pilot in manuallypositioning the control surfaces. Control surfaceson large aircraft are much too large to moveunaided. On smaller high-speed aircraft, the highair pressure makes it nearly impossible to movethe controls unaided. In the manual mode ofoperation, the hydraulic booster package isconnected between the pilot’s control stickand the control surface. It provides hydraulicassistance to the pilot in much the same mannerthat power steering aids the driver of a car ortruck.

Mode two of the hydraulic booster packageuses electrical signals from the AFCS to move theflight control surfaces. In this mode, the boosterpackage connects the AFCS to the control surfaceand provides the muscle to move the surface.Synchro devices on the boost package providefeedback signals to the AFCS (fig. 8-27).

The surface position transmitter sends theAFCS a signal representing the amount anddirection of control surface displacement from thestreamline position. This signal acts as a follow-up to prevent overshoot of the controls. Also,it serves to return the control surface to thetrimmed condition as the original signal returnsto zero.

The modulating piston is displaced only whenthe control surface is in motion. The position ofthe modulating piston is monitored to sense thecontrol surface rate of movement. This actiongenerates a signal to dampen control surfacemovement.

Hydraulic load sensors determine the amountof pressure the AFCS is applying to the controlsurface so the pilot can properly trim the aircraftbefore disengaging the AFCS. If the aircraft isnot properly trimmed and the control pressure issuddenly relieved, the control surface movesrapidly, causing sudden aircraft movement.

Some aircraft use both manual and AFCSmodes simultaneously. Aircraft stabilization isprovided from the AFCS, while the pilot manuallycontrols the aircraft. All flight control systemshave a method of disconnecting the boosterpackage. This method gives the pilot manualcontrol of the flight control surfaces if the boostermalfunctions or failure of the hydraulic systemoccurs.

AUTOMATIC TRIM. —Some AFCS systemsincorporate automatic trimming. When a signalis present in the control channel, there isan unbalance in fluid pressure at the inputto the hydraulic booster. The hydraulic loadsensor (fig. 8-28) detects this unbalance. Itssignal is amplified and drives the trim servo-motor. The engagement clutch engages whenthe AFCS is operating and the automatictrim system is functioning properly. The slipclutch allows the pilot to override the auto-matic trim in case of a malfunction. MostAFCS use autotrim in the pitch channel;however, it can be used for yaw and roll aswell.

THEORY OF OPERATION

The AFCS has three main control channels tocontrol movement of the aircraft about its axis.These channels control the yaw (rudder), roll(aileron), and pitch (elevator). Each of thecontrol channels has similar functional equipmentgroupings. The groupings include controls,sensors, signal coupling circuits, AFCS servoloops, and aircraft flight controls.

Each channel of the AFCS supplies control sig-nals to the flight controls and receives error signalsfrom the sensors. The error sensors and the signal cou-pling circuits in use in each control channel dependon the mode of operation. The AFCS servo loops

Figure 8-27.-Hydraulic booster block diagram.

Figure 8-28.—Automatic pitch trim block diagram.

and the aircraft flight control system provide the final NOTE: The AFCS channel shown in figureamplifying link to move the flight control surfaces. 8-29 represents the basic design of the channel and

Look at figure 8-29. Each AFCS control is for instructional use only. For further study ofchannel is basically the same; however, its design the AFCS, you should refer to the MIM for yourdepends on the type and mission of the aircraft. particular aircraft.

Figure 8-29.-Basic AFCS control

channel.

The amplifier/computer (ANC) receivresvarious error and control signals. These weaksignals are coupled (summed), modified, andamplified to develop the control surface commandsignal to drive the electrohydraulic actuators. Theactuator moves a certain direction for a specificdistance, depending on the command signal’spolarity and magnitude.

The flight control surfaces (rudder, aileron,and elevator) are mechanically linked to theelectrohydraulic actuator. As the control surfacemoves, a position transmitter synchro developsa feedback signal having the opposite polarity tothe error signal. The magnitude of the feedbacksignal increases as the control surface displace-ment increases. When the error signal andfeedback signal are equal and opposite inmagnitude and polarity, the control surface willno longer move.

Movement of the control surface causes theaircraft to displace about its axis (or reference).This movement corrects the original error signalas sensed by the sensors. Without any errorsignal inputs to the control channel of theamplifier/computer, only the feedback signal ispresent. The feedback signal is opposite in polarityand magnitude to the original error signal. Thisdrives the electrohydraulic actuator an equaldistance in the opposite direction and brings thecontrol surface back to the null or referenceposition.

AFCS MODES

This part of the chapter contains a briefdescription of the various modes available in atypical automatic flight control system. The pilotselects one of these modes on the AFCS controlpanel by moving the control stick or using knobson some instruments. Because the circuitry iscomplex and varies among the different weaponssystems, no specific mode will be diagramed.

Stability Augmentation Mode

The stability augmentation (STAB AUG)mode provides improved control of the aircraftby automatically damping oscillations about thepitch, roll, and yaw axes. Signals from rategyroscopes command control surface movementthrough electrohydraulic actuators. In somehigh-speed aircraft, STAB AUG is consideredcritical to safe flight. For this reason, the STABAUG engagement switch connects in series withall other modes of the AFCS. This arrangement

ensures that STAB AUG is engaged before anyother mode of the AFCS.

Attitude Hold

The attitude hold mode is the basic, hands-off mode of operation. With attitude hold modeengaged, the AFCS maintains aircraft attitude atthe time of engagement in pitch and roll. Thisparticular mode is governed by the actual degreesof bank or pitch of the aircraft. For example, atypical attitude hold mode will release when theaircraft exceedsdegrees in roll.

Altitude Hold

With altitude

±60 degrees in pitch or ±70

hold engaged, the aircraft willmaintain the altitude at the time of engagement.If the aircraft is climbing or diving at engagement,the aircraft returns to the altitude that existed atengagement. The AFCS receives control signalsfor this mode from the air data computer.

Heading Hold

With heading hold engaged, the AFCSmaintains aircraft heading at the time ofengagement. If the pilot is flying a heading of 180degrees and engages the heading hold, the AFCSmaintains the heading of 180 degrees.

Control Stick Steering/ControlWheel Steering

The control stick steering/control wheelsteering mode lets the pilot manually (moving thestick/wheel) change the attitude of the aircraftwith the AFCS engaged without disengaging it.After achieving the new attitude, the pilot releasesthe stick/wheel, and the AFCS resumes controlof the aircraft.

Heading Select

In the heading select mode, the aircraft willautomatically turn to a course selected by thepilot. Upon engagement, the aircraft will assumea fixed maximum roll and turn to the selectedheading.

Mach Hold

The Mach hold function maintains the Machnumber existing at the time of Mach hold

engagement. In this mode, the air data computercommands a pitch-up or pitch-down whenairspeed is above or below the selected mathnumber.

Automatic Carrier Landing System (ACLS)

In the ACLS mode, the pilot can make a“hands-off” carrier landing. The aircraft followscommand signals generated by the data linkreceiver.

Ground Control Bombing

Similar to ACLS, the aircraft follows com-mand signals from personnel on the ground.

Q1. What two types of altitude signals can theAFCS use to maintain a constant altitude?

Q2. Why is the AFCS synchronized with theflight controls before engaging the AFCS?

Q3. What unit is considered the heart of theAFCS?

Q4. What AFCS component provides signaloutputs representing yaw, pitch, and rollrates?

Q5. What type of heading signal does the AFCSreceive from the compass/INS system?

Q6. What are the two modes of operation forelectrohydraulic servo actuators?

Q7. What unit of the AFCS sends a signal thatacts as a follow-up to flight controlmovement?

Q8. What AFCS mode automatically dampensoscillations in the yaw, pitch, and roll axes?

HELICOPTER AFCS

In many respects, the helicopter differsradically from conventional fixed-wing aircraft.However, rotary-wing aerodynamics are verysimilar to fixed-wing aerodynamics. A review ofAirman, NAVEDTRA 14014, will help youunderstand the material in this discussion.

The AFCS is an electrohydromechanicalsystem. It provides inputs to the flight controlsystem to aid the pilot in maneuvering andhandling the helicopter. The AFCS consist ofthree major subsystems—the stability augmenta-tion system (SAS), the stabilator system, and thedigital automatic flight control system (DAFCS).All engagement controls for the three subsystemsare on the AFCS and stabilator control panels.Each subsystem operates independently of theother two subsystems, and they complement oneanother. The pilot engages autopilot functions bypushing the AUTO PLT push button on theAFCS CONTROL panel. The AFCS systemprovides the following features:

. Pitch, roll, and yaw stability augmentation

l Stabilator control

l Cyclic, collective, and pedal trim

l Pitch and roll attitude hold

. Airspeed hold

l Heading hold

. Barometric altitude hold

l Radar altitude hold

l Pitch and roll hover augmentation/gustalleviation

l Turn coordination

. Maneuvering stability

l Automatic approach to hover

. Hover coupler

• Automatic depart

• Crew hover

• Longitudinal stick gradient augmentation(pitch bias actuator)

• Blade-fold assist

• Automatic preflight check

• Diagnostics (mode failure display)

AFCS Control Panels

The pilot controls the AFCS from the AFCSCONTROL panel and the stabilator control panel(figs. 8-30 and 8-31). The stabilator control panelcontains all the operating controls for the stabilator.All the other AFCS controls are on the AFCSCONTROL panel. All detectable AFCS mode failures,except the stabilator, illuminate the AFCS

Figure 8-30.-Helicopter automatic flight controlsystem panel.

Figure 8-31.-Stabilator control panel.

DEGRADED light on the caution/advisory panel. Theywill also illuminate the appropriate mode failurecapsule on the failure advisory section of the AFCSCONTROL panel. Stabilator failures illuminate theSTABILATOR caution light on the caution/advisorypanel and generate an aural warning tone in thepilot’s and ATO’s headsets.

The AFCS CONTROL panel has switches toelectronically engage the SAS 1, SAS 2, and aSAWBOOST HYD switch. The SAS/BOOST HYDswitch controls pressure application to SAS actuatorsand pitch boost servos.

Stability Augmentation System (SAS)

The SAS 1 system provides electrical controlsignals proportional to sensor inputs to the pitch, roll,and yaw servo valves. The servo valves convertelectrical control signals into hydraulic commands forthe SAS actuators. The SAS actuators respond to thehydraulic commands and produce mechanicalmovement of the flight control linkages withoutmoving the cyclic sticks and pedals. The flight controllinkages direct changes in main rotor and tail rotorpitch.

With both SAS 1 and SAS 2 engaged, the SASactuators each have 10 percent authority of flightcontrol movement. Each SAS system has 5 percentauthority. With only SAS 1 engaged, the gain of theSAS amplifier is double, but the control authorityremains at 5 percent. The AFCS control panelprovides a signal to the SAS amplifier to indicate SAS2 engagement.

The SAS 1 pitch channel provides dynamicstability for the helicopter’s pitch axis. The No. 1 pitchrate gyro senses changes in pitch rate and applies thisrate to the No. 1 stabilator amplifier. The No, 1stabilator amplifier filters this pitch rate signal andapplies it to the SAS amplifier. The SAS amplifierprocesses the pitch rate signal to

remove long term rate signals and applies acorrection signal to the pitch SAS servo valve. Theservo valve controls hydraulic pressure to the SASactuator. The SAS actuator provides mechanicalmovement of the flight controls, producingrotor head movement opposing the sensed pitchrate.

The SAS 1 roll channel provides dynamicstability and limited roll stability for thehelicopter’s roll axis. A roll attitude signal isprovided to the SAS amplifier from the ATO’sAttitude. Heading Reference system. The SASamplifier limits this signal and uses it with a rollrate signal from an internal roll rate gyro. Theamplifier applies the resultant correction signalto the roll SAS servo valve. This servo valvecontrols the actuator, which positions the flightcontrols to oppose the sensor inputs.

The SAS 1 yaw channel provides dynamicstability for the helicopter’s yaw axis and turncoordination. At airspeeds of less than 50 knots,the No. 1 stabilator amplifier provides a +12 to+15 VDC airspeed switch discrete. The discreteinhibits No. 1 lateral acceleration and roll ratesignals in the SAS amplifier. The discrete controlsthe internal yaw rate gyro in the SAS amplifieras the yaw channel sensor. When sensing a yawrate, the SAS amplifier processes it to remove longterm yaw rate signals. It then develops a shortterm correction signal and applies it to the yawSAS servo valve. The servo valve controls theactuator, which moves the flight controls tooppose the yaw rate.

At speeds below 50 knots, the airspeedtransducer is applied to stabilator amplifier No.1 and sent to SAS 1 amplifier (discrete signal).This discrete signal allows the yaw SAS amplifierand flight controls to use the internal yaw rategyro. Above 50 knots, the SAS amplifier recievesa –12 to –15 VDC airspeed switch discrete fromthe No. 1 stabilator amplifier. This enables filteredNo. 1 lateral acceleration and roll rate to sum withthe yaw rate signal. Airspeed discrete from No.1 stabilator amplifier is inhibited for this situation(above 50 knots). Airspeed discrete is not removedfrom SAS amplifier.

The AFCS consists of many components. Tounderstand the entire system, you need to knowwhat each component does. The followingdiscussion describes the components that makeup the AFCS.

AFCS POWER SWITCHING ASSEMBLY. —The AFCS power switching assembly providespower switching for the digital automatic flight

control system computer and the SAS andBOOST shutoff valves.

NO. 1 LATERAL ACCELEROMETER. —The accelerometer provides lateral acceleration tothe SAS amplifier via the No. 1 stabilatoramplifier for the turn coordination function ofSAS 1.

NO. 1 PITCH RATE GYRO. —The gyroprovides pitch rate to the SAS amplifier via theNo. 1 stabilator amplifier for the pitch dynamicstability function. The pitch rate gyros are alsopart of the stabilator control system.

SAS ACTUATORS. —The SAS actuators, onthe pilot-assist servos, convert electrical signalsfrom the analog SAS 1 and digital SAS 2 tomechanical motion to move the flight controls.The actuators are electrohydraulic, operatingfrom pressure supplied by the No. 2 transfermodule, applied through the SAS shutoff valve.When the SAS/BOOST HYD switch is pressed,power is removed from the SAS shutoff valve,opening the valve. Opening the valve allows theactuators to move with signals applied from SAS1, SAS 2, or both. With pressure removed fromthe actuators, a spring-loaded device locks theactuator piston in center position. Each actuatoroutput piston connects to linkage. When theactuator piston becomes locked, it forms a fixedpivot point on one end of the linkage. This allowsthe linkage to move with stick, pedal, or triminputs.

SAS AMPLIFIER. —The SAS amplifier con-tains the No. 1 roll and yaw rate gyro powersupplies and processing circuitry for SAS 1. Theoutputs of the internal roll and yaw rate gyros areused by SAS 1 and the digital automatic flightcontrol system computer. The amplifier processesrate and proportional signals. Amplifier outputsoperate the SAS actuators on the pilot-assist servoassembly.

NO. 1 STABILATOR AMPLIFIER. —TheNo. 1 stabilator amplifier provides filtered lateralacceleration and pitch rate signals and an airspeeddiscrete to the SAS amplifier. The stabilatoramplifier is part of the stabilator control system.

Stabilator System

The stabilator system optimizes trim attitudesfor cruise, climb, and autorotation. Also, it

provides pitch stability augmentation to comple-ment the SAS system for additional redundancy.The stabilator system is completely independentof the other two AFCS subsystems except forcommon airspeed sensors, lateral accelerometers,and pitch rate gyros. The stabilator control systemis a completely automatic fly-by-wire controlsystem with a manual backup slew control. Theprimary purpose of the stabilator control systemis to stop undesirable noseup attitudes. Thenoseup attitude is caused by rotor downwashimpinging on the horizontal stabilator duringlow-speed flight and transition to a hover.

The stabilator panel contains an automaticcontrol (AUTO CONTROL) switch, a TEST pushbutton, and a manual slew (MAN SLEW) switch.The AUTO CONTROL switch is used to engagethe automatic mode or to reset the stabilator ifit should fail. The pilot can manually position thestabilator to any position within the stabilatorlimits by moving the MAN SLEW switch. TheTEST push button is used to check the automaticmode fault detector.

Two electric jackscrews, working in series,position the stabilator. Each actuator providesone-half the input to position the stabilator andis controlled by a separate and redundantstabilator amplifier. The stabilator travels from40 degrees trailing edge down for hover andlow-speed flight below 30 knots to 8 degreestrailing edge up for cruise and maneuveringflight. Four inputs are required to position thestabilator—airspeed, collective stick position,lateral acceleration, and pitch rate.

Each stabilator amplifier receives thesefour inputs, but they receive the inputs fromindependent sensors. The DAFCS computermonitors each of these sensors for malfunctions,and the stabilator control system monitors andcompares the position of the two actuators. Anysystem malfunction caused by a differencebetween the two stabilator actuator positionsresults in the stabilator remaining in the lastposition. A malfunction also causes an automaticpower shutdown to both actuators, an aural toneto the pilot, and a STABILATOR caution lightilluminating. The shutdown threshold between thetwo stabilator actuator positions is 10 degrees forairspeeds less than 50 KIAS and 4 degrees forairspeeds greater than 120 KIAS with a linearvariation between 50 KIAS and 120 KIAS.

The airspeed input aligns the stabilator withthe main rotor downwash during slow-speedflight. The collective stick position input decouplesaircraft pitch attitude with collective position.

Pitch rate and lateral acceleration inputs improvethe dynamic response of the aircraft, especiallyin gusty air conditions. The pitch rate inputsupplements the dynamic stability provided by theSAS and DAFCS. The lateral accelerometer inputdecouples the aircraft pitch response with changesin tail rotor lift caused by changes in airflow onthe canted tail induced with sideslip.

If a malfunction of the stabilator systemoccurs, the pilot can manually position thestabilator with the manual slew switch. Themanual slew switch bypasses the stabilatoramplifier automatic mode, applying powerdirectly to the actuators through relays in theamplifiers. A stabilator position indicator aids thepilot in positioning the stabilator to any positionbetween the stabilator travel limits. However,total travel is restricted if the malfunction is anactuator failure. Stabilator travel is restricted to35 degrees if an actuator fails in the full-downposition; 30 degrees if an actuator fails in thefull-up position. The stabilator control rate islimited to ±6 degrees per second.

The following is a description of the com-ponents of the stabilator system. To understandthe system, you will need to know what thesecomponents do.

AIRSPEED TRANSDUCER. —The stabi-lator position program is a function of four sensorinputs. Each function has dual sensors forfail-safe operation. Airspeed for the No. 1stabilator system is sensed by the airspeedtransducer. It connects into the ATO pitot staticsystem and produces a dc output voltage that isproportional to airspeed.

AIR DATA TRANSDUCER. —An air datatransducer accomplishes airspeed sensing in theNo. 2 stabilator system. The air data transducerconnects into the pilot’s pitot static system. Itproduces dc output voltages proportional toairspeed, altitude, and altitude rate. Airspeed datais used by the stabilator system and AFCScomputer. Altitude and altitude rate are appliedto the DAFCS computer only for altitude holdand depart modes of operation.

LATERAL ACCELEROMETERS. —Thereare two lateral accelerometers—No. 1 and No 2.They each produce a dc output signal proportionalto helicopter lateral acceleration. Each accel-erometer dc signal goes to its related stabilatoramplifier, where it is conditioned. The filteredlateral acceleration signal is used to position the

stabilator to counteract tail rotor downwash onits upper wing surface. No. 1 and No. 2 filteredlateral acceleration is used by the DAFCScomputer for signal quality comparison andsoftware generation. No. 1 filtered lateralacceleration is also used in analog SAS for slipor skid correction above 50 knots.

STABILATOR POSITION INDICATOR. —Stabilator position is displayed by an indicator onthe center of the cockpit instrument panel. Thisindicator is a synchro-type device driven by asynchro transmitter mounted in the stabilatorposition transmitter and limit switch assembly inthe tail pylon.

STABILATOR CONTROL PANEL. —Con-trol functions for the system are provided by thestabilator control panel. The panel consistsof an AUTO CONTROL PUSH TO RESETpush-button switch, a TEST push button, and aMAN SLEW UP/DOWN switch. Engagement ofthe system is automatic, upon application ofhelicopter ac and dc power, provided all interlocksare in their proper condition.

TEST Push Button. —A TEST push button,operational below 50 knots, provides a check ofthe system fault monitors by inserting an airspeedderived test signal into the No. 1 system. Thissignal drives only the No. 1 stabilator actuator,which produces a difference between the twoactuators. The fault monitor circuit in either theNo. 1 or No. 2 amplifier, or both, shoulddisengage the automatic mode of operation whenthe programmed threadhold trips.

MAN SLEW UP/DOWN Switch. —If theautomatic mode disengages, and cannot be resetdue to a malfunction, the MAN SLEW switch isused to manually position the stabilator. Relaysin the No. 1 and No. 2 stabilator amplifiers areoperated by dc power from the switch, when itis placed to UP or DOWN. Using the switch,when the automatic mode is engaged, willdisengage the automatic mode. As a result, theSTABILATOR caution light will go on, and abeeping tone will be heard in the ICS.

COLLECTIVE STICK POSITION SENSORS. –Collective stick position also affects the stabilatorposition schedule. Stick position is sensed by twocollective stick position sensors. No. 1 and No.2 collective stick position sensors each producesa dc output signal proportional to the collective

stick position. Both signals are used by theDAFCS computer for signal level comparison andfor software generation of collective-to-yaw pedalcoupling.

STABILATOR AMPLIFIER. —Processing ofairspeed, lateral acceleration, collective stickposition, and pitch rate is accomplished withineach stabilator amplifier. The amplifiers containa power supply, processing and feedback circuits,and a fault monitor circuit. Sensor inputs areprocessed, summed, and applied to a motor drivercircuit. The motor driver circuit output is applied,through contacts of relays, to the respectivestabilator actuators. Any difference of actuatorposition is sensed by the fault monitor circuit ineither or both amplifiers, causing an automaticmode disengagement.

ACTUATORS. —Two actuators position thestabilator. Each actuator contains an electricmotor (geared to a jackscrew), limit switches, anda feedback potentiometer. The potentiometerprovides actuator position feedback to eachamplifier. The actuators extend or retract, asnecessary, to position the stabilator.

PITCH RATE GYROS. —There are two pitchrate gyros—No. 1 and No. 2. Each rate gyroproduces a dc output signal relative to the pitchrate of the aircraft. Each rate gyro signal goes toits respective stabilator amplifier, where it isconditioned. The stabilator system uses thefiltered pitch rate signal to enhance the AFCSsystems ability to correct short-term pitchdisturbances. The DAFCS computer uses No. 1and No. 2 filtered pitch rate signals for signalquality comparison and software generation. TheNo. 1 filtered pitch rate signal is also used inanalog SAS for short-term pitch correction of therotor head.

Digital Automatic Flight Control System(DAFCS)

The central component of the DAFCS is thedigital computer. The computer commands thepitch bias actuator (PBA), the inner-loop SASactuators, and the outer-loop trim actuators in allfour control channels. The computer also providesself-monitoring, fault isolation, and failureadvisory.

The DAFCS uses two types of control–identified as inner loop and outer loop. Theinner loop (SAS) uses rate damping to improve

helicopter stability. This system is fast in response,limited in authority, and operates without causingmovement of the flight controls. The outer loop(AUTO PILOT) provides long-term inputs bytrimming the flight controls to the positionrequired to maintain the selected flight regime.It can drive the flight controls throughout theirfull range of travel (100-percent authority). Theouter-loop drive rate is limited to 10 percent persecond. Both inner and outer loops allow forcomplete pilot override through the normal useof the flight controls.

The DAFCS computer processes incominginformation from various sensors (fig. 8-32)aboard the aircraft and stores this information inits memory. The central processing unit (CPU)uses the sensor information to compute requiredcorrection signals. Inner-loop correction signalsgo to the SAS actuators, and outer-loop signalsoperate trim servos and actuators.

TRIM SYSTEM. —The parallel trim actuatorassemblies provide the flight control forcegradients and detent positions and the outer-loopautopilot control functions. The trim actuatorscommand full control authority in all four controlchannels, but are rate-limited to 10 percent persecond. Pressing the trim release switch (cyclictrim release, collective trim release, and pedalrelease) disengages the respective trim functionand allows free control motion. Releasing the trimrelease switch re-engages trim. For yaw trimrelease above 50 knots, the pilot must press thepedal microswitches and the cyclic trim switch.Below 50 knots only the pedal microswitches haveto be pressed. The pilot can override the trimcontrol forces in all channels.

AUTOPILOT. —The autopilot maintainshelicopter pitch and roll attitude, airspeed, andheading during cruise flight and provides acoordinated turn feature at airspeeds above 50knots. To engage the autopilot function, the pilotpresses the control panel SAS 1 or SAS 2 switches,TRIM switch, and then the AUTO PLT pushbutton. The autopilot may be disengaged bypressing the AUTO PLT push button or pressingthe AFCS release button. The computer alsoprovides command signals to the trim actuatorsto reposition the flight controls using the trimsystem.

ATTITUDE AND AIRSPEED HOLD. —Attitude and airspeed hold are engaged withAUTO PLT. In the pitch channel, at airspeeds

of less than 50 knots, attitude changes arecommanded by changing the cyclic stick position.The pilot can use the TRIM REL switch or thefour-direction (beeper) TRIM switch to changecyclic stick position. This causes the cyclic stickto move, and the helicopter attitude to changeabout 5 degrees per second. When cyclic move-ment stops, the autopilot stabilizes the helicopteraround the new stick position and attitude. Above50 knots and bank angles less than 30 degrees,the system becomes airspeed sensitive in pitch.Operating the four-direction TRIM switch causesthe cyclic stick to move and the helicopter tochange airspeed reference at 6 knots per second.Because of variations in the pitot-static systemduring gusty conditions, integrated longitudinalacceleration is used for short-term correction. Theairspeed sensor is used for long-term updatesthrough a 3-second filter.

The roll channel autopilot holds roll attitudeof the helicopter. Attitude information is suppliedto the computer from the pilot’s and copilot’sA/A24G vertical gyros. The command signal isapplied to roll SAS 1 and SAS 2 and the roll trimsystem. When the pilot actuates the four-directionTRIM switch, the helicopter roll attitude willchange at about 6 degrees per second. In additionto the attitude hold feature, the system includesan automatic wing-leveling capability. Duringtransitions from hover to airspeeds above 50knots, this feature automatically retrims theaircraft from a left roll attitude in a hover to awings level attitude at 50 knots. After establishinga level attitude, the attitude hold feature maintainsthat attitude until a new roll attitude iscommanded by the pilot.

HEADING HOLD. —The yaw channel of theautopilot provides the heading hold feature forhover and forward flight. It is engaged wheneverthe AUTO PLT PBS is illuminated. Heading holdis an outer-loop function, operating through theyaw trim actuator; therefore, it will only workwhen the yaw trim is engaged. Releasing all pedalswitches at a given heading synchronizes thetrim system to the established heading. Apotentiometer in the yaw trim actuator applies atrim position feedback signal to the computer.This signal cancels the drive signal at the desiredposition, stopping the motor. The yaw autopilotalso uses a collective stick position sensor to holdreference heading for yaw excursions caused bymain rotor torque changes. The collective stickposition sensor is controlled by an airspeed signal

Figure 8-32.-DAFCS input/output block diagram.

that reduces its gain as airspeed increases. When Above 50 KIAS, actuation of the switch for lessthe heading hold is engaged, the HDG TRIM than 1 second provides a 1-degree heading change.(slew) switch on the collective lets the pilot make Actuation for greater than one second providesheading changes without retrimming. Below 50 a 1 degree per second coordinated turn. TheKIAS, the aircraft slews at 3 degrees per second. heading hold is re-engaged following a turn when

the following conditions are maintained for 2seconds:

• Aircraft roll attitude is within 2 degrees ofwings level.

• Yaw rate issecond.

The heading hold

less than 2 degrees per

is disengaged by the weight-on-wheels switch when the aircraft is on theground.

ALTITUDE HOLD. —Either barometric orradar altitude hold is selectable from the AFCSCONTROL panel. With the altitude hold modeon, the DAFCS computer uses a referencealtitude. The reference altitude comes from eitherthe air data transducer or the radar altimeter,depending on whether barometric altitude holdor radar altitude hold is selected. The DAFCScomputer uses altitude and rate from thebarometric or radar altitude systems (dependingon which hold mode is selected) and verticalacceleration to command the collective SAS andtrim actuators. The computer also monitorsengine torque to prevent dual engine torque fromexceeding 116 percent whenever the collective trimis positioning the collective. Barometric altitudehold is engaged at any altitude and airspeed bydepressing the BAR ALT PBS with SAS 2 andautopilot engaged. Depressing the collectivetrim release button temporarily disengagesthe mode. Upon release of the trim switch,barometer altitude hold automatically re-engagesand maintains the altitude at the time ofre-engagement. Radar altitude hold is engaged atany altitude from 0 to 5,000 feet AGL and at anyairspeed by depressing the RDR ALT PBS withSAS 2 and autopilot engaged.

When in the hover coupler mode, altitude holdis referenced to the altitude selected on the AFCSCONTROL panel HVR ALT potentiometer.Depressing collective trim release temporarilydisengages the mode. Upon release of thetrim switch, radar altitude hold automaticallyre-engages to the altitude selected on the AFCSCONTROL panel HVR ALT potentiometer.When in the hover coupler mode, transitionfrom one altitude to another is made withthe HVR ALT knob on the AFCS CONTROLpanel. If the radar altitude mode should fail

while engaged, the barometric altitude holdautomatically engages. Integrated verticalacceleration provides short-term radar altitudecorrections, and rate information from the radaraltimeter altitude signal provides long-termupdates.

HOVER AUGMENTATION/GUST ALLE-VIATION. —An additional feature of theSAS, provided only through SAS 2, is hoveraugmentation/gust alleviation. It furtherimproves aircraft stability at low airspeed usingattitude retention and longitudinal and lateralacceleration to eliminate drift.

TURN COORDINATION. —Automatic turncoordination is provided at airspeeds greater than50 knots. Turn coordination lets the pilot fly acoordinated turn with directional control providedby the AFCS. The AFCS uses lateral accelerationand roll rate to determine if the aircraft is out ofbalanced flight. It also provides the yaw SAS andyaw trim with the inputs necessary to maintainan automatic coordinated turn. Automatic turncoordination is engaged and heading holddisengaged when roll attitude is greater than 1degree, and any of the following conditionsexist:

• Lateralpercent

cyclic force is greaterstick displacement.

than 3.0

• Cyclic trim release is pressed.

• Roll attitude is beeped beyond 2.5 degreesof bank angle.

Actuation of the collective-mounted headingslew for greater than 1 second provides a 1 percentper second coordinated trim.

MANEUVERING STABILITY. —Pitchcontrol forces are increased to increase pilot effortrequired for a given pitch rate at bank anglesgreater than 30 degrees. The higher pitch controlforces help alert the pilot to g loading duringmaneuvering flight, and are provided through thelongitudinal trim actuator. A linear longitudinalstick force gradient is provided by trimming 1percent forward stick for each 1.5 degrees angleof bank between 30 degrees and 75 degrees. At75 degrees angle of bank, the longitudinal stick

force is equivalent to 30 percent of stickdisplacement. The maneuvering stability featureis engaged whenever the AUTO PLT PBS isilluminated.

AUTOMATIC APPROACH TO HOVER. —The DAFCS provides the capability to performan automatic approach to a zero longitudinal andany lateral ground speed selected on the LAT VELcontrol knob. Also, the DAFCS can perform anautomatic approach to any radar altitude selectedon the HVR ALT control knob between 40 feetand 200 feet. If the HVR ALT is set below 40 feet,the approach will be made to 40 feet and thencontinued to the HVR ALT setting when themode is switched from APPR to HVR. Theautomatic approach can start at any airspeed andaltitude. The automatic approach is an outer looponly function, and it commands the aircraft todecelerate or descend until meeting the approachprofile conditions. If the approach mode isselected when aircraft conditions are below theapproach profile, the DAFCS commands theaircraft to decelerate. The aircraft will decelerateat 1 knot) second while in the radar altitude holdmode until the approach conditions are met. Ifthe approach mode is selected when the aircraftis above the approach profile, the DAFCScommands the aircraft to descend. The descentoccurs at 360 feet/minute when the aircraft ismore than 50 feet above the approach profile. Itoccurs at 120 feet/minute when the aircraft is lessthan 50 feet above the profile. During theseconditions, the DAFCS uses the radar altimeteruntil the approach profile conditions are met.When the approach profile conditions are met,the aircraft simultaneously decelerates at 1knot/second and descends at 120 feet/minute.This profile is maintained until the aircraft attains1 knot of Doppler ground speed and comes towithin 1 foot of the selected radar altitude. If theselected altitude is below 40 feet, the aircraft fliesto 40 feet and zero longitudinal ground speed andthen descends to the selected attitude. Whenground speed equals 1 knot or less and the aircraftaltitude is within 2 feet of the selected altitude,the hover coupler mode automatically engages.The aircraft then accelerates to the selectedlongitudinal ground speed.

HOVER COUPLER. —The hover couplerprovides longitudinal and lateral ground speedcontrol and stabilization about the selected groundspeed and automatic altitude retention. Thelongitudinal and lateral ground speed and the

altitude are selectable on the AFCS CONTROLpanel. Longitudinal and lateral ground speed canalso be beeped ± 10 knots with the cyclic trimswitch about the ground speed selected on theAFCS CONTROL panel, The hover couplermode can automatically engage at the terminationof the automatic approach. Also, the pilot canmanually engage it when the aircraft is hoveringwith less than 5 knots longitudinal groundspeed.To do this, the pilot presses the APPR/HVRbutton on the AFCS CONTROL panel with SAS2, TRIM, and AUTO PLT engaged. The hovercoupler engages when the longitudinal groundspeed is less than 5 knots if engaged manually.After engagement, the aircraft accelerates to thelongitudinal and lateral ground speeds selected onthe AFCS CONTROL panel. Pressing andreleasing the cyclic TRIM REL removes cyclictrim switch inputs. This action returns the aircraftto the LONG VEL and LAT VEL settings on theAFCS CONTROL panel. Because of Dopplernoise, short-term longitudinal and lateral groundspeed is obtained from integrated longitudinal andlateral inertial acceleration. Long-term correctionis obtained from the Doppler sensor using a7-second filter.

AUTOMATIC DEPART. —The automaticdepart mode provides the capability to performan automatic departure from a coupled hover orfrom an automatic approach. This mode can takethe aircraft to a cruise airspeed of 100 KIAS andaltitude of 500 feet. If the coupled hover or theautomatic approach feature is engaged, engagethe automatic depart mode by depressing thehover depart button on the cyclic grip. Depressingthe DEPART PBS a second time disengages theautomatic depart mode, but it doesn’t auto-matically re-engage the RAD ALT hold. Depress-ing the hover depart button a second timedisengages the automatic depart mode and returnsaircraft control to the pilot. Upon engagement,the aircraft accelerates at 2 knot/second andclimbs at 480 feet/minute. During the departure,the DAFCS computer monitors engine torque toensure that it does not exceed 116 percent. At 100KIAS, the airspeed hold automatically engages;at a radar altitude of 500 feet, the radar altitudehold automatically engages. Any alternate cruiseairspeed or altitude condition of less than 100KIAS and 500 feet is available to the pilot. Thepilot attains an alternate condition by depressingthe cyclic trim release and collective trim releaseat the desired airspeed and altitude, respectively.If either trim release button is depressed and

released, the hold mode (airspeed or altitude)associated with that control axis is engaged. Theaircraft continues to follow the depart profile forthe other axis until the final cruise condition forthat axis is met.

The automatic depart mode is an outer-loopfunction operating through the pitch, roll, andcollective trim actuators. As in the automaticapproach mode, above 60 KIAS, roll attitude ismaintained, and below 60 KIAS, the DAFCScommands roll to eliminate lateral drift.

CREW HOVER. —The crew hover feature letsthe crewman position the helicopter during hoistand rescue operations. The crewman controls theaircraft from the crew hover-trim panel. The crewhover controller has a control authority of ± 5knots. This authority is laterally and longi-tudinally about the reference values selected onthe AFCS CONTROL panel plus the speedsbeeped from the cyclic trim beep switch, The crewhover feature is activated from the AFCSCONTROL panel by depressing the CREW HVRbutton. It can only be activated if the hovercoupler mode is already engaged.

PITCH BIAS ACTUATOR (PBA). —ThePBA provides longitudinal cyclic displacementproportional to airspeed. The DAFCS commandsthe PBA as a function of pitch attitude, pitch rate,and airspeed. The PBA is an electromechanicalseries actuator with ±15 percent control authorityand ±3 percent per second rate limit. The PBAfunctions automatically upon application ofpower to the DAFCS computer, and it isn’tselectable on the AFCS control panel. TheDAFCS computer monitors the PBA position toconfirm correct response to the input commands.If the PBA fails, the DAFCS lights the BIASadvisory light and flashes the AFCS DEGRADEDlight. Also, the DAFCS commands the PBA toa predetermined position, depending on the typeof failure. The PBA is driven by the DAFCScomputer as a function of airspeed, pitch rate,and pitch attitude. PBA failure modes are asfollows:

Attitude failure—bias actuator centered

Pitch rate failure—faded out pitch ratecomponent

Airspeed failure—actuator goes to120-knot position and attitude and ratecontinues to function.

• Actuator failure—power removed fromactuator.

If the system malfunctions, the BIAS FailAdvisory light on the AFCS CONTROL panelwill go on, and, in some cases, remove powerfrom the actuator. If the malfunction that causedthe shutdown was of an intermittent nature, theactuator operation can be reset by pressing theappropriate MODE RESET button.

AUTOMATIC STABILIZATIONEQUIPMENT

Learning Objective: Recognize the operat-ing principles and triodes of the automaticstabilization and coupler system, andidentify control panel and channel monitorpanel functions.

Personnel in the AE rating maintain theautomatic stabilization equipment (ASE) dis-cussed in this chapter. The ASE is used in SH-3aircraft, and is the system used to illustrate ASEequipment. For detailed information on ASE, youshould refer to the appropriate maintenanceinstruction manual (MIM).

The ASE is to helicopters as the AFCS is tofixed-wing aircraft. ASE is designed to improvethe handling characteristics of the helicopter. ASElets the pilot fly the helicopter in automaticcruising flight and in hands-off flight. Whenenergized, the ASE activates the attitude stabiliza-tion and heading retention systems in thehelicopter. Also, ASE places the barometricaltitude control in a synchronizing condition soit can be engaged when desired.

ASE provides selective attitude retention in allflight axes, and it also improves handlingcharacteristics by introducing stability corrections.Barometric altitude (BAR ALT) maintains theaircraft at the desired altitude manually selectedby the pilot. The coupler (CPLR) enables thehelicopter to find and keep pilot-selected speeds,drifts, and altitudes during automatic cruise flight.During antisubmarine warfare (ASW) missions,the ASE enables the helicopter to make atransition from forward flight to hover. It alsoaids in maintaining a hover during sonar search.An attitude indicating system provides a visualdisplay of the helicopter’s attitude on the pilot’sand copilot’s attitude indicators.

ASE, with the coupler system, uses associatedelectronic and hydraulic systems to provideautomatic hovering and automatic cruising. Thecoupler system consists of a Doppler radar system(ground speed measurement) and a verticalvelocity system (used with a low-altitude radaraltimeter). Since radar systems are not maintainedby the AE, they are not discussed in detail. Radarsystems are mentioned only because they serve assignal sources for ASE operation.

ASE COMPONENTS

ASE, like the AFCS, consists of manycomponents, including electrical/electronic,

hydraulic, and mechanical. You will be able tounderstand how the system functions if you candescribe the components and understand how theywork.

Electrical/Electronic Components

The following paragraphs describe the majorelectrical/electronic components that make upan ASE system. These electrical/electroniccomponents make the ASE function.

ASE CONTROL PANEL. —The ASE controlpanel (fig. 8-33) is on the cockpit center consolebetween the pilot and copilot. The panel containsfive push-button switches and five control knobs.These switches and knobs control the ASE modesof operation. The control also contains two toggleswitches used with the other controls. When thepilot presses the ASE button, the pitch, roll, andyaw control circuits start operating.

When the pilot presses the BAR ALT button,the barometric altitude controller in the collectivechannel engages. The collective channelbarometric altitude controller disengages when thepilot depresses the BAR OFF button. If thebarometric altitude controller is engaged, the pilotcan momentarily release it by pressing the BAR

Figure 8-33.-ASE control panel.

REL button on the collective pitch stick grip. Thisaction makes possible a change in altitude.

By pressing the CPLR button, the pilotengages the coupler and the ASE collectivechannel. The ASE must be on before engagingthe coupler. The push button marked HOVERTRIM ENG can only be engaged when the CYCCPLR switch is in DOPP position. WhenHOVER TRIM is engaged, hover trim controltransfers from the pilot to the crew member atthe hoist position. Hover trim allows the hoistoperator to maneuver the helicopter fore and aftand port and starboard during hoisting or similaroperations. Hover trim control authority is limitedto 11 ±4.5 knots of the speed established by theASE speed control. The coupler release button,marked CPLR REL, on each collective pitch leverdisengages the coupler portion of ASE. It alsodisengages the hover trim control. If the operatordesires to remain in a Doppler hover, the hovertrim may also be disengaged by cycling the cycliccoupler switch to OFF and then to DOPP.

Control knobs on the ASE panel havedefinitive characteristic shapes that permit thepilot to identify them without looking at them.Specific control knobs and their identifying shapesare described in the following paragraphs.

Altitude Set Knob. —This knob is crossshaped. With basic ASE and the radar altitude(RDR ALT) coupler engaged, the helicopter canaccurately maintain any selected altitude from0 to 200 feet. This mode of operation isadvantageous during hover operations, as the pilotcan vary altitude by rotating a single control.

Various models of ASE have different scaleson the attitude set knob at different increments.(In one model, the scale runs from 0 to 1,000 feetwith varying increments. In another model, thescale runs from 0 to 500 feet.) In the ASEdiscussed in this chapter, the ALTITUDE controlscale is from 0 to 200 feet in 5-foot increments.The 0 to 200 foot configuration offers accuracythroughout the full range of travel of the controlknob.

Yaw Trim Knob. —This knob is triangularshaped. The pilot turns the yaw control knob totrim the heading of the helicopter during flight.One rotation of the knob will turn the helicopter72 degrees. Each increment on the knob equals1 degree of movement.

Ground Speed Set Knob. —This knob is anindented circle. The pilot turns the ground speed

set knob to preselect the forward or aft groundspeed to be maintained by the coupler. Frictionreduction built into the speed control permits thepilot to make small changes in speed.

Drift Knob. —This knob is bar shaped. Thepilot turns the drift set knob to preselect thecompensation for the helicopter’s lateral driftabout its ground track. The roll coupler mode willmaintain this compensation. Full-scale sensitivity yis low enough to allow the pilot to make smallchanges in drift.

CC Trim Knob. —This knob is clover shaped.The pilot turns the CG (center of gravity) trimknob to adjust the pitch in small increments bothfore and aft. The CG trim signal is summed withthat of the position sensor to maintain apreselected pitch.

Cyclic Coupler Switch. —The cyclic couplerswitch selects the coupler mode of operation. Theswitch has three-positions with the center positionbeing OFF. When the switch is in the DOPPposition, information comes from the Dopplerradar. It also activates the Doppler mode of pitchand roll coupler operation.

With selection of the CABLE ANGLEposition, the sonar operator controls the pitch androll trim angle using the cable angle control panel.This control allows for compensation of helicopterdrift over the sonar transducer.

Altitude Coupler Switch. —The altitudecoupler switch is used with the coupler button foraltitude reference. It is also a three-positionswitch. The RDR ALT position engages the radaraltitude mode for absolute altitude reference. TheCABLE ALT position engages the transducercable altitude for absolute altitude reference.When the CPLR button is depressed with eitherRDR ALT or CABLE ALT selected, the BARALT mode automatically engages. The VERTACCEL position of the altitude coupler switchengages the vertical accelerometer mode toeliminate collective pumping.

CHANNEL MONITOR PANEL. —Thechannel monitor panel (fig. 8-34) is to the rightof the pilot’s seat. This panel contains fourON-OFF toggle switches across the top row andfour guarded switches marked HARDOVERacross the middle of the panel. It also includesa meter select switch, and a vertical gyro switch.

Figure 8-34.-Channel monitor panel.

The four ON-OFF toggle switches are PITCH,ROLL, COLL, and YAW. These switches let thepilot disengage any of the four channels of theASE. Normally, the switches are left in the ONposition. They are positioned to OFF to disengagea malfunctioning channel. The four guardedswitches are for maintenance personnel tointroduce HARDOVER signals to the individualchannels when they test or troubleshoot the ASE.

NOTE: The channel monitor test switch onthe pilot’s compartment dome light panel MUSTbe in the test position before using the HARD-OVER switches.

The METER SELECTOR switch lets eitherASE or CPLR inputs be monitored with the hoverindicator (covered later in the chapter) in the Amode.

The VERTICAL GYRO SELECTOR switchis marked PORT and STBD. This switch lets thepilot select the port or starboard vertical gyro forASE pitch and roll reference sources.

CABLE ANGLE CONTROL PANEL. —Thecable angle control panel (fig. 8-35) is located atthe sonar operator’s console. This panel contains

Figure 8-35.-Cable angle control panel.

an engagement switch (push-button type) andtwo potentiometers labeled ROLL DRIFT andPITCH DRIFT. The sonar operator has controlof the ASE when the CPLR mode is on, and thecyclic coupler switch is in the CABLE ANGLEposition. The sonar operator may engage andcontrol the transducer cable angle by trimmingthe helicopter’s pitch and roll attitude. Thislets the operator maintain the cable angleperpendicular, relative to the horizon.

HOVER TRIM CONTROL PANEL. —Thispanel (fig. 8-36) features two rotary knobs and

Figure 8-36.-Hover trim control panel.

a spring-centered hover trim stick. One of the tworotary knobs is marked ROLL BIAS, withdirectional arrows for left and right. The otherrotary knob is marked PITCH BIAS, withmarked arrows for forward and aft. The hovertrim switch is outlined in its corners with the lettersR (for right), L (for left), F (for forward), andA (for aft).

The hoist operator uses the knobs on the hovertrim panel to control helicopter drift. With thehover stick centered, the hoist operator adjuststhe roll and pitch bias knobs to trim the aircraftfor hover. Signals introduced into the amplifierby the hoist operator correct any slight driftingof the helicopter while the aircraft is hovering.To energize the hover trim circuits, the pilot mustdepress the HOVER TRIM ENG button on theASE control panel with the cyclic coupler switchin the DOPP position.

DUAL-CHANNEL LAG AMPLIFIER. —The function of the dual-channel lag amplifier(fig. 8-37) is to delay (lag) applied signals to ensureproper response in helicopter movement. Thevertical velocity, roll stick position sensor, and thepitch and roll cable angle sensor signals are laggedby this amplifier.

ASE AMPLIFIER. —The ASE amplifier (fig.8-38) compares pilot-selected attitude, altitude,and heading with the helicopter’s actual attitude,altitude, and heading. The amplifier also provideserror signals that maintain helicopter stability,within limited authority, regardless of combinedaerodynamic disturbances.

The ASE amplifier integrates the input signalsfrom the sensors and directs control operation.It also provides an output to the hydraulic sectionsof the control system activating servo units. The

Figure 8-37.-Dual-channel LAG amplifier.

Figure 8-38.-ASE amplifier.

ASE amplifier contains eleven amplifiers. Theamplifiers include four basic ASE modules, twocoupler integrator modules, two coupler amplifiermodules, one yaw synchronizer module, onecollective coupler module, and one verticalaccelerometer amplifier module. The ASEamplifier also contains nine gain capsules and ayaw rate gyro. The ASE amplifier modules andcapsules are discussed below.

Modules. —The four basic ASE modules arepitch, roll, collective, and yaw. They serve toamplify, shape, and derive the rate signalsused during ASE operation. These modules areinterchangeable.

The pitch and roll coupler amplifier modulesare used during pitch and roll coupler operation.These modules are also interchangeable.

The pitch and roll coupler integrator modulesare also interchangeable. They provide steady-state error signals during pitch and roll coupleroperation.

The collective coupler module amplifies signalsduring collective coupler operation.

The vertical accelerometer amplifier operateswith selection of the CABLE ALT mode ofcollective coupler operation. The vertical accel-erometer amplifier filters, amplifies, and lags thevertical accelerometer signal in the CABLE ALTmode.

The yaw synchronizer modules keep the yawchannel input nulled, except when ASE isoperating and the pilot’s feet are not on thecontrol pedals.

Each ASE amplifier module and each of theyaw synchronizer modules contain a regulatedpower supply. Individual ac and dc fuses preventfailure in one channel from affecting anotherchannel.

Gain Capsules. —Each module, except the yawsynchronizer module and the vertical accelerom-eter amplifier module, contains a plug-in gaincapsule. A gain capsule consists of resistors andjumpers that determine module gain and signalpath. Since each module has a particular function,you must be sure the proper gain capsule isinserted in the amplifier.

ATTITUDE DEMODULATOR. —When theASE is operating in the C mode, the pitch androll vertical gyro signals are demodulated by theattitude amplifier demodulator (fig. 8-39). The dcoutput of the demodulator is compared with thecable angle drift sensor dc voltage. The sum ofthe two dc voltages goes to the hover indicators.

Figure 8-40.-Hover indicator.

indicator, which moves horizontally leftcoinciding with the horizontal scale.

There is a mode selector switch on

HOVER INDICATOR. —There are two hoverindicators on the instrument panel (fig. 8-40). Theindicators show Doppler information duringautomatic cruise flight. During an automatictransition from forward flight to a hover, theindicator displays speed drift and rate of descent.

The hover indicator contains scale incrementmarks across the center vertical and horizontalaxes and along the vertical (left side) and bottomof the dial face. Two movable bars coincide withthe center vertical and horizontal axes scale markson the dial. When the aircraft is hovering, thesebars intersect at a small circle on the face. Thereare two arrowhead pointers. One pointer is on theleft side of the indicator, which moves verticallyup or down, coinciding with the vertical scale, Theother pointer is located at the bottom of the

Figure 8-39. -Attitude demodulator.

or right,

he lowerleft edge of the hover indicator case marked withpositions A, C, and D. The hover indicator alsocontains a mode selector window on the dial facethat operates with the mode selector switch. Oneof the letters (A, C, or D) is displayed here to showthe mode of operation.

Operation in the A mode connects the hoverindicator to ASE. With the meter selector on thechannel monitor panel in ASE position, the hoverindicator operates as a null indicator. The hoverindicator horizontal bar monitors the pitchchannel. The vertical bar monitors the rollchannel; the vertical pointer, the altitude channel.The horizontal pointer monitors the yaw channel.

Operation in the D mode connects the hoverindicator to the Doppler radar. The horizontal barshows forward or aft velocities, and the verticalbar indicates left or right drift. Each incrementon the hover indicator horizontal and verticalscales equals 10 knots ground speed, with amaximum indication of 40 knots. The verticalpointer shows vertical velocity, with each incre-ment equal to 250 feet per minute. Full-scaledeflection is equal to 1,000 feet per minute up ordown. To show forward flight, the horizontal barmoves downward, To indicate drift, the verticalbar moves in a direction opposite to the directionof drift; therefore, the pilot flies into the bar forcorrection.

Operation in the C mode connects the hoverindicator to the cable pickoff potentiometersmounted on the sonar hoist assembly. Thehorizontal and vertical bars of the indicator then

represent cable angle, with each increment equalto about 2.5 degrees of cable angle. The verticalpointer shows sonar depth as an error signal. Thenull position represents the selected sonar searchaltitude. Each increment above and below the nullpoint represents 10 feet of error.

In both C and D) modes, the yaw pointer isinoperative and should not move. An OFF flagon the upper dial face of the hover indicator canappear in all three modes of operation. In the Amode, the flag disappears when the ASE isengaged. In the D mode, the flag disappears whenthe Doppler signal is reliable. In the C mode, theflag disappears when the sonar equipment isturned on.

INERTIAL VELOCITY SYSTEM COM-PUTER. —The inertial velocity system computer(fig. 8-41) derives separate inertial velocity signalsfor pitch and roll ASE channels during the cableangle mode of operation. The computer containstwo separate channels for pitch and roll signals.It receives inputs from the pitch and rollaccelerometers, the vertical gyro, and the Dopplersystem.

ASE operation is controlled by the sonarsystem when the transducer is submerged. TheDoppler velocity signals are routed past theinertial velocity signals and then delayed for 40seconds. The Doppler velocity signals thenintegrate with the accelerometer and vertical gyrosignals, resulting in an inertial velocity signal. Thissignal represents both short-term velocity sensedby the accelerometer and long-term velocityproduced by the delayed Doppler signal.

ALTITUDE CONTROLLER. —The altitudecontroller (fig. 8-42) senses changes in barometric

Figure 8-41.-Inertial velocity system computer.

Figure 8-42.-Altitude controller.

altitude from a reference altitude and generatesa voltage proportional to the altitude difference.The excitation voltage and the output voltage willbe 180 degrees out of phase at altitudes below theestablished reference. They will be in phase ataltitudes above the established reference.

A purifier chamber lessens the moisturecontent in the altitude controller by routing theairflow through a desiccant cartridge in thischamber.

CYCLIC STICK. —The pilot uses the cyclicstick (fig. 8-43) to control the direction in which

Figure 8-43.-Cyclic stick grip.

the helicopter moves. By moving the cyclic stickin any direction, the pilot causes the tip-path planeof rotation of the rotary wing blades to tilt in thatdirection. As a result, the helicopter moves in thesame direction.

COLLECTIVE PITCH LEVER. —The pilotuses the collective pitch lever (fig. 8-44) tochange the pitch of all the rotary-wing bladessimultaneously. This angular change is equal onall the main rotor blades. Changing the pitch willincrease or decrease the lift accordingly.

RUDDER PEDALS. —The pilot uses therotary rudder pedals to change the pitch andthrust of the rotary rudder, thus changing the

heading of the helicopter. By pressing the leftpedal, the pilot causes the rotary rudder bladepitch to increase. This increases blade thrust andturns the helicopter to the left. By pressing theright pedal, the pilot causes the rotary rudderblade pitch to decrease. This decreases thrust andlets the rotor torque reaction turn the helicopterto the right.

ASE Sensors

ASE signal voltages for controlling helicoptermovement come from the following sensors andassociated components:

• Collective position sensor and clutch

• Pitch-and-roll position sensors

• Accelerometers

• Gyros

• Tilt table

COLLECTIVE POSITION SENSOR ANDCLUTCH. —The electromechanical collective clutchnulls the collective position sensor (fig. 8-45, view A)

Figure 8-44.-Collective pitch lever grip. Figure 8-45.-Position sensors.

output when barometric altitude is disengaged. Theclutch consists of a stationary coil and self-centeringsprings. The input male shaft can rotate continuouslywhen disengaged. The output female shaft is spring-centered and follows, without slip, any input shaftrotation through a minimum of 45 degrees of rotation.The rotation can be in either direction from the 0-degree de-energized position. The clutch has nocommon ground between the coil and case and has anoutput zero rest position repeatability of one-quarterdegree.

PITCH-AND-ROLL POSITION SENSORS. —The pitch-and-roll position sensors (fig. 8-45, views Band C) are identical single-phase, salient-pole,variable-induct ion transformers. The sensor rotoracts as the primary and the stator acts as thesecondary. The stator is nonsinusoidally wound. Thestator windings are isolated for practically unlimitedsensor resolution.

As the flight controls mechanically rotate therotor, a signal is produced. This signal is the sensoroutput required to cancel the vertical gyro signals inthe pitch and roll channels. The signal

Figure 8-46.-Stick trim valve and schematic diagram.

dampens and repositions the collective stickduring open-loop operation in the altitudechannel. The pitch-and-roll sensors mechanicallylink to the cyclic sticks. The collective sensormechanically links to the collective stick throughthe collective clutch.

ACCELEROMETERS. —Pitch, roll, andvertical accelerometers provide signal voltagesproportional to linear input acceleration. Thesignals are phased to show the direction ofacceleration.

RATE GYRO. —The rate gyro senses theturning rate of the helicopter and produces an acoutput signal proportional to the rate of turn. Theac output signal goes to the yaw channel and alsoto the attitude indicating system.

TILT TABLE. —The tilt table consists of thestarboard vertical gyro; the pitch, roll, and verticalaccelerometers; the dc rate gyro; and the starboardrate switching gyro. These components providesignals to ASE and associated equipment pro-portional to direction and magnitude. Duringmaintenance procedures, the tilt table com-ponents, along with the mobility of the table, alsoprovide simulated flight conditions in pitch androll. Changes in pitch and roll provide accurate

displacement rate and acceleration signals to theASE amplifier and associated equipment. Thesesignals can then be measured as a knowncalibrated output.

Hydraulic Components

The hydraulic components actuated andcontrolled by ASE signals include stick trim valvesand servo valves.

STICK TRIM VALVE. —There are two sticktrim valves (fig. 8-46), one for pitch and one forroll. These valves are continuous-duty, dual-inputtype solenoid valves. The valves reposition thecyclic stick when the pilot actuates the trim switch.When the release relay de-energizes, all foursolenoids on the trim valves energize. Whenthe release relay energizes, each solenoid isindependently controlled when the operator usesthe cyclic stick trim switch.

ASE SERVO VALVES. —Error signalscaused by changes in pitch and roll are sensed atthe servo valves. The servo valve (fig. 8-47)converts the error signal into a mechanical output,making the auxiliary power piston move. Thisaction affects the primary servo valve and changesthe pitch or roll attitude of the rotary-wing blades.

Figure 8-47.-Servo valve.

PRINCIPLES OF OPERATION OF ASECONTROL CHANNELS

ASE stabilizes the helicopter at referenceattitudes (pitch and roll), altitude (collective), andheading (yaw) selected by the pilot. ASE may beengaged before takeoff and remain engagedduring an entire mission. Stability corrections areintroduced into the flight control system so thepilot always has complete control of the helicopterthrough normal use of the primary flight controls.The flight control system consists of basic ASEand coupler systems. Basic ASE has pitch, roll,collective, and yaw stabilization channels; and thecoupler has related pitch, roll, and collectivestabilization channels. To stabilize the helicopterduring basic ASE and coupler operation, signalsrepresenting actual attitudes, altitude, and headingare continuously fed into the ASE amplifier.These signals are compared with correspondingsignals representing reference attitudes, altitude,and heading. The amplifier ties together all sensorand control panel input signals to produce stabilitycorrection output signals.

ASE compensates for any combination oftemporary or continuous aerodynamic dis-turbances met by the helicopter. The attitudeindicating system uses signals from the ASE rategyro, the vertical gyros, and a dc rate gyro. Thesesignals provide the pilot and copilot with pitchand roll attitude and rate-of-turn information.

ASE has time delay, system interface, andsystem interlock circuits. These circuits ensure thecompatibility of modes and associated equipmentwhen these modes and equipment are selected foruse with ASE.

ASE has four basic channels of operation—pitch, roll, collective, and yaw. These channelsare discussed in the following paragraphs.

Pitch Channel

Automatic pitch attitude stabilization ismaintained continuously after ASE engagement.The vertical gyro produces a signal proportionateto displacement of the helicopter fuselage fromthe horizon (fig. 8-48). This signal also goesthrough a rate network to produce a signalequivalent to the rate of change of pitch attitude.The direction of the displacement and rate signalscause the rotor blades to produce lift, returningthe fuselage to the desired attitude.

A combination of CG trim and cyclic stickposition serves to provide the reference attitude.In a balanced state, the following occurs: The

vertical gyro displacement and rate signals areequal and opposite to the CG trim/position sensorsignal at the summing point. Therefore, there isno command for flight control movement. If thepitch attitude changes without the cyclic stickbeing moved, the vertical gyro senses themovement. Signals from the vertical gyro thencause the aircraft to return to its original pitchattitude. If the pilot desires to change the pitchattitude, the cyclic stick is moved forward or aft.This action produces an error signal at thesumming point. As the helicopter changes attitudeto correspond with the cyclic stick position, thevertical gyro pitch displacement signal increasesuntil the two signals are equal and opposite again.This becomes the new pitch reference attitude.

During CPLR operation in the DOPP mode,the system compares Doppler ground speed withthe speed selected on the ASE control panel. Anydifference between the actual ground speed andthe selected ground speed produces an errorsignal. The error signal causes the helicopter tochange speed until the two voltages balance. Anaccelerometer monitoring the change in speedproduces a rate signal to oppose any change inspeed. This signal prevents large rapid changes inpitch attitude. The hover trim control panel allowsthe hoist operator to provide small attitudechanges in hovering flight.

The integrator circuits monitor any steady-state error signal that may exist because of thedifference between Doppler speed and selectedspeed (could be caused by head or tail winds). Thelarger the output of the integrator, the morecorrection is made until the Doppler speed equalsselected speed.

In the cable angle mode, pitch attitude fromthe vertical gyro is compared with the cable angle(in respect to the deck of the helicopter). Thisallows the ASE to maintain the sonar cablevertical to the earth’s surface regardless of aircraftattitude. The cable angle error signal is processedin a similar manner as the Doppler error signal.

An inertial velocity system (IVS) circuitbecomes operational in the cable angle mode whenthe IVS control relay energizes. The voltage toenergize the relay develops when the sonartransducer submerges to a depth below 11 feet.When the IVS circuit is operational, outputs fromthe accelerometer and the vertical gyro combinewith the Doppler velocity signal. This combinationproduces a very accurate inertially derived signal.

The output of the coupler also goes to a sticktrim actuator circuit. When the coupler outputreaches a certain level, a solenoid of the pitch

Figure 8-48.-ASE/coupler pitch channel.

stick trim valve is excited. The solenoid causesautomatic repositioning of the cyclic stick, therebyextending pitch channel authority.

Roll Channel

Roll channel operation is similar to the pitchchannel. (See fig. 8-49.) In basic ASE, the positionsensor is lagged by an amplifier before beingadded to or subtracted from the rate-plus-displacement signal. In the Doppler mode, thepilot input is the DRIFT control on the ASEcontrol panel. Roll signals are otherwise processedin the same manner as the pitch signals.

Collective Channel

After ASE is engaged, the pilot engages thecollective channel by the BAR ALT engage buttonto stabilize helicopter altitude. Part of thealtitude error signal goes to the collective couplerintegrator during BAR ALT operation. (Seefigure 8-50.) If the altitude control error signalis steady-state (due to variables such as fuel,load, weather, etc.), a signal is developedto compensate for this steady-state error andto improve engaged altitude retention. Whenthe pilot presses the BAR ALT REL button, theintegrator cancels any remaining compensatingsignal.

Figure 8-49.-ASE/coupler run channel.

As the servo valve corrects the altitude errorsignal through the auxiliary power piston androtary wing blades, the altitude error signaldiminishes. When the helicopter returns to theengaged altitude, the altitude control output errorsignal returns to null. However, as the servo valvecorrects the altitude error signal, it places thecollective servo system in an open-loop condition.

The open-loop spring and a balance springsupport the collective stick and linkage weight sothe stick retains the desired position. The stickshould remain in this position with little or noexternal force applied to hold the stick. The

open-loop spring also allows the barometricaltitude control to make a larger than 10 percentcorrection over a period of time. The pilot mayoverride open-loop operation by applying anopposing force on the collective stick. As thealtitude error signal diminishes, the positionsensor signal returns the collective pitch controlstick to the engaged altitude position. While BARALT is on, the pilot applies collective anddepresses the BAR REL button on the collectivestick grip momentarily to change altitude. Thecollective clutch nulls the output of the positionsensor and the barometric altitude input while theBAR REL button is depressed.

Figure 8-50.-ASE/coupler collective channel.

During collective coupler operation with radaraltitude mode engaged, the coupler seeks andretains the altitude selected on the control panel.Any difference between radar altitude and theselected altitude produces an error voltage. Thisvoltage causes the total lift developed by the rotorblades to change and seek the selected altitude.The radar altitude error signal is also fed to theintegrator amplifier circuit. This circuit providesa signal to compensate for a steady-state errorsignal input. The compensating signal is added tothe other portion of the radar altitude error signal.The combined signal then couples to an outputfade-in circuit.

The radar navigation set produces a signal(vertical velocity) proportional to the rate ofchange in altitude. This rate signal is laggedapproximately 1 second in the lag amplifier. Therate signal then goes through the ASE controlpanel, and couples to the output fade-in circuit.The coupler altitude error signal is coupled to thecollective ASE amplifier and processed in a similarmanner as the basic ASE altitude error signal.

In the cable attitude mode, the cable altitudeoutput signal is compared to the radar altitudeoutput signal. Any difference between signalsgoes to the follow-up integrator circuit. Thefollow-up output signal is fed back to the cablealtitude input to correct the cable altitude signal.The corrected cable altitude input signal issummed with the ALTITUDE control. Theresultant signal is coupled through the isolation

amplifier to the fade-in circuit. The isolationamplifier output is summed with the verticalaccelerometer signal, and the resultant is com-bined with the integrator amplifier output.

When operating in the collective couplervertical accelerometer mode, the cable signal isdisconnected. In this mode only the verticalaccelerometer and radar altitude signals areapplied to the output fade-in or fade-out circuit.During collective coupler operation, accidentalmomentary disengagement of the altitude controlis prevented by a diode in the ASE control panel.This diode makes the momentary BAR RELbutton inoperative.

Yaw Channel

When the ASE is on, the yaw channelstabilizes the heading of the helicopter. Thecompass system sends a heading signal to adifferential synchro in the ASE control panel.(See figure 8-51.) The physical position of thedifferential synchro is controlled by the ASEcontrol panel YAW TRIM control. The outputof the differential synchro is coupled to the yawsynchronizer module. Depending on the modeof the yaw channel, the signal becomes asynchronizing signal or a heading error signal.During the synchronizing mode (manual turnsusing cyclic stick and rudder pedals), headingsignals are nulled out and no error signal iscoupled to the yaw ASE amplifier. When the yaw

Figure 8-51.-ASE yaw channel.

channel is operating, heading error signals areadded to a rate signal developed by the yaw rategyro. The resultant signal goes through thechannel monitor control panel as differentialcurrent flow. The differential current flowprovides outputs to excite the yaw servo valve.This signal also couples to the hover indicators.

As the servo valve corrects the heading errorsignal through the auxiliary power piston androtary rudder blades, the heading error signaldiminishes. When the helicopter returns to itsreference heading, the compass system outputsignals return to their original value. As the servovalve corrects large heading error signals, it placesthe yaw servo system in an open-loop condition.The open-loop condition enables the yaw channelto make larger heading corrections if necessary.The pilot may override the open-loop operationby applying an opposing force on the pedals.

If the pilot uses the pedals to turn manuallywhile ASE is engaged, an artificial force is felt,which reduces the tendency to overcontrol. Theyaw rate gyro produces a signal proportional tothe turning rate of the helicopter. This rate signalis eventually coupled to the servo valve and placesthe yaw servo system in an open-loop condition,opposite to the turn direction.

APPENDIX I

GLOSSARY

ACCELEROMETER —A device that measuresthe acceleration to which it is subjected anddevelops a signal proportional to it.

AGONIC —An imaginary line of the earth’ssurface passing through points where the magneticdeclination is 0O degrees; that is, points where thecompass points to true north.

AMBIENT CONDITIONS —Physical condi-tions of the immediate environment, which maypertain to temperature, humidity, pressure, etc.

AMPERE —The basic unit of electricalcurrent.

AMPERE-TURN —The magnetizing forceproduced by a current of 1 ampere flowingthrough a coil of 1 turn.

AMPLIDYNE —A rotary magnetic ordynamoelectric amplifier used in servomechanismand control applications.

AMPLIFICATION —The process of increasingthe strength (current, power, or voltage) of asignal.

AMPLIFIER —A device used to increase thesignal voltage, current, or power, generallycomposed of solid-state circuitry called a stage.It may contain several stages in order to obtaina desired gain.

AMPLITUDE —The maximum instantaneousvalue of an alternating voltage or current,measured in either the positive or negativedirection.

ANALOG COMPUTER —A type of computerthat provides a continuous solution to a mathe-matical problem with continuously changinginputs. Inputs and outputs are represented byphysical quantities that may be easily generatedor controlled.

APPARENT DRIFT —The effect of theearth’s rotation on a gyro that causes the spinningaxis to appear to make one complete rotation in1 day. Also called APPARENT PRECESSIONor APPARENT ROTATION.

APPARENT PRECESSION —See APPAR-ENT DRIFT.

APPARENT ROTATION —See APPARENTDRIFT.

ARC —A flash caused by an electric currentionizing a gas or vapor.

ARMATURE —The rotating part of an elec-tric motor or generator. The moving part of arelay or vibrator.

ATTENUATOR —A network of resistorsused to reduce voltage, current, or power deliveredto a load.

ATTRACTION —The force that tends tomake two objects approach each other. Attractionexists between two unlike magnetic poles(north and south) or between two unlike staticcharges.

AUTOTRANSFORMER —A transformer inwhich the primary and secondary are connectedtogether in one winding.

AVB —Avionic bulletin.

AVC —Avionic change.

AXIS —A straight line, either real orimaginary, passing through a body around whichthe body revolves.

AZIMUTH —Angular measurement in thehorizontal plane in a clockwise direction.

BATTERY —Two or more primary orsecondary cells connected together electrically.The term does not apply to a single cell. A devicefor converting chemical energy into electricalenergy.

BATTERY CAPACITY —The amount ofenergy available from a battery. Battery capacityis expressed in ampere-hours.

BIAS —In vacuum tubes, the difference ofpotential between the control grid and thecathode; in transistors, the difference of potentialbetween the base and emitter and between the baseand collector; in magnetic amplifiers, the level offlux density in the core under no-signal conditions.

BLOCK DIAGRAM —A diagram in whichthe major components of an equipment or asystem are represented by squares, rectangles, orother geometric figures, and the normal order ofprogression of a signal or current flow isrepresented by lines.

BRIDGE CIRCUIT —The electrical bridgecircuit is a term referring to any one of a varietyof electric circuit networks, one branch of which(the bridge proper) connects two points of equalpotential; therefore, it carries no current when thecircuit is properly adjusted or balanced.

BRUSH —The conducting material, usually ablock of carbon, bearing against the commutatoror slip rings through which the current flows inor out.

BUS BAR —A primary power distributionpoint connected to the main power source.

CABLE HARNESS —A group of wires orribbons of wiring used to interconnect electronicsystems and subsystems.

CAGING (GYRO) —The act of holding a gyroso that it cannot precess or change its attitude withrespect to the body containing it.

CAPACITOR —Two electrodes or sets ofelectrodes in the form of plates, separated fromeach other by an insulating material called thedielectric.

CAPACITOR-START MOTOR —A type ofsingle-phase, ac induction motor in which astarting winding and a capacitor are placed in

series to start the motor. The values of andR are such that the main-winding and starting-winding currents are nearly 90 degrees apart, andthe starting torque is produced as in a two-phasemotor.

CARDIOPULMONARY RESUSCITATION –Procedure designed to restore breathing aftercardiac arrest. Includes clearing air passages tolungs and heart massage.

CELL —A single unit that transforms chemi-cal energy into electrical energy. Batteries aremade up of cells.

CHOKE COIL —A coil of low ohmicresistance and high impedance to alternatingcurrent.

CIRCUIT —The complete path of an electriccurrent.

CIRCUIT BREAKER —An electromagneticor thermal device that opens a circuit when thecurrent in the circuit exceeds a predeterminedamount. Circuit breakers can be reset.

CIRCULAR MlL —An area equal to that ofa circle with a diameter of 0.001 inch. It is usedfor measuring the cross section of wires.

COAXIAL CABLE —A transmission lineconsisting of two conductors concentric with andinsulated from each other.

COMMUTATION —The act of a commutatorin converting generator output from an ac voltageto a dc voltage.

COMMUTATOR —The copper segments onthe armature of a motor or generator. It iscylindrical in shape and is used to pass power intoor from the brushes. A mechanical device thatreverses armature connections in motors andgenerators at the proper instant so that currentcontinues to flow in only one direction. In effect,the commutator changes ac to dc.

COMPARATOR —A circuit that comparestwo signals or values, and indicates agreement orvariance between them.

COMPENSATING WINDINGS —Windingsembedded in slots in pole pieces, connected inseries with the armature, whose magnetic fieldopposes the armature field and cancels armaturereaction.

COMPOUND-WOUND MOTORS ANDGENERATORS —Machines that have a seriesfield in addition to a shunt field. Such machineshave characteristics of both series- and shunt-wound machines.

CONDUCTANCE —The ability of a materialto conduct or carry an electric current. It is thereciprocal of the resistance of the material, andis expressed in mhos.

CONDUCTIVITY —The ease with which asubstance transmits electricity.

CONDUCTOR —Any material suitable forcarrying electric current.

CORE —A magnetic material that affords aneasy path for magnetic flux lines in a coil.

COUNTER EMF —Counter electromotiveforce; an EMF induced in a coil or armature thatopposes the applied voltage.

COUNTING CIRCUIT —A circuit thatreceives uniform pulses representing units to becounted and produces a voltage in proportion totheir frequency.

COVALENT BOND —A type of linkagebetween atoms in which the atoms share valenceelectrons.

CPR —See CARDIOPULMONARY RESUS-CITATION.

CURRENT —The movement of electrons pasta reference point. The passage of electronsthrough a conductor. Measured in amperes.

CURRENT LIMITER —A protective devicesimilar to a fuse, usually used in high amperagecircuits.

CYCLE —One complete positive and onecomplete negative alternation of a current orvoltage.

D’ARSONVAL METER MOVEMENT —The permanent-magnet moving-coil movementused in most meters.

DEGREES OF FREEDOM (GYRO) —A termapplied to gyros to describe the number ofvariable angles required to specify the position ofthe rotor spin axis relative to the case.

DELTA —A three-phase connection in whichwindings are connected end to end, forming aclosed loop that resembles the Greek letter delta.A separate phase wire is then connected to eachof the three junctions.

DEMODULATOR —A circuit used in servosystems to convert an ac signal to a dc signal, Themagnitude of the dc output is determined by themagnitude of the ac input signal, and its polarityis determined by whether the ac input signal is inor out of phase with the ac reference voltage.

DIELECTRIC —An insulator; a term thatrefers to the insulating material between the platesof a capacitor.

DIFFERENTIAL —A mechanical computingdevice used to add or subtract two quantities.

DIGITAL COMPUTER —A type of com-puter in which quantities are represented innumerical form. It is generally made to solvecomplex mathematical problems by use of thefundamental processes of addition, subtraction,multiplication, and division. Its accuracy islimited only by the number of significant figuresprovided.

DIODE —A material of either germanium orsilicon that is manufactured to allow current toflow in only one direction. Diodes are used asrectifiers and detectors.

DIRECT CURRENT —An electric currentthat flows in one direction only.

DISCRIMINATOR —A dual-input circuit inwhich the output is dependent on the variationof one input from the other input or from anapplied standard.

DOPPLER EFFECT —An apparent change inthe frequency of a sound wave or electromagneticwave reaching a receiver when there is relativemotion between the source and the receiver.

EDDY CURRENT —Induced circulatingcurrents in a conducting material that are causedby a varying magnetic field.

EFFICIENCY —The ratio of output power toinput power, generally expressed as a percentage.

ELECTROLYSIS —A type of corrosion(chemical, decomposition) caused by current flowresulting from contact of dissimilar metals.

ELECTROLYTE —A solution of a substancethat is capable of conducting electricity. Anelectrolyte may be in the form of either a liquidor a paste.

ELECTROMAGNET —A magnet made bypassing current through a coil of wire wound ona soft iron core.

ELECTROMOTIVE FORCE (EMF) —Theforce that produces an electric current in a circuit.

ELECTRON —A negatively charged particleof matter.

ELECTRON SHELL —A group of electronsthat have a common energy level that forms partof the outer structure (shell) of an atom.

ENERGY —The ability or capacity to dowork.

EQUIVALENT CIRCUIT —A diagrammaticarrangement of component parts representing, insimplified form, the effects of a more complicatedcircuit to permit easier analysis.

ERECTING (A GYRO) —The positioning ofa gyro into a desired position and the maintainingof that position.

ERROR SIGNAL —(1) In servo systems, thesignal whose amplitude and polarity or phase areused to correct the alignment between thecontrolling and the controlled elements. (2) Thename given to the electrical output of a controltransformer.

E-TRANSFORMER —A magnetic device withan E configuration, used as an error detector.

FARAD —The unit of capacitance.

FEEDBACK —A transfer of energy from theoutput circuit of a device back to its input.

FIELD —The space containing electric ormagnetic lines of force.

FIELD WINDING —The coil used to providethe magnetizing force in motors and generators.

FLUX —(1) In electrical or electromagneticdevices, a general term used to designate collec-tively all the electric or magnetic lines of force ina region, (2) A solution that removes surfaceoxides from metals being soldered.

FLUX DENSITY —The number of magneticlines of force passing through a given area.

FLUX FIELD —All electric or magnetic linesof force in a given region.

FREE ELECTRONS —Electrons that areloosely held; consequently, they tend to move atrandom among the atoms of the material.

FREE GYRO —A gyro so gimbaled that itassumes and maintains any attitude in space. Thefree gyro has two degrees of freedom; torquecannot be applied to the rotor of a truly free gyro.

FREQUENCY —The number of completecycles per second existing in any form of wavemotion, such as the number of cycles per secondof an alternating current.

FULL-WAVE RECTIFIER CIRCUIT —Acircuit that uses both the positive and the negativealternations of an alternating current to producea direct current.

FUSE —A protective device inserted in serieswith a circuit. It contains a metal that will meltor break when current is increased beyond aspecific value for a definite period of time.

GAIN —The ratio of the output power,voltage, or current to the input power, voltage,or current, respectively.

GALVANOMETER —An instrument used tomeasure small dc currents.

GENERATOR —A machine that convertsmechanical energy into electrical energy.

GIMBAL —A frame in which the gyro wheelspins, and that allows the gyro wheel to havecertain freedom of movement. It permits the gyromotor to incline freely and retain that positionwhen the support is tipped or repositioned.

GROUND —A metallic connection with theearth to establish ground potential. Also, acommon return to a point of zero potential.

GROUND POTENTIAL —Zero potentialwith respect to the ground or earth.

GYROSCOPE —A wheel or disk so mountedas to spin rapidly about one axis and be free tomove about one or both of the two axes mutuallyperpendicular to the axis of spin.

HALF-WAVE RECTIFIER —A rectifierusing only one-half of each cycle to change ac topulsating dc.

HEAT SHUNT —A device (preferably aclip-on type) used to absorb heat and protectheat-sensitive components during soldering.

HERO —Hazardous electromagnetic radiationto ordnance.

HERTZ —A unit of frequency equal to onecycle per second.

HMI —Handbook Maintenance Instructions.

HORSEPOWER —The English unit of power,equal to work done at the rate of 550 foot-poundsper second. Equal to 746 watts of electrical power.

HYSTERESIS —A lagging of the magneticflux in a magnetic material behind the magnetizingforce that is producing it.

Hz —See HERTZ.

IMPEDANCE —The total opposition offeredto the flow of an alternating current. It mayconsist of any combination of resistance, inductivereactance, and capacitive reactance.

INDUCED CURRENT —Current caused bythe relative motion between a conductor and amagnetic field.

INDUCTANCE —The property of a circuitthat tends to oppose a change in the existingcurrent.

INDUCTION —The act or process ofproducing voltage by the relative motion of amagnetic field across a conductor.

INDUCTION MOTOR —A simple, rugged,ac motor with desirable characteristics. The rotoris energized by transformer action (induction)from the stator. Induction motors are used morethan any other type.

INDUCTIVE REACTANCE —The opposi-tion to the flow of alternating or pulsating currentcaused by the inductance of a circuit. It ismeasured in ohms.

INERTIA —The physical tendency of a bodyin motion to remain in motion and a body at restto remain at rest unless acted upon by an outsideforce (Newton’s first law of motion).

lNFINITE —(1) Extending indefinitely,endless. (2) Boundless, having no limits, (3) Anincalculable number.

IN PHASE —This term is applied to thecondition that exists when two waves of the samefrequency pass through their maximum andminimum values of like polarity at the sameinstant.

lNSULATION —A material used to preventthe leakage of electricity from a conductor andto provide mechanical spacing or support asprotection against accidental contact with theconductor.

INTEGRATING CIRCUIT —A circuit whoseoutput voltage is proportional to the product ofthe instantaneous applied input voltages and theirdurations.

INTEGRATOR —A computing device usedfor summing up an infinite number of minutequantities.

INVERSELY —Inverted or reversed inposition or relationship.

ISOGONIC LINE —An imaginary line drawnthrough points on the earth’s surface where themagnetic variation is equal.

JOULE. —A unit of energy or work. A jouleof energy is liberated by 1 ampere flowing for 1second through a resistance of 1 ohm.

JUNCTION —(1) The connection betweentwo or more conductors, (2) The contact betweentwo dissimilar metals or materials, as in athermocouple.

JUNCTION BOX —A box with a cover thatserves the purpose of joining different runs of wireor cable and provides space for the connectionand branching of the enclosed conductors.

KINETIC ENERGY —Energy that a masspossesses by virtue of its motion.

KNEE (OF A CURVE) —An abrupt changein direction between two fairly straight segmentsof a curve.

LAG —The amount one wave is behindanother in time; expressed in electrical degrees.

LAMINATED CORE —A core built up fromthin sheets of metal and used in transformers andrelays.

LEAD —The opposite of LAG. Also, a wireor connection.

LEAD-ACID CELL —A cell in an ordinarystorage battery in which electrodes are grids oflead containing an active material consisting ofcertain lead oxides that change in compositionduring charging and discharging. The electrodesor plates are immersed in an electrolyte of dilutedsulfuric acid.

LINE OF FORCE —A line in an electric ormagnetic field that shows the direction of theforce.

LOAD —The power that is being delivered byany power-producing device. The equipment thatuses the power from the power-producing device.

LOGIC CIRCUITS —Digital computercircuits used to store information signals and/orto perform logical operations on those signals.

MAGNETIC AMPLIFIER —A saturablereactor-type device that is used in a circuit toamplify or control.

MAGNETIC CIRCUIT —The complete pathof magnetic lines of force.

MAGNETIC FIELD —The space in which amagnetic force exists.

MAGNETIC FLUX —The total number oflines of force issuing from a pole of a magnet.

MAGNETIZE —To convert a material into amagnet by causing the molecules to rearrange.

MAGNETO —A generator that producesalternating current and has a permanent magnetas its field.

MATTER —Any physical entity that possessesmass.

METEOROLOGY AUTOMATED SYSTEMFOR UNIFORM RECALL AND REPORTING(MEASURE) —The Navy data processing systemdesigned to provide a standardized system for therecall, scheduling, and documenting of testequipment into calibration facilities.

MEGGER. —A test instrument used tomeasure insulation resistance and other highresistances. It is a portable hand-operated dcgenerator used as an ohmmeter.

MEGOHM —A million ohms.

METER —A device used to measure a specificquantity, such as current, voltage, or frequency.

METER MOVEMENT —The part of themeter that moves to indicate some value.

METER SHUNT —A resistor placed inparallel with the meter terminals; used to provideincreased range capability.

MICRO —A prefix meaning one-millionth.

MICROFICHE —A film negative card (fiche)developed for many purposes throughout theNavy where microfilming is used to reduceamounts of paper documents.

MICROMETER —A unit of length equal to10 -6 meter. Formerly a micron.

MICRON —See MICROMETER.

MIL —The diameter of a conductor equal to1/1000 (.001) inch.

MIL-FOOT —A unit of measurement forconductors (diameter of 1 mil, 1 foot in length).

MILITARY SPECIFICATIONS (MILSPEC) —Technical requirements and standards adopted bythe Department of Defense (DoD) that must bemet by vendors selling materials to DoD.

MILITARY STANDARDS (MILSTD) —Standards of performance for components orequipment that must be met to be acceptable formilitary systems.

MILLI —A prefix meaning one-thousandth.

MILLIAMMETER —An ammeter thatmeasures current in thousandths of an ampere.

MOTOR —A machine that converts electricalenergy to mechanical energy. It is activated by acor dc voltage, depending on the design.

MOTOR-GENERATOR —A motor and agenerator with a common shaft used to convertline voltages to other voltages or frequencies.

MULTICONDUCTOR —More than one con-ductor, as in a cable.

MULTIMETER —A single meter combiningthe functions of an ammeter, a voltmeter, and anohmmeter.

MULTIPHASE —See POLYPHASE.

MUTUAL INDUCTANCE —A circuit propertyexisting when the relative position of twoinductors causes the magnetic lines of force fromone to link with the turns of the other.

NAMP —The Naval Aviation MaintenanceProgram.

NANOMETER —A unit of length equal to10 -9 meter. Formerly millimicron.

NEGATIVE CHARGE —The electricalcharge carried by a body that has an excess ofelectrons.

NEGATIVE FEEDBACK —Feedback inwhich the feedback signal is out of phase with theinput signal.

NEGATIVE TEMPERATURE COEFFI-CIENT —A characteristic of a semiconductormaterial, such as silver sulfide, in which resistanceto electrical current flow decreases as temperatureincreases.

NEUTRON —A particle having the weight ofa proton but carrying no electric charge. It islocated in the nucleus of an atom.

NOISE —Any undesired disturbance withinthe useful frequency band; also, that part of themodulation of a received signal (or an electricalor electronic signal within a circuit) representingan undesirable effect of transient conditions.

NUCLEUS —The central part of an atom thatis mainly made up of protons and neutrons.It is the part of the atom that has the mostmass.

NULL —A point or position where a variable-strength signal is at its minimum value (orzero).

OHM —The unit of electrical resistance. Thatvalue of electrical resistance through which aconstant potential difference of 1 volt across theresistance will maintain a current flow of 1 amperethrough the resistance.

OHM’S LAW —The current in an electricalcircuit is directly proportional to the electromotiveforce in the circuit. The most common form ofthe law is E = IR, where E is the electromotiveforce or voltage across the circuit, I is the currentflowing in the circuit, and R is the resistance ofthe circuit.

OVERLOAD —A load greater than the ratedload of an electrical device.

PARAMETERS —In electronics, the design oroperating characteristics of a circuit or device.

PERMALLOY —An alloy of nickel and ironhaving an abnormally high magnetic permeability.

PERMEABILITY —A measure of the easewith which magnetic lines of force can flowthrough a material as compared to air.

PHASE —The angular relationship betweentwo alternating currents or voltages when thevoltage or current is plotted as a function of time.When the two are in phase, the angle is zero; bothreach their peak simultaneously. When out ofphase, one will lead or lag the other; that is, atthe instant when one is at its peak. The otherphase will not be at peak value and (dependingon the phase angle) may differ in polarity as wellas magnitude.

PHASE DIFFERENCE —The time in elec-trical degree by which one wave leads or lagsanother.

PHOTON (hv) —An elementary quantity ofradiant energy (quantium) whose value is equalto the product of Plank’s constant and thefrequency of the electromagnetic radiation.

PICKOFF —In gyros, a sensing device thatmeasures the angle of the spin axis with respectto its reference and provides an error signal thatindicates the direction and (in most cases) themagnitude of the displacement,

POLARITY —The character of having mag-netic poles, or electric charges.

POLYPHASE —A circuit that uses more thanone phase of alternating current.

POSITIVE CHARGE —The electrical chargecarried by a body that has become deficient inelectrons.

POSITIVE FEEDBACK —Feedback in whichthe feedback signal is in phase with the inputsignal.

POSITIVE TEMPERATURE COEFFI-CIENT —The characteristic of a conductor inwhich the resistance increases as temperatureincreases.

POTENTIAL —The amount of charge held bya body as compared to another point or body.Usually measured in volts.

POTENTIOMETER —A variable voltagedivider; a resistor that has a variable contact armso that any portion of the potential appliedbetween its ends may be selected.

POWER —The rate of doing work or the rateof expending energy. The unit of electrical poweris the watt.

POWER FACTOR —The ratio of the actualpower of an alternating or pulsating current, asmeasured by a wattmeter, to the apparent power,as indicated by ammeter and voltmeter readings.The power factor of an inductor, capacitor, orinsulator is an expression of their losses.

POWER SUPPLY —A unit that supplieselectrical power to another unit. It changes ac todc and maintains a constant voltage output withinlimits.

PRECESSION —The reaction of a gyro to anapplied torque, which causes the gyro to tilt itselfat right angles to the direction of the appliedtorque in such a manner that the direction of spinof the gyro rotor will be in the same direction asthe applied torque.

PRIMARY WINDING —The winding of atransformer connected to the electrical source.

PRIME MOVER —The source of the turningforce applied to the rotor of a generator, This maybe an electric motor, a gasoline engine, a steamturbine, and so forth.

PROTON —A positively charged particle inthe nucleus of an atom.

RADAR —An acronym for radio detectingand ranging.

RADAR ALTIMETER —Airborne radar thatmeasures the distance of the aircraft above theground.

RADIAN —In a circle, the angle includedwithin an arc equal to the radius of the circle. Acomplete circle contains 2 radians. One radianequals 57.3 degrees, and 1 degree equals 0.01745radian.

RATE GYRO —A gyro with one degree offreedom that has an elastic restraint, with orwithout a damper, and whose output will beproportional to the rate of the applied torque.

RATIO —The value obtained by dividing onenumber by another, indicating their relativeproportions.

REACTANCE —The opposition offered tothe flow of an alternating current by theinductance, capacitance, or both, in any circuit.

RECTIFIERS —Devices used to changealternating current to unidirectional current.These may be vacuum tubes, semiconductors,such as germanium and silicon, and dry-diskrectifiers, such as selenium and copper oxide.

RELUCTANCE —A measure of the opposi-tion that a material offers to magnetic lines offorce.

RESISTANCE —The opposition to the flowof current caused by the nature and physicaldimensions of a conductor.

RETENTIVITY —The measure of the abilityof a material to hold its magnetism.

RHEOSTAT —A variable resistor.

RIGIDITY —In gyros, the characteristics ofa spinning body that causes it to oppose allattempts to tilt it away from the axis in which itis spinning.

ROTATING FIELD —The magnetic field ina multiphase ac motor that is the result offield windings being energized by out-of-phasecurrents. In effect, the magnetic field is made torotate electrically rather than mechanically.

ROTOR —(1) The revolving part of a rotatingelectrical machine. The rotor may be either thefield or the armature, depending on the design ofthe machine. (2) The rotating member of asynchro that consists of one or more coils of wirewound on a laminated core. Depending on thetype of synchro, the rotor functions similarlyto the primary or secondary winding of atransformer.

SATURABLE REACTOR —A control devicethat uses a small dc current to control a large accurrent by controlling core flux density.

SATURATION —The condition existing inany circuit when an increase in the driving signalproduces no further change in the resultant effect.

SCHEMATIC —A diagram that shows, bymeans of graphic symbols, the electricalconnections and functions of a specific circuitarrangement.

S E C O N D A R Y —The output coil of atransformer.

SELF-INDUCTION —The process by whicha circuit induces an EMF into itself by its ownmagnetic field.

SERIES CIRCUIT —An arrangement whereelectrical devices are connected so that the totalcurrent must flow through all the devices;electrons have one path to travel from the negativeterminal to the positive terminal.

SERIES-WOUND —A motor or generator inwhich the armature is wired in series with the fieldwinding.

SERVO —A device used to convert a smallmovement into one of greater movement or force.

SERVOMECHANISM —A closed-loopsystem that produces a force to position an objectaccording to the information that originates at theinput.

SERVOMOTOR —An ac or dc motor used inservo systems to move a load to a desired positionor at a desired speed. The ac motor is usually usedto drive light loads at a constant speed, while thedc motor is used to drive heavy loads at varyingspeeds.

SERVO SYSTEM —An automatic feedbackcontrol system that compares a required condition(desired value, position, etc.) with an actualcondition and uses the difference to drive a controldevice to achieve the required condition,

SHIELDING —(1) A metallic covering usedto prevent magnetic or electromagnetic fields fromaffecting an object. (2) Technique designed tominimize internal and external interference.

SHUNT —A resistive device placed in parallelwith another component. Appreciable currentmay flow through it, and an appreciable voltagemay exist across it.

SLIP RINGS —Contacts that are mounted onthe shaft of a motor or generator to which therotor windings are connected and against whichthe brushes ride. Devices for making electricconnections between stationary and rotatingcontacts.

SOLENOID —An electromagnetic coil thatcontains a movable plunger.

SOLID-STATE DEVICE —An electronicdevice that operates by the movement of electronswithin a solid piece of semiconductor material.

SOURCE —(1) The object that produces thewaves or disturbance. (2) The name given to theend of a two-wire transmission line that isconnected to a source. (3) The device thatfurnishes the electrical energy used by aload.

SPECIFIC GRAVITY —The ratio betweenthe density of a substance and that of pure waterat a given temperature.

STATOR —(l) The stationary part of arotating electrical machine. The stator may beeither the field or the armature, depending on thedesign of the machine. (2) The stationary memberof a synchro that consists of a cylindrical structureof slotted laminations on which three Y-connectedcoils are wound with their axes 120 degrees apart.Depending on the type of synchro, the stator’sfunctions are similar to the primary or secondarywindings of a transformer.

STRANDED CONDUCTOR —A conductorcomposed of a group of wires, The wires in astranded conductor are usually twisted togetherand not insulated from each other.

STRANDS —Fine metallic filaments twistedtogether to form a single wire.

SUBASSEMBLY —Consists of two or moreparts that form a portion of an assembly or a unit.

SUPPORT EQUIPMENT (SE) —All theequipment on the ground or ship needed tosupport aircraft in a state of readiness for flight.

SYNCHRO —A small motor-like analogdevice that operates like a variable transformerand is used primarily for the rapid and accuratetransmission of data among equipments andstations.

SYNCHRO SYSTEM —An electrical systemthat gives remote indications or control by meansof self-synchronizing motors.

TACHOMETER —An instrument for indi-cating revolutions per minute.

TEMPERATURE COEFFICIENT —Theamount of change of resistance in a material perunit change in temperature.

TERTIARY WINDING —A third winding ona transformer or magnetic amplifier that is usedas a second control winding.

THERMISTOR —A resistor that is used tocompensate for temperature variations in a circuit.See also BOLOMETER.

THERMOCOUPLE —A junction of twodissimilar metals that produces a voltage whenheated.

TINNING —The process of applying a thincoat of solder to materials prior to their beingsoldered; for example, application of a light coatof solder to the filaments of a conductor to holdthe filaments in place prior to soldering of theconductor.

TORQUE —The turning effort or twist thata shaft sustains when transmitting power. A forcetending to cause rotational motion; the productof the force applied times the distance from theforce to the axis of rotation.

TOTAL RESISTANCE (RT) —The equivalentresistance of an entire circuit, For a series circuit:R T= R1 + R2 + R3 + . . . RN . For parallelcircuits:

TRANSFORMER —A device composed oftwo or more coils, linked by magnetic lines offorce, used to transfer energy from one circuit toanother.

TRANSFORMER EFFICIENCY —The ratioof output power to input powver, generallyexpressed as a percentage.

TRANSFORMER, STEP-DOWN —A trans-former constructed so that the number of turnsin the secondary winding is less than the numberof turns in the primary winding. This constructionwill provide less voltage in the secondary circuitthan in the primary circuit.

TRANSFORMER, STEP-UP —A trans-former constructed so that the number of turnsin the secondary winding is more than the numberof turns in the primary winding. This constructionwill provide more voltage in the secondary circuitthan in the primary circuit.

TRUE BEARING —Angle between a targetand true north measured clockwise in thehorizontal plane.

TRUE NORTH —Geographic north.

TUMBLE (GYRO) —To subject a gyro to atorque so that it presents a precession violentenough to cause the gyro rotor to spin end overend.

TURNS RATIO -The ratio of the number ofturns in the primary winding to the number ofturns in the secondary winding of a transformer.

VALE NCE —The measure of the extent towhich an atom is able to combine directly withother atoms. It generally depends on the numberand arrangement of the electrons in the outermostshell of the atom.

VALENCE SHELL —The electrons that formthe outermost shell of an atom.

VECTOR —A line used to represent bothdirection and magnitude.

VOLT —The unit of electromotive force orelectrical pressure. One volt is the pressurerequired to send 1 ampere of current through aresistance of 1 ohm.

VOLTAGE —(1) The term used to signifyelectrical pressure. Voltage is a force that causescurrent to flow through an electrical conductor.(2) The voltage of a circuit is the greatest effectivedifference of potential between any two con-ductors of the circuit.

VOLTAGE DIVIDER —A series network inwhich desired portions of the source voltage maybe tapped off for use in the circuit.

WATT —The unit of electrical power.

WHEATSTONE BRIDGE —An ac bridgecircuit used to measure unknown values ofresistance, inductance, or capacitance.

WIRING DIAGRAM —A diagram that showsthe connections of an equipment or its componentdevices or parts. It may cover internal or externalconnections, or both, and contains such detail asis needed to make or trace connections that areinvolved.

WORK —The product of force and motion.

WYE (Y) —A three-phase connection in whichone end of each phase winding is connected toa common ground.

X-AXIS —In a gyro, the spin axis of the gyro.

Y-AXIS —In a gyro, an axis through the centerof gravity and perpendicular to the spin axis.

Z-AXIS -In a gyro, an axis through the centerof gravity and mutually perpendicular to both theX (spin) and Y axes.

ZENER DIODE —A PN-junction, diodedesigned to operate in the reverse-bias breakdownregion.

APPENDIX II

SYMBOLS

A I I - 3

APPENDIX III

REFERENCES

Chapter Two

Aircraft Weapons Systems Cleaning and Corrosion Control, NAVAIR01-1A-509, Naval Air Systems Command, Washington, D.C., 1 July1988.

Avionics Cleaning and Corrosion Prevention/Control, NAVAIR 16-1-540,Chg 1, 18 July 1984, Naval Air Systems Command, Washington, D.C.,1 September 1981.

Cable Comparison Handbook, MIL-HDBK-299(SH), Department of Defense,3 April 1989.

Circuit Breakers, Selection and Use of, MIL-STD-1498B, Department ofDefense, Washington, D.C., 14 May 1985.

Electrical Connectors, Plug-in Sockets, and Associated Hardware, Selectionand Use of, MI L-STD-1353B, Department of Defense, Washington, D.C.,2 July 1980.

Electrical Insulation, NAVSEA 0960-LP-001-5010, Naval Sea SystemsCommand, Washington, D.C.

Electronics lnstallation and Maintenance Book (EIMB), General, NAVSEASE000-00-EIM-100, Naval Sea Systems Command, Washington, D.C.,1983.

AIII-1

Electronics Installation and Maintenance Book (EIMB), General Maintenance,NAVSEA SE000-00-EIM-160, Naval Sea Systems Command, Washington,D.C., 1981.

Electronics Installation and Maintenance Book (EIMB), Test Methods andPractices, NAVSEA 0967-LP-000-0130, NavalWashington, D.C., 1980.

Fuses, Fuseholders, and Associated Hardware,MIL-STD-1360A, Department of Defense,September 1979.

Sea Systems Command,

Selection and Use of,Washington, D.C., 26

Electronic Assembly Repair, NAVAIR 01-1A-23, change 2, 1 August 1985,Naval Air Systems Command, Washington D.C. 30 April 1982.

Huntron Tracker, 1000 Series Instructional Manual, Huntron InstrumentsInc., 15720 Mill Creek Blvd, Mill Creek, WA. 98012.

Instrument Calibration Procedure, Time Domain Reflectometer, NAVAIR17-20AY-32, Naval Air Systems Command, Washington, D.C., 1 Novem-ber 1980.

Naval Safety Precautions for Forces Afloat, OPNAVINST 5100.19A,Department of the Navy, Office of the Chief of Naval Operations,Washington, D.C., 26 January 1983.

NEETS, Module 3, Introduction to Circuit Protection, Control andMeasurement, NAVEDTRA 14175, Naval Education and Training ProgramDevelopment Center, Pensacola, Florida, 1985.

NEETS, Module 4, Introduction to Electrical Conductors, Wiring Techniques,and Schematic Reading, NAVEDTRA 14176, Naval Education and TrainingProgram Development Center, Pensacola, Florida, 1985.

Navy Occupational Safety and Health (NAVOSH) Program Manual,OPNAVINST 5100-23B, change 4, Department of the Navy, Office ofthe Chief of Naval Operations, Washington, D.C., 31 August 1983.

Standard General Purpose Test Equipment, MIL-STD-1364G (Navy),Department of Defense, Washington, D.C., 1 February 1985.

Tektronix Time Domain Reflectometer Instruction Manual, T.O. 33A1-4-78-1,United States Air Force, May 1980.

Tools and Their Uses, NAVEDTRA 14256, Naval Education and TrainingProgram Development Center, Pensacola, Florida, 1971.

Tracker 2000 Operation and Maintenance Manual, Huntron Instruments Inc.,15720 Mill Creek Blvd, Mill Creek, WA. 98012.

Maintenance Manual, All Levels for Automatic Transistor Analyzer Model900, ST810-AD-OPI-010, Commander, Naval Sea Systems Command,Washington, D.C., July 1985.

AIII-2

Chapter Four

Integrated Weapon Systems Functional Diagrams, NAVAIR 01-F14AAA-2-2-16.1,Change 7, May 1986, Naval Air Systems Command, Washington, D.C.,1 April 1981.

Electrical Power, Power Distribution and Lighting01-F14AAA-2-2-9, Naval Air Systems Command,March 1988.

Systems, NAVAIRWashington, D.C.,

Installation Practices Aircraft Electric and Electronic Wiring, NAVAIR01-1A-505, Naval Air Systems Command, Washington, D.C., 1 December1987.

Chapter Five

Organizational Maintenance Principles of Operation Rotor Systems,A1-H60BB-150-100, 31 March 1987, RAC 1, Commander, Naval AirSystems Command, Washington, D.C., 1 July 1988.

Organizational Maintenance Principles of Operation Power Plants and RelatedSystems, A1-F18AA-270-100, Commander, Naval Air Systems Command,Washington, D.C., 15 June 1981.

AIII-3

Organizational Maintenance Principles of Operation Fuel Systems,A1-F18AC-460-100, CHG 4, 15 June 1988, Commander, Naval Air SystemsCommand, Washington D.C., 1 July 1985.

Chapter Six

Electronics Installation and Maintenance Book (EIMB), General, NAVSEASE000-00-EIM-100, Naval Sea Systems Command, Washington, D.C.,1983.

NEETS, Module 4, Introduction to Electrical Conductors, Wiring Techniquesand Schematic Reading, NAVEDTRA 14176, Naval Education and TrainingProgram Development Center, Pensacola, Florida, 1985.

NATOPS Flight Manual, A-6E/A-6E Trim/KA-6D, NAVAIR 01-85ADF-1,Naval Air Systems Command, Washington, D.C., February 1986.

Chapter Seven

Attitude Heading Reference System, AN/ASN-50, NAVAIR 05-35LAA-1,Naval Air Systems Command, Washington, D.C., January 1984.

NEETS, Module 15, Principles of Synchros, Servos, and Gyros, NAVEDTRA14187, Naval Education and Training Program Development Center,Pensacola, Florida, 1985.

Chapter Eight

Automatic Flight Control Systems AN/ASW-40, NAVAIR 01-85ADF-18,Naval Air Systems Command, Washington, D.C., 1 May 1987.

General Information and Principles of Operation, Volume II, Avionics,NAVAIR 01-23HLH-1-1.2, Naval Air Systems Command, Washington,D.C., October 1983.

NATOPS Flight Manual, A-6E/A-6E Trim/KA-6D, NAVAIR 01-85ADF-1,Naval Air Systems Command, Washington, D.C., February 1986.

NATOPS Flight Manual, Navy Model SH-60B, NAVAIR A1-H60BB-NFM-000Naval Air Systems Command, Washington, D.C., 1 September 1987.

NEETS, Module 15, Principles of Synchros, Servos, and Gyros, NAVEDTRA14187, Naval Education and Training Program Development Center,Pensacola, Florida, 1985.

APPENDIX IV

REVIEW SUBSET ANSWERS

CHAPTER 2

Safety first.

Open and tag the circuit breakers and mainswitches.

Maintenance Instruction Manuals.

Anticollision light.

Energized.

Shorting the terminals together.

Endanger yourself.

A, B, and C.

Use the proper tool for its intendedfunction in the proper manner. Maintainall tools in proper working order and a safecondition.

(1) Analyze the symptom, (2) detect andisolate the trouble, and (3) correct thetrouble and test your work.

The ohmmeter.

Replacement of parts.

MIMs, schematics, and equipment records.

Visual inspection.

Broken wiring, loose terminals or plugconnections, faulty relays and switches.

Installation Practices, Electrical andElectronic Wiring, NAVAIR 01-1A-505.

M I M.

Make sure it works properly.

A9. Carbon dioxide (CO2 ) .

In parallel.

The highest range.

Open circuit.

A continuity test.

Circuit shorting to ground.

(1) The VOM can load the circuit, and(2) the meter movement is easy to damage.

A megger.

An oscilloscope.

An abnormal resistance or impedance thatinterferes with normal signal flow.

Differential voltmeter.

MIL-W-22759.

Aluminum.

15 inches maximum.

Ground.

Selecting the right soldering iron.

Aircraft Structural Hardware for AircraftRepair, NAVAIR 01-1A-8.

Corrosion, strength, size, length, magneticproperties, special features, and lubricationor coating of the substitution part.

To separate the wire bundle from theplumbing line.

(1) Lock wire, (2) shear wire, (3) seal wire.

To ensure that all metal parts of the aircrafthave the same electrical charge.

A1. Aircraft Weapons Systems Cleaning andCorrosion Control, NAVAIR 01-1A-509,and Avionic Cleaning and CorrosionPrevention/Control, NAVAIR 16-1-540.

A2. The capability to perform microminiaturecircuit repair.

A1. Common and peculiar.

A2. The MA-2 aircraft electrical power test set.

A1. A power-off condition.

A2. SND mode.

A3. Open circuit signature.

A4. The test leads are only conductive at thetips.

A5. 270 ohms.

A1. Jet ignition system tester.

A2. Make the ground connection for the powerlead.

A3. 95 to 135 Vac.

TTU-205C/E.

Tank unit probe capacitance and tank unitinsulation resistance.

The angle-of-attack.

A1. Self-check.

A2. Debugging new programs.

A1. 20 volts.

A2. Ground.

A3. 250,000 ohms.

A5. Right wing tip.

A6. Tanker aircraft.

A7. Amber.

A1. Hydraulics.

A2. Control the flow of the fluid.

A3. The pump.

A4. Useful work by linear/reciprocating-mechanical motion.

A5. To ensure the hook is up and locked beforeremoving hydraulic pressure.

A6. Nose wheel steering.

A7. The airstream pressure.

A8. 2,750 PSI.

A1. Ram air temperature.

A2. Crew safety and comfort.

A3. On the inner surface of the outer pane ofglass.

A4. Hot air from combustion heaters or enginecompressors.

A5. Electrical heating elements.

CHAPTER 5

A1. Start, constant speed drive, and air turbinemotor generator.

A2. Overspeed switch mechanism.

A3. Converts energy from compressed air toshaft power.

A4. Removing the positive potential.

A1. Jet engines run on a continuous burningfire.

A2. The ignition exciter.

A3. Between 45 percent and 65 percent of theengine rated speed.

A4. 34 pulses.

A5. A flameout or during aircraft weapons fire.

A1. Feather, ground stop, run, and airstart.

A2. When the power lever is above 66°Fcoordinator and the temp datum switch isin AUTO.

A3. Below 94 percent RPM.

A4. Chromel and Alumel.

A5. A variable potentiometer, discriminatingdevice, and a cam-operated switch.

A1. 65 percent RPM.

A2. 16, 65, and 94 percent switches.

A3. 10 percent.

A4. To accommodate the temperature datumvalve.

A5. Drain valves.

A1. Friction of the air over the aircraft createsenough heat to prevent icing.

A2. Hot bleed air and electrical.

A3. Air at the inlet guide vanes.

CHAPTER 4

A1. To provide specialized exterior lighting andilluminating the interior.

A2. To prevent filament fatigue failure.

A3. Single and double contact bayonet.

A4. Navigation lights.

A4. Deicing removes ice already built upwhereas anti-icing prevents ice buildup.

A1. Short discriminator circuit.

A2. As the temperature increases, the resistanceof the sensing element decreases.

A3. The fire warning will not come on. Thesystem senses a short.

A4. Bromotrifluoromethane (CF3Br).

A1. Fuel tanks 2 and 3.

A2. When the fuel level drops below the highpilot valves in tanks 1 and/or 4.

A3. When fuel tank 1 transfers at a faster ratethan fuel tank 4.

A1. Closed.

A2. The electrical control assembly (ECA).

A1. Flight and ground operation ranges.

A2. Synchrophasing governing mode.

A3. Phase and trim control unit.

A4. Engine RPM.

A5. The limiting circuit (two precent).

A1. Thrust.

A2. Low pitch stop.

A3. Variable hydraulic low pitch stop.

A4. Pitch lock reset.

A5. Feather valve.

A6. Automatic feathering system.

A7. Energize the pressure cutout overridebutton.

A1. Engine power.

A2. The throttles override the system.

A3. A three-position temperature switch.

A1. Transmission pressure switch.

A2. Counterclockwise.

A3. Yes.

A4. To relieve stress built up on the segmentgear.

A5. About 50 open.

A1. Mach number.

A2. Two flight control computers.

CHAPTER 6

A1. By operating principles and the job theyperform.

A2. The weight of the air above that altitude.

A3. 12.23 pounds per square inch.

A4. Airspeed, altimeter, and vertical speedindicator.

Airspeed indicator.

Mach 2.

AGL—Altitude above ground level, andMSL —altitude above mean sea level.

The distance between the aircraft and theterrain it is flying over.

Bimetal yoke.

Total pressure (pitot), indicated staticpressure, indicated angle of attack, andtotal temperature.

The force of air against the aircraft.

80,000 feet of altitude and Mach 2.5.

Angle of attack.

The number of gimbals the gyro has.

Rigidity in space and precession.

Turn-and-bank indicator.

Frequency.

How the permanent magnet aligns to theflux of the two coils.

A junction of two unlike metals used tosense temperature.

Magnets (ring magnet and bar magnet).

The Bourdon tube and the synchro system.

The outer shaft.

A1. White paint.

A2. One.

A3. Brown and orange.

A4. Flammable material, toxic and poisonousmaterials, anesthetics and harmfulmaterials, and physically dangerousmaterials.

CHAPTER 7

A1. Position.

A2. Course.

A3. True north and the direction in which youare pointing.

A4. Longitude.

A5. The prime meridian.

A6. Variation.

A7. East or west, respectively, of true north.

A8. Deviation.

A9. Compass error.

A10. Add.

A11. Dead reckoning.

A12. Inertial navigation systems.

Non characterized and characterized.

Temperature.

The compensator probe.

Integrated position indicator.

It allows the platform to retain the originalorientation regardless of aircraft.

When the displacement gyroscope ismalfunctioning.

Newton’s first law of motion.

Acceleration.

Velocity.

To keep track opposition.

The accelerometer.

Mounting another accelerometer per-pendicular to the first one.

The inner roll gimbal.

84.4 minutes.

Centripetal errors.

Terrestrial, celestial, and inertial.

Earth.

The wander-azimuth system.

Analytic inertial navigation system.

The gyros are not torqued.

Universal screw and the loose magnet type.

When coefficient A amounts to 2° or more. .in either direction.

CHAPTER 8

A1. Ease pilot workload and provide aircraftstability at all speeds.

A2. Any part of an aircraft designed to producelift.

A3. The airfoil stalls.

Rudder.

Angle of attack on the wing.

By increasing the airfoil’s angle of attack(blade pitch).

When the pilot applies collective to themain rotor.

Radar and barometric altimeter signals.

To prevent sudden and violent maneuversupon engagement.

The air navigation computer.

Three axis rate gyro.

A clutched heading signal.

Manual and AFCS engaged.

Surface position transmitter.

Stability augmentation mode.

Stability augmentation system, stabilatorsystem, and digital automatic flight controlsystem.

5-percent control authority.

To stop undesirable noseup attitudes.

When airspeed is below 50 knots.

To improve helicopter stability.

10 percent per second.

0 to 5,000 feet.

120 feet/minute.

REVIEW SUBSET NUMBER 5 A4. The cable angle control panel.

A1. The Doppler radar system and the vertical A5. Speed drift and rate of descent.velocity system.

A2. The crew member at the hoist position. A6. Two, one for pitch and one for roll.

A3. 72 degrees. A7. Pitch, roll, collective, and yaw.

Assignment Questions

Information: The text pages that you are to study are provided at the beginning of the assignment questions.

A S S I G N M E N T 1

Textbook Assignment: “Basic Physics.” Pages 1-1 through 1-15.

*Assignment 1 has been deleted.

Textbook Assignment: “Basic Physics” and “Electrical Maintenance and Troubleshooting.”Pages 1-20 through 2-17.

*Questions 2-1 through 2-39 have been

deleted. This Assignment begins with

Question 2-40. On the answer sheet,

leave questions 2-1 through 2-39 blank.

Learning Objective: Identify safetyprecautions regarding aircraft,personnel, material, and tools.

2-40. Under what category(ies) can allmaintenance performed on navalaircraft be grouped?

1. Preventive only2. Unscheduled only3. Scheduled and unscheduled4. Scheduled and preventive

2-41. An accident-free naval career canbest be achieved by following whichof the following courses of action?

1. Constantly reading technicalmanuals

2. Reading all naval rules andregulations

3. Making a list of all potentialwork hazards

4. Taking a common-sense approachtowards safety

2-42. If electrical equipment is to berepaired, what action should youtake before beginning the actualwork?

1. Remove the fuses for theassociated circuits

2. Short out the main supplyswitches

3. Secure the main power switchesin the open position andproperly tag them

4. Station an individual with afire extinguisher near the workarea

2-43. When an aircraft anticollision lightis operating on the flight line, itwarns personnel of what potentialhazards?

1. Aircraft refueling in progress2. Ordnance loading in progress3. Liquid oxygen converters being

filled4. Aircraft engines are operating

2-44. If you are working on high-voltagecircuits or around wires havingexposed surfaces, you should keeptools and equipment that have metalparts what minimum number of feetfrom the work area?

1. 5 feet2. 2 feet3. 5 feet4. 4 feet

2-45. To prevent low-voltage shock, youshould be careful handling equip-ment having what minimum circuitvoltage?

1. 13 volts2. 12 volts3. 10 volts4. 6 volts

2-46. The intensity of electrical shock isdetermined by which of the followingproperties?

1. Current2. Voltage3. Impedance4. Electromagnetic force

2-47. Which of the following is the primaryreason why a person should not moveabout after receiving an electricalshock?

1. The heart is temporarily weakened2. Muscles have been damaged3. Nerves have been damaged4. The brain is impaired

2-48. When fighting an electrical fire, youshould use which of the followingfire-extinguishing agents?

1. Foam2. Water (H=O)3. Soda and water4. Carbon dioxide (CO=)

2-49. If you swallow gasoline, which of thefollowing actions should you take?

1. Swallow three glasses of saltwater to induce vomiting

2. Drink large amounts of milk orwater and take 4 tablespoons ofvegetable oil, if available

3. Take two aspirins and two glassesof water

4. Swallow a solution of bicarbonateof soda and water

2-50. Which of the following statementsdescribes the hazards of compressedair?

1. It can inject minute foreignbodies into the skin

2. It can rupture cell tissue andcause severe wounds

3. It can pass through clothing andmay cause fatal injury

4. Each of the above

2-51. When using compressed air to cleanout fixtures and jigs, you shouldobserve the proper safetyprecautions. Also, you shouldmaintain the air pressure below whatmaximum value?

1. 10 PSI2. 20 PSI3. 30 PSI4. 40 PSI

2-52. When using tools, you should observewhich of the following rules?

1. Use tools for their intendedpurpose

2. Maintain tools in good repair3. Replace tools that are damaged

or not working properly4. Each of the above

2-53. Which of the following is the causeof most accidents in electrical andelectronic work centers?

1. Moving machinery2. Carelessness3. Improper grounding4. Exposed electrical fixtures

2-54. If one of your tools becomes worn,damaged, or broken, you shouldreport this fact to what person?

1. Crew leader2. Division officer3. Work center supervisor4. Material control officer

2-55. What alley is used to make mostnonmagnetic tools?

1. Cadmium2. Nickel-iron3. Beryllium-copper4. Copper-Constantine

2-56. At the local level, you may insulatetools for use on what type ofcircuit?

1. High-voltage2. Low-voltage3. Low-resistance4. High-resistance

2-57. When you find a damaged power toolelectrical cord, what action shouldyou take?

1. Cover it with rudder tape2. Shorten the cord to remove the

damaged part3. Repair the damage with insulating

tape4. Replace the cord

2-58. Which of the following is NOT asafety practice to follow when usinga soldering iron?

1. Provide ventilation for the ironwhile it is on its rest rack

2. Hold small soldering jobs withpliers or clamps

3. Disconnect the iron duringtemporary absences from the workarea

4. Shake the iron to get rid ofexcess solder

2-59. While using an electric drill, youexperience an electrical shock.Which of the following conditions isthe most likely cause?

1. The voltage source is too high2. The voltage source is too loW3. An incorrectly grounded drill4. An overloaded drill

2-60. What color is the safety ground wirefor electrical tools?

1. Black2. White3. Green4. Red

2-61. You need to apply voltage to a powertool having a three-wire system.The receptacle is a two-wire type.To connect the tool to this voltagesource, you should use an adapterwith an external ground wire andconnect it in which of the followingconfigurations?

1. Tape the exposed ground wireterminal

2. Connect the safety ground to agood ground before plugging inthe tool

3. Connect the safety ground wireof the adapter to the tool case

4. Connect the safety ground wireto the center screw of thereceptacle before plugging inthe tool

2-62. Discrepancies found before, during,or after a flight require what typeof maintenance?

1. Preventive2. Unscheduled3. Scheduled4. Field

2-63. You are troubleshooting anelectrical device that is notreceiving any power. What checkshould you make first?

1. Check fuse or circuit breaker2. Check the power source3. Check for loose connector pins4. Check for visible indications of

trouble

2-64. Which of the following meters shouldyou use when troubleshooting an opencircuit?

1. Voltmeter2. Wattmeter3. Ohmmeter4. Ammeter

After planning the job of correct-ing an electrical power failureon an aircraft, the AE assembledthe necessary tools and equipment.Then the AE performed the follow-ing steps:

A. DISCONNECTED THE AIRCRAFTBATTERY

B. REMOVED A SECTION OF OXYGENLINE TO GAIN ACCESS TO ANELECTRICAL POWER CONNECTOR

C. DISCONNECTED THE CONNECTOR

D. REPLACED A CORRODED PIN INTHE CONNECTOR

E. REASSEMBLED THE CONNECTOR ANDTHE OXYGEN LINE AND RECON-NECTED THE BATTERY

Figure 2A

IN ANSWERING QUESTION 2-65, REFER TOFIGURE 2A.

2-65. What error did the AE commit, andwhy was his/her action incorrect?

1. The battery was disconnected;the main circuit protectorshould have been disconnected

2. The connector was repaired inthe aircraft; it should havebeen repaired in the shop

3. The oxygen line was removed andreplaced; this action shouldhave been done by an AME

4. One pin in the connector wasreplaced? the complete connectorshould have been replaced

2-66. Before replacing a major componentin an aircraft, the AE should makewhich of the followingdeterminations?

1. Whether the component isdefective

2. Whether the intended replacementis a suitable substitute

3. Whether the repair will requirea test flight

4. Whether the appropriate workcenter has been assigned thereplacement task

2-67. Before making an adjustment to anysystem, you should consult which ofthe following publications?

1. MIM2. IPB3. NAVSUP 20024. NATOPS

Learning Objective: Recognize thevarious types of general-purposetest equipment used in aircraftelectrical maintenance, and identifytests the AE makes using theseequipments. (This objective iscontinued in assignment 3.)

An ammeter is connected into acircuit in which of the followingways?

1. In parallel with the circuit2. In series-parallel with the

circuit3. Both 1 and 2 above4. In series with the circuit

When a multimeter is not being used,the selector switch should be inwhich of the following positions?

1. High resistance2. High dc volts3. High ac volts4. Low resistance

Which of the following conditions isthe most probable cause for agrounded circuit?

1. A blown fuse2. A tripped circuit breaker3. Frayed insulation on wiring4. Loose terminal lugs

For precise resistance measurements,you should useinstruments?

1. Wheatstone2. Multimeter3. Ohmmeter4. Megger

which of the following

bridge

What is the value of RX in the dcresistance bridge?

1. 42 ohms2. 400 ohms

3. 420 ohms4. 4,200 ohms

Which of the following types ofmeters are contained in themultimeter?

1. Voltmeter, frequency meter, andohmmeter

2. Frequency meter, ammeter, andvoltmeter

3. Frequency meter, ohmmeter, andammeter

4. Ammeter, voltmeter, and ohmmeter

A permanent-magnet, moving-coilmeter mechanism can be adapted tomeasure alternating current andvoltage if it is used with which ofthe following devices?

1. A transformer2. A transponder3. A rectifier4. A reactor

Which of the following is a standardfeature of the Fluke Model 8100Adigital multimeter?

1. A selectable input filter2. A full four-digit readout3. A power source of 115 volts or

230 volts can be used

Figure 2B

IN ANSWERING QUESTION 2-72, REFER TOFIGURE 2B.

Textbook Assignment: “Electrical Maintenance and Troubleshooting.”Pages 2-17 through 2-74.

Learning Objective (continued):Recognize the various types ofgeneral-purpose test equipment usedin aircraft electrical maintenance,and identify tests the AE will makeusing these equipments.

3-1. To test insulation for highresistance, grounds, and leakage,what meter should you use?

1. VTVM2 . Megger3. Ohmmeter4. Multimeter

3-2. A megger is prevented from exceedingits rated output voltage by theaction of a

1. friction clutch2. voltage regulator3. diode limiter4. variable resistor

3-3. Which of the following values can bemeasured by using an oscilloscope?

1. Frequency2. Voltage amplitude3. Phase differences4. Each of the above

3-4. What term is used to define abnormalresistance or impedance thatinterferes with the normal signalflow?

1. Discontinuity2. Distortion3. Reflectometry4. Reduction

3-5. What instrument should you use totroubleshoot fuel quantity coaxialcables?

1. Time-domain reflectometer2. Whetstone bridge3. Ammeter4. Phase-angle voltmeter

3-6. The output of the phase detector ina phase-angle voltmeter isproportional to the signal amplitudemultiplied by which of the followingangles of phase difference?

1. Sine2. Cosine3. Tangent4. Cotangent

3-7. Maximum deflection of the phase-angle voltmeter occurs when whatphase relationships exists betweenthe two signals?

1. 0° or 90°2. 0° or 180°3. 90° or 120°4. 90° or 180°

3-8. What type of voltmeter is a precisionvoltmeter that compares an unknownvoltage with an internal referencevoltage?

1. Phase-angle voltmeter2. Time-domain reflectometer3. Differential voltmeter4. Fluke Model 8100A

Learning Objective: Recognizeaircraft wire and cablecharacteristics and various means ofidentifying and splicing wire andcables.

3-9. To be classified as a cable, a singleconductor must have which of thefollowing characteristics?

1. Be insulated and designed tocarry RF energy

2. Be insulated and have a metallicbraided shield

3. Be covered by a metal shield anddesigned to carry RF energy

4. Be covered by a metal shield andhave at least a 00 wire size

3-10. To replace an aircraft electricalwire, you must determine the correctsize and type of wire to use. Tomake this determination, whatpublication should you consult first?

1. The aircraft MIM2. The aircraft IPB3. Military Specifications, MIL-W-

5088 (latest edition)4. Installation Practices for

Aircraft Electric and ElectronicHiring, NA 01-1A-505

3-11. Aluminum has the tendency to flowaway from a point where pressure isapplied. This tendency is known as

1. flowing2. crystallization3. creep4. feed through

3-12. When stamping wires or cables, thedistance between markings should notexceed what maximum distance?

1. 24 inches2. 15 inches3. 3 inches4. 6 inches

3-13. You are reading a wireidentification code. Which of thefollowing types of information canyOu gain?

1. Circuit function2. Wire size3. Wire segment4. Each of the above

3-14. Which of the following letters isNOT used to identify a wire segment?

3-15. What letter suffix in the wireidentification identifies the wireas being a ground?

1. E2. O3. X4. Z

1. A2. B3. C4. N

Heat-shrinkable tubing has which ofthe following advantages?

1. It insulates wire terminals2. It waterproofs wire splices3. Both 1 and 2 above4. It replaces wire terminal covers

When used on aircraft electricalwiring, the recommended power ratingrange for general-use soldering ironsis within which of the followingranges?

1. 13 to 60 watts2. 20 to 500 watts

3. 55 to 600 watts4. 60 to 200 watts

Learning Objective: Recognize theuses for and characteristics ofaircraft electrical and mechanicalhardware.

Before reusing items of mountinghardware, what determination shouldyou first make?

1. Whether they exceed thespecifications for their intendeduse

2. That they are the same size andshape of the specified items

3. If their reuse is prohibited byexisting directives

4. That they are not damaged andexceed the specification for theitems required by the IPB

You have temporarily installedsuitable substitute mounting parts.When should these parts be replaced?

1. During the next periodicinspection

2. During the time the aircraft isat NADEP

3. When the substitute parts becomedefective

4. When the required parts becomeavailable

Which of the following consider-ations should you observe whensubstituting mounting parts?

1. Color2. Availability3. Magnetic properties4. Accessibility

Refer to figure 2-15 in your text.Which of the following statements iscorrect concerning the replacementof shock mounts?

1. Those shown in both view A andview B can be replaced, but allfour must be replacedsimultaneously

2. Those shown in view B can beindividually replaced, and thosein view A cannot be replaced

3. Those shown in view A can beindividually replaced, and thosein view B cannot be replaced

Refer to figure 2-15, view B, inyour text. One vibration insulatoris cracked. What corrective actionshould you take?

1. Replace the complete shock mountassembly

2. Replace the cracked insulator3. Replace all the vibration

insulators

3-23. What is the reason for moisture-proofing solder-type electricalconnectors?

1. To reinforce the connector2. To reduce the possibility of the

connector cracking3. To improve the connector’s

dielectric characteristics4. To protect the connector for

future use

3-24. The preferred method for attachingcable terminals to terminal blocksrequires the use of what items ofhardware?

1. An anchor nut and lock washer2. An anchor nut and flat washer3. A standard nut and lock washer4. A standard nut and washer

3-25. Shielded conduit should be sup-ported by the use of what type ofclamp?

1. AN 7422. Strap3. Bonded4. Nonbonded

3-26. When long runs of cable betweenpanels need to be supported, which ofthe following types of clamps ispreferred?

1. Strap2. Plastic3. Bonded4. AN 742

When installing a cable through alightening hole, you should use agrommet (rubber cushion) if thecables’s distance from the edge ofthe hole is less than what minimumdistance?

1. 1/4 inch2. 3/8 inch

3. 1/2 inch4. 5/8 inch

In addition to supporting andprotecting electrical wires, whatother advantage does conduit offer?

1. It protects against heatradiation

2. It provides radio shielding3. It provides bullet deflection4. It supports adjacent cables

What type of safety wire should youuse to secure an oxygen regulator?

1. Lockwire2. Seal wire3. Shear wire4. Soft steel wire

If an aircraft were improperlybonded, which of the followingconditions would exist?

1. An increased likelihood of fireand a noisy radio receiver

2. An increased likelihood oflightening strikes

3. A decreased chance of lighteningstrikes

4. The likelihood of electricalfailures increases

3-31. A primary objective of bonding is toprovide an electrical path of

1. high dc resistance and high RFimpedance

2. low dc resistance and high RFimpedance

3. high dc resistance and low RFimpedance

4. low dc resistance and low RFimpedance

Learning Objective: Recognizeelectrical motor and generatormaintenance procedures.

3-32. When using methyl chloroform to cleanelectrical equipment, you shouldremove the equipment from thesolution within what maximum lengthof time?

1. 5 minutes2. 5 to 15 minutes3. 15 to 30 minutes4. 30 to 60 minutes

3-33. A generator is not delivering itsrated voltage or current. Before youremove this generator, you shouldcheck which of the followingcircuits?

1. Control2. Feeder3. Regulator4. Each of the above

3-34. When a printed circuit ismanufactured by the photoetchingprocess, what portions of theplastic or phenolic sheet areactually photographically exposed?

1. All areas covered by light-sensitive enamel not covered bythe circuit template

2. Only areas covered by thecircuit template

3. Only areas where circuitcomponents, such as resistors,are attached

4. Only areas that act as wires

3-35. Exposed copper is removed during theetching process, and the unexposedcopper surfaces are protected by

1. enamel2. solder3. the printed circuitry overlay4. the exposed copper smear

Learning Objective: Identifyfunctions, capabilities, andoperating characteristics of varioustypes of aircraft test equipment.

3-36. What color paint is normally usedfor line test equipment?

1. Yellow2. Orange3. White4. Red

Learning Objective: Identify thecharacteristics of printed circuits,including modules and pottedcomponents, and recognize circuitconstruction features.

3-37. Which of the following statementsdescribes actions you should takewhen troubleshooting aircraftelectrical equipment?

1. Use line test equipment inconjunction with shop testequipment

2. Follow the instructions in theapplicable MIM

3. Use test equipment to isolatemalfunctions to a specificcomponent

3-38. The MA-2 Aircraft Electrical PowerTest Set (stand) is NOT used toperform a complete test and check ofwhich of the following components?

1. Generator outputsVoltage regulator circuits

3. Reverse current relay operation4. Generator spline wear

3-39. The load bank of the MA-2 is capableof supplying dc loads from 0 to whatmaximum value?

1. 500 amperes2. 750 amperes

3. 1,000 amperes4. 1,500 amperes

3-40. When using the Huntron Tracker 1000,which of the following statements isa test requirement?

1. All circuit power must be off2. The signal fuse must be installed3. The test leads attached to the

device under test4. Each of the above

3-41. The Huntron Tracker 2000 quadrant 3shows which of the followingdisplays?

1. Positive voltage and negativecurrent

2. Negative current and negativevoltage

3. Positive current and negativevoltage

4. Positive current and positivevoltage

3-42. When using the Huntron Tracker 1000or 2000, you should begin testing inwhich of the following ranges?

1. Low2. High3. Medium

3-43. When testing analog circuits ordevices with a Huntron Tracker, youshould use the low range of thetester for which of the followingreasons?

Defects will show easier, andthe internal impedance makes itmore likely the device undertest will load the testerDefects are easier to find inthe medium rangeThe internal impedance willcause the device under test to

load the testerDefects will show easier , andthe internal 54 ohm impedancemakes it less likely that partsin parallel with the deviceunder test will load the tester

3-44. In the SND mode of the AutomaticTransistor Analyzer Model 900, a gooddevice is indicated by the Sonalertwhen it beeps in what way?

1. Out of phase with the amber light2. In phase with the amber light3. In phase with the green light4. Out of phase with the red light

3-45. When using the Automatic TransistorAnalyzer Model 900 to test atransistor, you can connect thetransistor to the tester in (a) whattotal number of ways of which (b)what total number are correct?

1. (a) Six (b) one2. (a) Four (b) two3. (a) Four (b) one4. (a) Six (b) two

3-46. You should use the jet ignitionsystem tester to check theinput/output of which of thefollowing components?

1. Spark plug2. Power output of the ignition unit3. Power input to the ignition unit4. Each of the above

ITEMS 3-47 AND 3-48 PERTAIN TO THE TTU-27ETACHOMETER INDICATOR TEST SET.

3-47. The test set can be used to testwhich of the following components?

1. Two-pole and four-pole tachometergenerators for speed and outputvoltage under load

2. Tachometer indicator calibrationaccuracy

3. Tachometer indicator voltage4. Each of the above

3-48. After the tachometer generatorto be tested is. mounted on thetester, what component will controlthe speed at which it will bedriven?

1. The master tachometer generator2. A variable dc drive motor in the

test set3. The two-speed test pad4. The gearbox drive

ITEMS 3-49 AND 3-50 PERTAIN TO THE JETCALIBRATION (JETCAL) ANALYZER.

3-49. When performing a functional groundtest of the EGT system, heat for thethermocouples is provided by

1. heater probes2. exhaust gas3. the JETCAL potentiometer4. the aircraft heating and air-

conditioning system

3-50. An external ac power supply isrequired to supply electrical powerto a JETCAL analyzer that is beingused to make which of the followingchecks?

1. Engine speed2. The EGT circuit for shorts and

grounds3. EGT indicators4. The resistance of the EGT

circuit

3-51. Which of the following pulses is/aregenerated by the synchrophaser testset?

1. Slave pulse only2. Master pulse only3. Slave and master pulses4. Tachometer pulse

The gain test readout from thesynchrophaser test set is provided bywhat front panel component?

1. The null meter2. The galvanometers3. The calibrated potentiometer4. The feedback potentiometer

Which of the following test setsprovides regulated pitot and staticpressure for evaluating theperformance characteristics of airdata systems, aircraft pneumaticinstruments, and other auxiliaryequipment?

1. TF-202. TTU-205C/E3. TTU-27E4. AN/PSM-17A

Of the following lists, which onegives the power requirements for theTTU-205C/E test set?

1. 28 Vdc, three-phase, 360 hertz2. 28 Vdc, 115 Vat, single phase,

400 hertz3. 115 Vac, single phase, 400 hertz4. 115 Vac, three-phase, 400 hertz

The TF-20 liquid quantity test setcan be used to perform which of thefollowing functions’?

1. Measure tank unit probecapacitance and insulationresistance

2. Simulate the total capacitance ofa probe for checking the aircraftfuel quantity indicator

3. Calibrate the test set’smegohmmeter scales andcapacitance indicator dial

When using the AN/FSM-17A todynamically test the transmitter ofthe angle-of-attack system, whattype of energy is fed tothe transmitter probe?

1. Hydraulic or pneumatic2. Vacuum or air pressure

Mechanical3.

4. Electrical

The AN/PSM-21A air-conditioning testset is used to troubleshoot andcheck which of the followingsystems?

1. Cabin pressure2. Pressure suit3. Equipment air conditioning4. Each of the above

The basic configuration of a typicalVAST station consists of which ofthe following components?

1. A computer subsystem, a DTU, anda stimulus and measurementsection

2. A computer processing unit, aDTU, and a stimulus andmeasurement section

3. A computer subsystem, a CPU, anda stimulus and measurementsection

4. A CPU, a DTU, and an outputmonitor

What component of the VAST systemserves to synchronize instructionsbetween the computer and the VASTsystem’s functional building blocks?

1. Stimulus and measurement section2. MTTU3. CPU

4. Each of the above 4. DTU

3-60. What are the three levels of faultdetection used in the VAST?

1. Primary, secondary, and buildingblocks

2. Basic block, core block, andperipheral block

3. Self-test, auto-check, and self-check

4. Manual, semiautomatic, andautomatic

3-61. What test should you run on a VASTstation before you apply power to aUUT?

1. Self-test2. Auto-test3. Continuity test4. Dynamic functional test

3-62. Programmed halts are used during VASTtesting to permit the operator tomanually intervene in order toperform which of the followingprocedures?

1. Self-tests2. Observations only3. Adjustments only4. Observations and adjustments

Learning Objective: Recognize thehazards to ESD-sensitive devices, toinclude proper handling and packagingtechniques.

3-63. What is the lowest voltage that willdestroy or damage an ESD-sensitivedevice?

1. 10 volts2. 20 volts3. 30 volts4. 25 volts

3-64. The generation of static electricityon an object by rubbing is known asthe

1. electrostatic charge2. dielectric effect3. triboelectric effect4. prime charge

3-65. Under which of the followingconditions is generated staticelectricity decreased?

1. Dry air2. Humid air3. Cold air4. Hot air

3-66. Which of the following is NOT an ESDprime generator?

1. Synthetic mats2. V i n y l3. Common solder suckers4. Carbon impregnated polyethylene

3-67. What is the minimum resistance forpersonnel ground straps?

1. 25,000 ohms2. 150,000 ohms3. 250,000 ohms4. 500,000 ohms

3-68. Which of the following proceduresshould NOT be performed when workingon ESD-sensitive devices?

1. Ground the work area, wriststraps, and equipment

2. After testing, replace shortingdevices and protective packaging

3. Perform dielectric strengthtests

4. Use of a Simpson 260 meter orequivalent to test components

Textbook Assignment: “Power Generation and Control Systems” and “Aircraft ElectricalSystems.” Pages 3-1 through 4-2.

*Assignment 4 has been deleted.

Textbook Assignment: “Aircraft Electrical Systems” and “Aircraft Power Plant Systems.”Pages 4-4 through 5-6.

Learning Objective: Identify thetypes, purposes, and uses of aircraftexternal lighting.

5-1. The minimum requirements fornavigation lights on all militaryheavier-than-air aircraft areestablished by what agency?

1. Department of Defense2. Department of Transportation3. Federal Transportation

Administration4. Federal Aviation Administration

5-2. Flight safety is the primary purposeof which of the following lights?

1. Position lights2. Formation lights3. Fuselage lights4. Anticollision lights

5-3. Refer to figure 4-6 in your text.The actuation of what device releasesthe magnetic brake allowing thelanding light to extend or retract?

1. Mechanical linkage actuation only2. Retract-extend switch actuation

only3. Mechanical linkage and retract-

extend switch actuation4. Gear train actuation

5-4. What type of light provides thelanding safety officer with a visualindication of a carrier aircraft’ssafe or unsafe landingconfiguration?

1. Index lights2. Position lights3. Formation lights4. Approach lights

IN ANSWERING (QUESTIONS 5-5 THROUGH 5-7,MATCH THE ANGLE-OF-ATTACK INDICATION INCOLUMN A WITH THE CORRECT COLOR APPROACHLIGHT IN COLUMN B. ALL ANSWERS ARE NOTUSED.

A. ANGLE-OF-ATTACK B. APPROACHINDICATION LIGHT

5-5. Angle of attack is 1. Ambertoo low

2. Green5-6. Angle of attack is

too high 3. Red

5-7. Angle of attack is 4. Whiteoptimum

5-8. What lights provide the pilot withangle-of-attack information?

1. Approach lights2. Indexer lights3. Instrument lights4. Position lights

With reference to the indexer lights,an inverted V indicates to the pilotthat the angle of attack is in whichof the following positions?

1. Slightly low2. Slightly high3. Very high4. Very low

What is the landing configuration ofan aircraft if the landing signalofficer (LSO) observes flashingapproach lights?

1. The landing gear is up only2. The landing and arresting gears

are up3. The arresting gear is not fully

extended4. Normal operation

What is the function of the arrestinggear override switch?

1. It activates a light showing theLSO that the arresting hook isextended for landing

2. It allows the approach lights tosignal that the aircraft isunprepared to land

3. It allows the approach lights tofunction properly while thearresting hook is up

4. It shows the pilot the positionof the arresting hook

The fuselage formation lights areconnected in parallel with andcontrolled by the same switches aswhat other lights?

1. The fuselage signal lights only2. The wingtip formation lights only3. The fuselage signal and wingtip

formation lights only4. The fuselage signal, wingtip

formation, and navigation lights

5-13. What is the purpose of the in-flightrefueling probe light?

1. To illuminate the drogue of therefueling aircraft only

2. To illuminate the probe of theaircraft being refueled only

3. To illuminate the probe of thereceiver aircraft and the drogueof the refueling aircraft

4. To indicate fuel flow andstoppage of fuel flow into theaircraft being refueled

5-14. When a controllable light is mountedin the nose of a helicopter, it haswhat total number of degrees ofazimuth travel?

1. 90°

2. 180°3. 270°4. 360°

Learning Objective: Identify thetypes, purposes, and uses ofaircraft internal lighting.

5-15. When replacing aircraft interiorlamps or light covers, you shouldmake sure that replacements meet thesame specifications as theoriginally installed units for whichof the following reasons?

1. Fire hazards can be easilycreated in lighting fixturealterations

2. The aircraft’s power circuitrywill be unbalanced if substitutelighting is used

3. Original specifications werebased on scientific consider-ations of necessity and crewcomfort

5-16. What are the advantages of the grain-of-wheat instrument lamps over othertypes of lamps used in instrumentsystems?

1. Longer life only2. More rugged only3. Better illumination only4. Longer life, more rugged, and

better illumination

5-17. What lighting feature is provided toaid a crew member who is reading achart?

1. Red floodlights2. White floodlights3. Extension lights4. Momentary contact switches to

bypass the rheostats onfloodlights

5-18. Which of the following statementsdescribes the push-to-test feature onwarning lights?

It provides a means for checkingthe condition of the warninglight bulb onlyIt provides a means for checkingthe system’s circuits onlyIt provides a means for checkingthe system’s circuits and thecondition of the warning lightbulbIt provides a means formomentarily activating allequipments and circuits in therespective systems

IN ANSWERING QUESTIONS 3-19 THROUGH 5-22,REFER TO FIGURE 4-11 IN YOUR TEXT, ANDMATCH THE COLOR OF THE LEGEND-TYPE LIGHT INCOLUMN B TO THE FUNCTION IT REPRESENTS INCOLUMN A. EACH ANSWER IS USED AT LEASTONCE.

A. FUNCTION

It indicates

B. COLOR

there 1. Yellowis a requirementimmediate action

A malfunction isindicated

A safe or normalindicated

The condition or

for2. Red

3. Green

configuration is

performance ofequipment is indicated

Learning Objective: Recognize theoperating principles andcharacteristics of aircraftelectrohydraulic and pneumaticsystems.

5-23. In aircraft hydraulic systems, theAE maintains circuits that controlthe fluid

1. viscosity2. flow3. shape4. pressure

5-24. All hydraulic systems contain aminimum of which of the followingbasic components?

1. Pump, selector valve, actuator,and reservoir

2 . Pump, pressure regulator, switch,and reservoir

3. Selector valve, filter, pump, andactuator

4. Selector valve, actuator,pressure lines, and return lines

5-25. Which of the following componentsdirects the fluid flow in a hydraulicsystem?

1. Reservoir2. Pump3. Selector valve4. Actuating unit

IN ANSWERING QUESTIONS 5-26 THROUGH 5-29,REFER TO FIGURES 4-13 AND 4-14 IN YOUR TEXT.

Hydraulic pressure is supplied toAFCS components by which of thefollowing components of the hydraulicsurface control booster system?

1. Modulator piston transducer2. Main control valve3. Engagement valve4. Transfer valve

When the hydraulic transfer valve isin the static state and the AFCS isengaged, which of the followingforces hold(s) the plunger mechanismin the center position?

1. Equal hydraulic pressure on bothsides of the modulator piston

2. Equal current flow in thetransfer valve coils

3. Equal tension on both centersprings

4. Both 2 and 3 above

5-28. The hydraulic surface controlbooster system operates in which ofthe following modes?

1. One--hydraulic2. Two--manual and AFCS3. Three--hydraulic, manual, and

pneumatic4. Each of the above

5-29. In the AFCS, what is the purpose ofthe modulator piston lineartransducer?

1. To send rate signals to the AFCScomputer

2. To send position signals to theAFCS computer

3. To reposition the modulatorpiston

4. To detect hydraulic failures

5-30. Which of the following functions iscommon to both the left and rightmain gear torque-link switches?

1. Preventing the throttles frombeing placed in the reversepropeller range while airborne

2. Furnishing power for the landinggear control lever lockingsolenoid

3. Disabling wing station externalstores circuits

4. Energizing the bomb bay doorcontrol circuit

5-31. Refer to figure 4-15 in your text.What component prevents the landinggear control lever from being placedin the wheel-up position when theweight of the aircraft is on thelanding gear?

1. A manually operated solenoid2. An electrically operated but de-

energized solenoid3. The safety switch in the right

main gear torque-link switchhousing

4. An electrically operated solenoidthat energizes when the aircraftlands.

IN ANSWERING QUESTIONS 5-32 AND 5-33, REFERTO FIGURE 4-16 IN YOUR TEXT.

5-32. Retraction of the arresting hook iselectrically controlled andhydraulically actuated; however,extension of the hook is accomplishedby the use of what means?

1. Hydraulic power only2. Electrical power only3. Electrical or hydraulic power4. Mechanical

5-33. What is the purpose of the time-delay relay in the relay panel?

To dampen hydraulic pressuresurges when the system is firstengagedTo ensure that the arresting hookis fully up and locked beforehydraulic pressure is removedTo prevent the arresting hookfrom dropping if the handle isinadvertently moved to the downpositionTo ensure that the arresting hookis completely down beforehydraulic pressure is applied

IN ANSWERING QUESTIONS 5-34 THROUGH 5-37,SELECT FROM COLUMN B THE COMPONENT THATPERFORMS EACH FUNCTION OF THE STEER-DAMPERUNIT LISTED IN COLUMN A. EACH ANSWER ISUSED ONCE.

A. FUNCTION B. COMPONENT

Controls 1. Steeringhydraulic shutoffpressure to valvethe steer-damper unit 2. Servo valve

Permits 3. Unidirectionalrestricted restrictorshydraulicfluid flow 4. Fluidto the bypass compensatorvalve to dampennosewheel shimmy

Controls the flow of fluid to theactuator

Prevents system pressure frombecoming excessive in the steer-damper unit

IN ANSWERING QUESTIONS 5-38 THROUGH 5-41,SELECT THE COMPONENT FROM COLUMN B THATPERFORMS THE FUNCTION LISTED IN COLUMN A.(THESE QUESTIONS CONCERN THE ELECTRICALUNITS IN THE STEERING SYSTEM.) EACH ANSWERIS USED AT LEAST ONCE.

Contains a 1. Command

swivel disconnect potentiometer

switch to preventreverse steering 2. Steering

amplifierFeeds a signalto the steering 3. Steeringamplifier as the feedback

nosewheel is moved potentiometer

Provides a nonlinear steeringresponse as it is varied mechanicallyby the rudder pedals

Receives the turn signal ordered bythe rudder pedals from the commandpotentiometer

5-42. The catapulting system’s launch barwarning light illuminates when whichof the following conditions exist?

1. The launch bar is up and lockedwith weight off the landing gear

2. The launch bar is up and lockedwith weight on the landing gear

3. Solenoid B is energized4. Solenoid A is energized

5-43. The launch bar control switch isplaced to retract, and the warninglight remains on. What componentfailure is indicated?

1. Valve position switch2. Weight on gear switch3. Selector valve4. Control switch

5-44. Which of the following is adescription of a speed brake controlswitch?

Speed brake extension will stopwhen the switch is released fromthe OUT positionSpeed brake retraction will stopwhen the switch is momentarilyheld in the IN positionThe speed brake moves fully tothe indicated position when theswitch is momentarily held ineither the IN or OUT positionThe degree of extension orretraction is controlled by thedegree of movement of the switchlever to or from the STOPposition

5-45. Refer to figure 4-19 in your text.To retract the speed brakes, thespeed brake control switch is movedto the IN position, causing solenoidA and solenoid B to be in which ofthe following states?

1. Energized2. De-energized3. Energized and de-energized,

respectively4. De-energized and energized,

respectively

QUESTIONS 5-46 THROUGH 5-48 PERTAIN TO THECANOPY SYSTEMS OF AN A-6 AIRCRAFT.

5-46. Which of the following statementsbest describes the operationalprinciple of the canopy selectorvalve?

1. It is hydraulically orelectrically actuated andmanually operated

2. It is manually or hydraulicallyactuated and manually operated

3. It is hydraulically or manuallyactuated and electricallyoperated

4. It is manually or electricallyactuated and hydraulicallyoperated

5-47. To close the canopy, which of thefollowing conditions must exist?

1. The control circuit breakerclosed, canopy switch closed, andsolenoid 1 energized

2. The control circuit breakerclosed, canopy switch open, andsolenoid 2 de-energized

3. The control circuit breakerclosed, the canopy switch closed,and solenoid 2 energized

4. The control circuit breaker open,the canopy switch closed, andsolenoid 1 energized

5-48. To open the canopy, which of thefollowing conditions must exist?

The control circuit breakerclosed, the canopy switch open,and solenoid 1 energizedThe control circuit breaker open,the canopy switch open, andsolenoid 1 energizedThe control circuit breakerclosed, the canopy switch closed,and solenoid 2 energizedThe control circuit breaker open,the canopy switch open, andsolenoid 1 de-energized

IN ANSWERING QUESTIONS 5-49 THROUGH 5-52,SELECT FROM COLUMN B THE PNEUMATIC POWERSYSTEM COMPONENT THAT PERFORMS EACHFUNCTION LISTED IN COLUMN A. EACH ANSWERIS USED ONCE.

Ensures the fluid 1. Pressurein the hydraulic sensingmotor is not switchsurging

2. FlowSimultaneously regulatorpasses voltage tothe hydraulic 3. Dumpselector valve valveand dump valve

4. CheckProtects the valvemotor againstreverse fluid pressure

Enables accumulated moisture toescape from the air system

Learning Objective: Recognizeterms, definitions, components, andoperating principles and features ofaircraft environmental systems,including those used to controltemperature, pressure, and icing.

5-53. All of the following conditionscreate a demand for cabin airconditioning in aircraft flying atextreme airspeeds. Which one is theprincipal cause of cabintemperatures rising above the levelat which the crew can maintain topphysical and mental efficiency?

1. Engine heat2. Solar heat3. Body temperature4. Ram-air friction

Temperature that is measured from apoint at which there is no molecularmotion is known as the

1. standard temperature2. absolute temperature3. critical temperature4. ambient temperature

Which of the following is a correcttemperature equivalent?

1. 100°C = 212°F2. 32°C = 212°F3. 32°F = 100°C4. 0°F = 32°C

The cabin air-conditioning andpressurization system maintains thecabin air temperature and pressure ata comfortable and safe level byforcing which of the following kindsof air through cockpit diffusers?

1. Dehumidified refrigerated air2. Hot engine bleed air3. Both 1 and 2 above4. Ambient refrigerated air and

humidified hot engine bleed air

In an aircraft, cabin air pressure iscontrolled by the operation of whichof the following components?

1. A safety valve and a manual dumpcontrol

2. A pressure regulator and a safetyvalve

3. A manual dump control and apressure regulator

4. All of the above

5-58. With the cockpit switch in the ONposition, the desired cabintemperature is maintained by whichof the following means?

1. Variation in the opening of theram-air valve

2. Proportional amounts of enginebleed air and refrigerated airbeing mixed by the dualtemperature control valve

3. Proportional amounts of ram airand hot engine bleed air beingmixed by the ram-air valve

4. Proportional amounts of ram airand refrigerated air being mixedby the ram-air valve

5-59. When the cockpit switch is in theRAM AIR position, what are theconditions of the various valves?

The dual temperature controlvalve is in the full hotposition, the cabin bleed-airvalve is closed, and the cabinram-air valve is selected by theMAN/RAM AIR switchThe dual temperature controlvalve is selected by the MAN/RAMAIR switch, and the cabin ram-air valve is closedBoth the dual temperaturecontrol valves are closed, andthe cabin ram-air valve is asselected by the MAN/RAM AIRswitchBoth the dual temperaturecontrol valves are open, and thecabin ram-air valve is selectedby the MAN/RAM AIR switch

5-60. Cabin temperature changes areanticipated by what component(s)?

1. Cabin temperature sensor only2. Cabin temperature and duct

sensors3. Cabin duct dual temperature

sensor4. Temperature control wheel

5-61. At what altitude is cabinpressurization automaticallyinitiated?

1. 6,000 feet2. 8,000 feet3. 10,000 feet4. 12,000 feet

A. CABIN PRESSURE IS MAINTAINEDAT SEA-LEVEL PRESSURE AT ALLALTITUDES.

B. UP TO 8,000 FEET, CABINPRESSURE IS HELD AT THE8,000-FOOT ALTITUDE, ANDABOVE 8,000 FEET, A 5-PSIDIFFERENTIAL IS MAINTAINEDBETWEEN CABIN AND AMBIENTPRESSURES.

C. UP TO 8,000 FEET, THE CABINPRESSURE IS MAINTAINED ATSEA-LEVEL PRESSURE, AND ABOVE8,000 FEET, CABIN PRESSURE IS

HELD AT THE 9,000-FOOTPRESSURE LEVEL.

D. UP TO 8,000 FEET, THE CABINPRESSURE IS THE SAME ASAMBIENT. ABOVE 8,000 FEET,THE CABIN PRESSURE REMAINS ATTHE 8,000-FOOT PRESSURE LEVELUNTIL A DIFFERENTIAL OF 5-PSIEXISTS BETWEEN CABIN ANDAMBIENT PRESSURES. BEYONDTHIS ALTITUDE THE PRESSUREDIFFERENTIAL REMAINS AT 5-PSI.

Which of the following is thepressure schedule maintained by thecabin pressure regulator?

1. A2. B3. C4. D

Under what set of conditions willthe right forward and aft equipmentcompartments receive moist, cooledbleed air?

When the equipment cooling valveiS open, the ram-air valves areopen, and the temperature isbelow 46°CWhen the equipment cooling valveis open, the ram-air valves areclosed, and the temperature iSabove 46°CWhen the equipment cooling valveis open, the ram-air valves areopen, and the temperature isabove 46°CWhen the equipment coolIng valveis closed, the ram-air valvesare open, and the temperature isabove 46°C

What is the function of the cabinduct limit bridge?

1. To limit the temperature of thecabin inlet air

2. To anticipate sudden cabintemperature changes

3. To select the desired cabintemperature

4. To override the cabin ductanticipator bridge

Figure 5A

IN ANSWERING QUESTION 5-62, REFER TO FIGURE5A.

What is the purpose of the voltagefrom the feedback potentiometer inthe cabin dual temperature controlvalve?

To reduce the starting voltageonce the valve actuator motor hasstarted rotatingTo ensure the error feedbacksignal to the modulator circuitis the correct phaseTo prevent oscillations of thevalve actuator motorTo cause the valve actuator motorspeed to increase, giving morepositive control to thetemperature regulation

Windshield overheating is preventedby the combined actions of a shutoffvalve and what other component?

1. A thermostat2. A dump valve3. A thermistor4. An electronic temperature

controller

What is the location of the heatingelement for windshield anti-icing anddefogging?

1. On the outer surfaces of bothglasses

2. On the outer surface of the outerglass

3. Between the outer glass and thevinyl plastic core

4. Between the inner glass and thevinyl plastic core

5-68. What method is used to deice the

empennage of P-3 aircraft?

1. A de-powered, at-controlledsystem providing constant heatto the empennage

2. Hot engine bleed air iscirculated under the surfaces

3. The leading edges of thevertical and horizontalstabilizers are electricallyheated

4. Heat from the cabin routedthrough control valves heats theempennage

Learning Objective: Recognize theoperating principles andcharacteristics of aircraft startingand ignition systems.

Jet engine starters must have whichof the following capabilities?

1. Low starting torque and lowspeed

2. High starting torque and highspeed

3. Low starting torque and highspeed

4. High starting torque and lowspeed

What two sections make up theconstant-speed drive/starter (CSD/S)unit?

1. A turbine motor and a planetarydifferential transmission

2. A turbine motor and a generator3. A generator and a planetary

differential transmission4. A generator/turbine motor and a

planetary differentialtransmission

5-71. The CSD/S is maintaining an output of8,000 RPM from the engine input driveshaft. In what mode is the CSD/Soperating?

1. The air turbine motor generatordrive mode

2. The shaft mode only3. The start and air turbine motor

generator drive modes4. The constant-speed drive mode

5-72. Refer to figure 5-3 in your text.The holding relay receives itspositive potential after the startswitch is momentarily pressed throughthe

1. multiengined starting systemselector switch

2. normally closed stop switch3. pressure cutout switch4. starter overspeed switch

5-73. For a fire to occur, what elementsmust be present?

1. Oxygen and heat only2. Heat and a combustible material

only3. Oxygen and a combustible material

only4. Heat, oxygen, and a combustible

material

5-74. In an electronic ignition system,what component develops the voltagethat produces a spark?

1. The exciter2. The dynamotor3. The transformer4. The booster coil

5-75. In an electronic ignition system,ignition is discontinued when whatpercentage of the rated engine speedis reached?

1. Between 15 and 20 percent2. Between 25 and 45 percent3. Between 45 and 65 percent4. Between 65 and 75 percent

Textbook Assignment: “Aircraft Power Plant Systems.” Pages 5-9 through 5-50.

6-4.Learning Objective: Recognizeoperating principles andcharacteristics of aircraft enginetemperature control and engine startcontrol systems.

6-1. In turbine-powered aircraft, whatrelationship, if any, exists betweenengine power and turbine temperature?

1. They are directly proportional2. They are inversely proportional3. They tend to cancel each other 6-5.4. None

6-2. By what means does an enginetemperature control system onturboprop engines control enginetorque?

1. A converging and diverging panelcircuit

2. Electronic fuel trimming3. A restrictive airflow fan circuit4. Variable airflow

6-3. The fuel shutoff valve is elec-trically closed on engine shutdown byplacing the condition lever in whatposition?

The engine coordinators function tocoordinate the power and conditionlevers along with what othercomponent(s)?

1. The fuel control only2. The propeller and fuel control

only3. The electronic fuel trimming

circuit only4. The fuel control, the propeller,

and the electronic fuel trimmingcircuit

The discriminating device willcomplete the feather cycle when thecondition lever is placed in featherand the power lever is placed inwhat position?

1. In any position below the flightidle position only

2. In any position above the flightidle position only

3. In any taxi range position only4. In any position

1. Feather2. Run3. Ground stop4. Air start

6-6. The reference temperature and turbineinlet temperature signals sent to thetemperature datum control indicate adifference greater than 1.9°C, and acontrol signal is sent to thetemperature datum valve. With thepower lever in the temperaturecontrolling range and the TEMP DATUMswitch in the AUTO position, thecontrol signal causes the temperaturedatum valve to

adjust the power lever assembly,regulating fuel flow to theengine being controlledcontrol a fuel-flow stabilizingpump on the engine beingcontrolledregulate the fuel flow to theengine being controlledreadjust the engine coordinatorand trimming circuit on theengine being controlled

6-7. Engine speed is less than 94% and theengine coordinator is set above 66°.fit what temperature is the normallimiting temperature automaticallyset?

1. 730°C2. 830°C3. 978°C4. 1,077°C

6-8. Dual unit thermocouples are radiallymounted in what part of the engine?

1. Turbine inlet case2. Compressor section3. Inlet guide vanes4. Turbine section

By which of the following methodsare thermocouples electricallyconnected to provide an averagetemperature?

1. In series2. In parallel3. Either 1 or 2 above, as both

will provide an averagetemperature

4. In series-parallel

An air turbine starter can beoperated by compressed air from aGTC, an APU, or what other device?

1. An operating engine2. The starter3. An outside air vent4. An emergency air tank

What is the function of the enginestart system’s speed-sensitivecontrol?

1. To synchronize all engine speeds2. To control engine RPM during

start cycles3. To prevent the engine from

exceeding maximum RPM4. To activate internal switches at

predetermined intervals relativeto the engine’s normal speed

The ignition exciter provides whichof the following voltages?

1. A stepped-up voltage for firingthe ignition plugs

2. A 28-volt excitation voltage forclosing the ignition relay

3. A 28-volt excitation voltage foractivating the speed-sensitivecontrol

4. A sine-wave voltage forapplication to the ignitionrelay solenoid

6-13. What behavior of the paralleling lampindicates that the secondary elementof a fuel pump has failed?

1. It illuminates continuously2. It never illuminates3. It illuminates above 65% RPM4. It illuminates between 16% and

65% RPM

6-14. Along with the temperature datumvalve, the fuel control functions toprovide a starting fuel flow scheduleand to

1. reduce nominal fuel requirements2. prevent engine overtemperature

and compressor surge3. operate hydraulically actuated

fuel cutoff valves4. close the fuel shutoff valve

during compressor surges

6-15. At what percentage, if any, of therated engine RPM does the fuelcontrol shutoff valve open to permitfuel flow to the engine?

1. 16% RPM2. 65% RPM3. 94% RPM4. None

6-16. The temperature datum valve islocated in which of the followingpositions?

1. Between the fuel tank and thefuel control

2. Between the fuel control and theengine fuel nozzles

3. Between the primary fuel pump andthe secondary fuel pump

4. Between the fuel tank and theprimary fuel pump

6-17. What is the function of compressorbleed-air valves?

1. To sequentially close the eightparts on the compressor housing

2. To bleed air from the com-pressor’s fifth stage into itstenth stage

3. To reduce the compressor loadduring starts

4. To minimize the starter load onthe compressor

IN ANSWERING QUESTIONS 6-18 THROUGH 6-21,REFER TO FIGURE 5-8 AND TABLE 5-1 IN YOURTEXT.

6-18. The engine selector switch is placedin the engine No. 1 position, andthe fuel and ignition switch forengine No. 1 is ON. What willhappen to (a) the fuel control relayand (b) the temperature datum relay?

1. (a) Energize (b) de-energize2. (a) De-energize (b) energize3. (a) Energize (b) energize4. (a) De-energize (b) de-energize

6-19. Which of the following conditionsexists when the start control valveis energized?

The engine start switch isdepressedThe selector switch is inposition 1The speed-sensitive controlswitch (65%) is closedEach of the above

6-20. The yellow starter valve lightsilluminate when which of thefollowing conditions exists?

1. The primer switch is depressed2. The primer valve solenoid is

energized3. The start control valve is open4. The speed-sensitive control is

inoperative

6-21. Which of the following events occurswhen approximately 65% RPM isreached?

1. Switches in the speed-sensitivecontrol will close

2. The ignition relay will energize3. The fuel pumps will operate in

series4. A series light will illuminate

Learning Objective: Recognize theoperating principles andcharacteristics of aircraft powerplant, anti-icing and deicing, firewarning and extinguishing, fueltransfer, and oil temperature controlsystems.

6-22. Refer to figure 5-9 in your text.Anti-icing air is fed to the enginewhen the solenoid is energized.Energizing the solenoid causes thepressure in the anti-icing valve to

1. decrease in the poppet valve body2. increase in the poppet valve body3. increase on the face of the main

poppet4. decrease on the face of the main

poppet

6-23. For any slow-cycle operation, thetiming cycle for propeller deicersis such that current is supplied tothe heating elements forapproximately what time period?

1. 17 to 22 seconds2. 25 to 35 seconds3. 40 to 75 seconds4. 80 to 120 seconds

6-24. The propeller deice timer motor ischanged from fast speed to slowspeed by which of the followingactions?

1. Switching a filter in the motorcircuit

2. Bypassing the variable resistorwith two fixed resistors

3. Adjusting the variable resistorto provide maximum resistance

4. Switching an additional fixedresistor in series with thevariable resistor

6-25. Under what condition will a firewarning light illuminate?

When the control unit does notmonitor resistance changesWhen the resistance of thesensing element does not changewith a change in enginecompartment temperatureWhen the resistance of thesensing element decreases to apredetermined level due to anincrease in temperatureWhen the resistance of thesensing element increases to apredetermined level due to anincrease in temperature

6-26. What is the purpose of the shortdiscriminator circuit in the firewarning system?

1. To illuminate the fire warninglights during a test

2. To activate the fire warningsystem if an short occurs in thesystem

3. To prevent the fire warningsystem from actuating when an

open occurs in the circuit4. To prevent the fire warning

system from actuating when ashort occurs in the circuit

6-27. Which of the following statementsdescribes the means by which CF3Brextinguishes

1. It formsengine’sthe fire

2. It cools

an aircraft engine fire?

a blanket around theair passages, smothering

the burning area to anextremely low temperature,extinguishing the fire

3. It uses the oxygen in thecompartment at a rapid rate,making the air incapable ofsupporting a fire

4. It displaces the air in thenacelle, making the air incapableof supporting a fire

6-28. What component controls the oilcooler door position when the oilcooler switch is in the automaticmode?

1. A thermostat2. A thermistor3. A magnetic brake4. A solenoid valve

On an aircraft, what control systemserves to vary the exhaust escapearea to obtain the desired thrustand to maintain safe operatingconditions?

1. Afterburner control system2. Variable exhaust nozzle system3. Main fuel control system4. MIL control system

What VEN system components serve toschedule, compute,engine operation?

1. VEN power unit2. ECA3. VEN actuator4. MFC

When the throttle is moved to theidle POSition, the VEN area rapidlymoves to almost full open. The fullopen VEN area has which of the

and control

following effects on engineperformance?

1. It reduces running temperatureand exhaust temperature

2. It allows higher idle speed;reduces acceleration time

3. Aids starting; lowers thrust4. Both 2 and 3 above

Decreasing the VEN area has whateffect, if any, on EGT?

1. Increases EGT2. Decreases EGT3. Increases EGT up to 9,000 feet;

decreases EGT above 9,000 feetof altitude

4. None

Learning Objective: Recognize theoperating principles andcharacteristics of the aircraftvariable exhaust nozzle controlsystem.

6-33. What component provides feedback tothe ECA to ensure the VEN ispositioned correctly?

1. The VEN synchronizing shaft2. The VEN position transmitter LVDT3. The VEN torque motor feedback

circuit4. The metering valve position

transmitter LVDT

6-34. What is the function of the VEN powerunit?

1. To provide electromechanicalfeedback to position the VEN

2. To provide hydraulic pressure tothe actuator to position the VEN

3. To provide power for theelectrical control assembly

4. To provide hydraulic pressure todrive the servomotor whensignaled by the ECA

Learning Objective: Recognize theoperating principles andcharacteristics of the propellersynchrophasing systems.

6-35. What is the function of the propellergovernor?

To control engine speed byvarying the pitch of thepropellerTo control engine speed byvarying the rate of fuel flow tothe fuel nozzlesTo control propeller pitch byvarying hydraulic pressureTo control the propeller byvarying engine speed

The pitch of the propeller blade isvaried by porting hydraulic fluiddirectly to which of the followingparts of the propeller piston?

1. Inboard side2. Outboard side3. Both 1 and 2 above4. Geared cam

The synchrophaser will function inwhich of the following modes?

1. Normal2. Synchrophasing3. Both 1 and 2 above4. Mechanical

What is the function of the pulsegenerator in the synchrophasersystem?

1. To provide pulses for phase andspeed control of the propellers

2. To establish the phaserelationship between propellers

3. To translate synchrophaserelectrical signals intomechanical bias

4. To provide pulses directly tothe corresponding servomotor

What is the purpose of the phase andtrim control tn the synchrophasersystem?

1. To control the servomotor train2. To provide pulses for speed

control3. To set the phase relationship

between master and slavepropellers

4. To convert electrical signalsinto mechanical motion

6-40. What is the function of the speedbias servo assembly?

1. To set the phase relationshipbetween the master and slavepropellers

2. To translate synchrophaserelectrical signals intomechanical bias

3. To establish the reference pulsefor the propellers

4. To provide pulses for speed andphase control

6-41. The synchrophaser provides which ofthe following servomotor controlvoltages?

1. An ac voltage 90° or 270° out ofphase with the excitation voltage

2. An ac voltage in phase or 180°out of phase with the voltageapplied to the reference

3. An ac voltage in phase or 180°out of phase with the excitationvoltage

4. A negative or positive dc signalvoltage

6-42. In the input windings of the magneticmodulator, what parameters of thecurrent control the amplitude andphase of the modulator’s output?

1. Amplitude controls magnitude, andphase controls direction

2. Magnitude controls amplitude, andphase controls direction

3. Amplitude controls magnitude, anddirection controls phase

4. Magnitude controls amplitude, anddirection controls phase

In a synchrophaser, what governoringmode is used to provide improvedengine response to transient RPMchanges?

1. Throttle lever anticipation mode2. Master governoring mode3. Synchrophasing mode4. Normal mode

As the power lever is moved todecrease power, a more positivevoltage is applied to the speedderivative circuit. What is theresult of this action?

A leading voltage surge appearson the servomotor controlwinding of the mechanicalgovernorA change in the speed or phaseof the master engine withrespect to the slave engineResetting of the mechanicalgovernor towards a decrease inpropeller pitchResetting of the mechanicalgovernor towards an increase inpropeller pitch

In the synchrophaser, the speedderivative circuit senses changes inengine RPM and generates signalshaving what function?

1. To change the propeller pitch toalign the slave propellers tothe master propeller

2. To dampen engine RPM changes3. To change the propeller pitch to

correspond to engine RPM4. To dampen propeller pitch

transients

6-46. Refer to figure 5-36 in your text.The magnitudes of the signal input tothe No. 1 winding of the magneticmodulator vary as a result of the

1. synchrophaser output signalchanges

2. reflected impedance changes fromthe No. 2 winding and the dummyload

3. voltage changes in the input tothe magnetic modulator

4. frequency changes in the engine’stachometer generator

6-47. When the synchrophasing mode ofoperation is used, all engines exceptthe master engine operate withsynchrophasing. With what does themaster engine operate?

1. Normal governing only2. Mechanical qoverning only3. Normal and mechanical4. Hydraulic governing

governing

PROPELLER SYNCHRONIZING

SLAVE PULSES ARE COMPAREDWITH MASTER PULSES, AND ERRORVOLTAGES ARE DELIVERED TO THEAMPLIFIER AND SPEED BIASSERVOMOTOR SIMULTANEOUSLY.

PULSES FROM THE PROPELLER AREFED TO THE SYNCHROPHASER,SLAVE PULSES ARE COMPAREDWITH THE MASTER PULSES,ERROR VOLTAGES ARE DELIVEREDTO THE AMPLIFIER, AND THEAMPLIFIER OUTPUT IS DELIVEREDTO THE PILOT VALVE.

PULSES FROM THE PROPELLERSARE FED TO THE SYNCHROPHASER,SLAVE PULSES ARE COMPAREDWITH THE MASTER PULSES,AND ERROR VOLTAGES AREFED DIRECTLY TO THE TWO-PHASEMOTOR IN THE SERVO CONTROLASSEMBLY.

PULSES FROM PROPELLERS ARE FEDTO THE SYNCHROPHASER, SLAVEPULSES ARE COMPARED WITH THEMASTER PULSES, ERROR VOLTAGESARE ROUTED TO THE AMPLIFIER,THE SPEED BIAS SERVOMOTORREACTS TO THE AMPLIFIER SIGNAL,AND THE SLAVE ENGINE SPEEDCHANGES.

Figure 6A

IN ANSWERING QUESTION 6-40, REFER TOFIGURE 6A.

6-48. Of the sequences listed, which oneresults in the propellers of thesynchrophasing system becomingsynchronized?

1. A2. B3. C4. D

6-49. Refer to figure 5-37 in your text.During a propeller underspeedcondition, V206A experiences which ofthe following input voltage changes?

1. A slow change from positive tonegative

2. A sudden change from negative topositive

3. A sudden change from positive tonegative

4. A slow change from negative topositive

6-50. Refer to figures 5-37 and 5-40 inyour text. Zener diodes VR405 andVR408 function to connect the speed-error circuit to the summing pointwhen an off-speed condition exists.What is their function, if any,during an on-speed condition?

To connect the speed-errorcircuit potentiometer to groundTo prevent formation oftransients in the speed-errorcircuit potentiometerTo isolate the speed-errorcircuit potentiometer from thesignal summing pointNone; they function identicallyduring on- and off-speedconditions

6-51. What synchrophaser circuit preventsthe slave engine from following anoverspeeding or underspeeding masterengine?

1. The2. The3. The4. The

2% limiting circuit20% limiting circuitspeed bias servo assemblythrottle lever anticipation

potentiometer

6-52. What is the purpose ofresynchrophasing?

1. To overcome small errors of leadand lag about a set point

2. To provide for antihunt bycorrecting for overshoot

3. To move the feedbackpotentiometer and cancel aportion of the error signal

4. To correct for accumulated one-direction errors

Learning Objective: Recognize theoperating principles andcharacteristics of the propellercontrol l nd approach powercompensator systems.

6-53. What assembly maintains the minimumdesired low-pitch angle?

1. Magnetic latching2. Hydraulic3. Electrical4. Low-pitch

leversolenoidstop

6-54. What is the purpose of the Beta 6-58. When needed, the negative torque

follow-up stop?

1. To provide a secondary stopsetting at the 15° blade angleonly

2. To provide a secondary low-pitchstop

3. To prevent negative-torque systemfailure

4. To prevent minor reductions inblade angle

The function of the pitchlockmechanism in the propeller is toprevent which of the followingconditions?

1. Propeller overspeeding2. Low-pitch oscillations3. Propeller gyrations4. High-pitch oscillations

Pitchlock is blocked out betweenblade angles of +57° and +86°. Whataction does this permit?

1. RPM surges during approaches2. RPM changes during landings3. The blade angle reduction for

takeoffs and landings4. The blade angle reduction for air

starting

If the fuel governor and thepropeller pitchlock test switch wereto be placed in the TEST position,what would be the result?

1. The blade angle would be reset tothe proper angle for an air start

2. The propeller governor RPM wouldbe reset to permit ground checkof pitchlock and fuel governorfunctions

3. A momentary blockout of thepitchlock mechanism

4. An increase in pitch

system functions in what way?

1. To eliminate propeller cyclingactions

2. To generate a negative torque3. To limit positive horsepower4. To increase blade angle

If the negative torque system failsduring in-flight unfeatheroperations, what system limitsnegative torque?

1. NTS INOP warning2. Feather position activation3. Airstart blade angle4. Emergency override

The process of aligning thepropeller to the airstream tominimize drag during engine shutdownconditions is known as

1. negative torquing2. feathering3. pitchlock4. Beta follow-up

The function of the feather pumppressure cutout override switch isto bypass the feather pressureswitch to prevent which of thefollowing occurrences?

1. The premature commencement ofthe minimum feather position

2. The late commencement of theminimum feather position

3. The premature termination of thefull feather position

4. The late termination of the fullfeather position

When is the automatic featheringsystem used?

1. During takeoffs only2. During landings only3. During takeoffs and landings4. During emergencies

To activate the thrust signal deviceto cause autofeathering, what must bethe condition of (a) theautofeathering arming switch and (b)the power lever quadrant switch?

1. (a) Closed (b) open2. (a) Open (b) closed3. (a) Closed (b) closed4. (a) Open (b) open

What condition is indicated when thefour autofeather system indicatorlamps are illuminated?

1. All autofeather systems are armed2. The entire autofeather system is

malfunctioning3. Only one of the propellers is

being automatically feathered4. All of the propellers are being

automatically feathered

6-65. When the power levers (throttles)are set in the taxi range, thepropeller blade pitch and enginefuel flow are controlled by which ofthe

following means?

The prop governors control theprop pitch, and the temp datumsystem controls the engine fuelflowThe synchrophaser systemcontrols the prop pitch, and thepower levers control the enginefuel flowThe temp datum system controlsboth the blade pitch and theengine fuel flow throughhydromechanical linkageThe power levers control boththe blade pitch and the enginefuel flow through thehydromechanical linkage

6-66. Which of the following is a functionof the approach power compensatorsystem?

1. To position the aircraft controlsurfaces automatically duringlanding approaches

2. To provide sudden surges of fuelfor additional altitude whenrequired

3. To control the engine powerautomatically during landingapproaches

4. To maintain the speed of theaircraft in moderate turbulencewithin ±10 knots

6-67. In the approach power compensatorsystem, what component/aircraftsystem provides the computer withaircraft approach angle information?

1. The pitot-static system2. The static system3. The angle-of-attack indicator4. The angle-of-attack transducer

6-68. Which of the following statementsdescribes the approach powercompensator system?

The computer provides nocorrective signal to repositionthe throttle as long as theacceleration is 1 g, and theangle of attack is optimum forlanding approachesThe system is capable ofoperating in three ambienttemperature ranges to controlengine performanceThe rotary actuator motorpositions the engine controllinkage after receiving acorrective signal from thecomputerEach of the above

6-69. After spreading the blades on theH-60 helicopter, the bladefoldactuators run in the fold directionto relieve the stress on whatcomponent?

1. The bladefold actuator2. The segment gear in the bladefold

actuator3. The blade lockpins4. The hinge lug

6-70. At supersonic speeds greater thanMach 2, the variable inlet duct rampsystem allows which of the followingevents to occur?

Supersonic air to enter theinlet duct onlySubsonic air to enter the engineonlyBoth subsonic l nd supersonic airto enter the inlet ductSupersonic air to increase invelocity before entering theengine

Textbook Assignment: “Aircraft Instruments.” Pages 6-1 through 6-46.

7-4.Learning Objective: Recognize theoperating principles and features ofthe pitot-static system.

7-1. Which of the following statementsdescribes the earth’s atmosphere?

1. The air molecules are closertogether at the bottom of theatmosphere than at the top 7-5.

2. The weight of the air pressingdown from above determines theair pressure at any givenaltitude

3. The air is more dense on theearth’ s surface than at analtitude of 1,000 feet

7-2. The altimeter is part of whataircraft system?

1. Compass2. AFCS3. Pitot-static4. Radio

7-3. In the pitot-static system, the termpitot represents what type ofpressure?

In computing airspeed, the airspeedindicator uses which of thefollowing pressures?

1. Static pressure only2. Impact pressure only3. The sum of impact and static

pressures4. The difference between static

and impact pressure

The accuracy of the airspeedindicator readings may be affectedby which of the followingconditions?

Temperature changes in theinstrumentAir turbulence around the pitottubeImperfect scaling of theindicator dialEach of the above

1. Impact2. Ambient3. Barometric4. Stationary

7-6. An aircraft’s Mach indicator reads0.5 when the airspeed indicator shows300 knots. If the airspeed were todouble to 600 knots, which statementwould reflect the aircraft’s speedand Mach indication?

1. The aircraft is at the speed ofsound, and the Mach indication is0.25

2. The aircraft is at the speed ofsound, and the Mach indication is1

3. The aircraft is at one-half thespeed of sound, and the Machindication is 1

4. The aircraft is at the speed ofsound, and the Mach indication is1.5

7-7. Which of the following is a meaningof the term altitude?

1. Elevation2. The distance above mean sea level

(MSL)3. The distance above ground level

(AGL)4. Each of the above

7-8. At what altitude does gravity actingon the atmosphere produce a pressureof 14.70 PSI and support a column ofmercury to a height of 29.92 inches?

1. Pressure altitude2. MSL3. Absolute altitude4. Density altitude

7-9. The altitude reading of a properlycalibrated altimeter referenced to29.92 inches of mercury (Hg) is knownas the

1. true altitude2. absolute altitude3. pressure altitude4. indicated altitude

7-10. If, at sea level, there is abarometric change of 0.03 inches,the altimeter reading would changeby how many feet?

7-11. At what altitude does the barometricpressure setting for aircraftaltimeters change from the localbarometric pressure to 29.92 inches?

1. 8,000 feet2. 13,000 feet3. 18,000 feet4. 23,000 feet

7-12. For the pilot to obtain the bestperformance from the aircraftengine, which of the followingaltitudes should be known?

1. True2. Density3. Indicated4. Calibrated

7-13. If an aircraft is in level flight,what will the vertical speedindicator (VSI) pointer indicate?

1. -0.52. 0.03. +0.54. +1.0

7-14. In a VSI indicator, the pointer isdriven by

1. the difference between diaphragmand case pressures

2. the difference between pitot andstatic pressures

3. a pneumatically driven motor4. a calibrated leak

Learning Objective: Recognize theprinciples and features of theairspeed indicating system.

7-15. What is the purpose of the air datacomputer (ADC)?

To provide accurate air data thatis free of aircraft configurationerrorsTo compute air data informationthat is gathered from sensorsisolated from air datadisturbancesTo sense the characteristics ofthe air surrounding the aircraftTo sense the characteristics ofthe air surrounding the aircraftand correct the data tocompensate for aircraft inducederrors

7-16. What are the four data sense inputsto the ADC?

1. Pitot pressure, static pressure,total temperature, and AOA

2. Pitot pressure, static pressure,angle of attack, and totalpressure

3. Pitot pressure, totaltemperature, cabin pressure, andangle of attack

4. Static pressure, cabin pressure,total pressure, and pneumaticdifferential pressure

7-17. Refer to table 6-1 in your text.What symbol represents the correctimpact pressure?

1. PSI2. Hp3. Pt4. Qc

7-18. Which of the following is themeaning of the term angle of attack?

The difference between theleading edge of the wing andnose of the aircraft relative tothe air through which it ispassingThe angle at which the leadingedge of the wing must pass toprovide adequate lift forsustained flightThe angle at which the leadingedge of the wing encounters theair massThe angle of the air passingover the elevators to providemore lift

7-19. What is the purpose of thepotentiometers in the angle-of-attack transmitter?

To provide an electricalindication of the aircraftfuselage in reference to theangle of attackTo provide a means forconverting voltage intomechanical motionTo convert electrical signalsinto mechanical signalsindicative of the angle ofattackTo convert mechanical motioninto proportional electricalsignals

7-20. What is the purpose of the angle-of-attack system?

1. To indicate aircraft totalpressure with respect to ambientpressure

2. To indicate the aircraft attitudewith respect to the surroundingair mass

3. To provide indications of the airdata sensor outputs

4. To compute the air datainformation for the ADCS

7-21. Static pressure errors become asignificant factor in the accuracy ofpressure indications at what relativespeeds?

1. Supersonic only2. Transonic only3. Supersonic and transonic4. Subsonic and transonic

7-22. There are two variables that causethe most significant errors inindicated static pressure as detectedby the aircraft static ports. One ofthese is the Mach number (aircraftspeed) and the other one is the

1. angle of attack2. altitude3. total temperature4. ambient temperature

7-23. What is meant by impact pressure(Qc)?

The weight of the air on theaircraftAtmospheric pressure, includingdisturbancesThe force of the air against theaircraftAtmospheric pressure free ofdisturbances

7-24. The two temperatures that make upthe total temperature are theambient temperature plus the

1. temperature of the engine intakeram air

2. temperature increase created bythe motion of the aircraft

3. temperature decrease created bythe surrounding air

4. temperatures of the ram airexternal to the aircraft and theengine intake ram air

7-25. What is the function of the AICS?

1. To acceleratethe engines

2. To accelerateprovide it to

3. To decelerateprovide it to

4. To decelerate

subsonic air for

supersonic air andthe enginesubsonic air andthe enginesupersonic air to

provide subsonic air for theengine

Learning Objective: Recognize theoperating principles and features ofthe automatic altitude reportingsystem.

7-26. In the automatic altitude reportingsystem, position and altitudereporting is accomplished by whichof the following components?

1. Transponders2. Transformers3. Flashing beacons4. Tachometer generators

The automatic altitude reportingsystem provides the aircraft’saltitude in which of the followingincrements?

What total number of dimensionsis/are automatically displayed on aradar presentation by thesemiautomatic air traffic controlsystem?

1. One2. Two3. Three4. Four

Which of the following is the primaryreason for the development of theAIMS system?

1. To improve radar maintenance2. To shorten the interval between

carrier landings3. To reduce the number of aircraft

in the air at any one time4. To improve air traffic control

within the United States

The maximum allowable differencebetween the aircraft’s altitude andthe ground display altitude in theAIMS system is what total number offeet?

1. ±50 feet

2. ±125 feet

3. ±250 feet

4. ±200 feet

Most high-performance naval aircraftuse which of the followingaltimeters as part of the AIMSsystem?

1. AAU-18/A2. AAJ-19/A3. AAU-21/A4. AAU-24/A

What are the operational limits forthe CPU-46/A altitude computer?

1. 25,000 feet altitude and Mach 82. 25,000 feet altitude and Mach

2.5 3. 80,000 feet altitude and Mach

2.54. 80,000 feet altitude and Mach 3

If the CPU-46/A altitude computerwere to fail, what would happen tothe AAU-19/A altimeter and thealtitude reporting encoder?

1. The altimeter automaticallyreverts to the pneumatic standbymode, and the encoder isdeactivated

2. The altimeter becomesinoperative and must be switchedto the standby mode; the encodercontinues to operate

3. The altimeter and encoder areinoperative

4. The altimeter is inoperative,and the encoder continues tofunction

The AAU-19/A altimeter operates as astandard altimeter when it is placedin which of the following modes?

1. Reset2. Servoed3. Baroset4. Standby

What components in the AAU-24/Aaltimeter overcome the effects of thestop-and-jump friction?

1. Servos2. Synchros3. Transponders4. Vibrators

When turning the baroset knob on theAAU-24/A altimeter, what devicedrives the barometric counter?

1. A counter-drum assembly2. A spur gear3. A magnetic coupler4. An aneroid assembly

Learning Objective: Recognize theoperating principles and features ofthe angle-of-attack (AOA) indicatingand stall warning systems andgyroscopic instruments.

When used in naval aircraft, theangle-of-sideslip system is usedalong with which of the followingsystems?

1. Crosswind landing system2. Gun firing system3. Bombing system4. Rocket firing system

The angle-of-attack indicating systemoperates by detecting

1. earth field variation2. airflow differential pressure3. the altitude/speed pressure

gradient4. the airflow/altitude change rate

If the angle-of-arrack of an air-craft is changed, which of thefollowing actions occur(s) in theself-balancing bridge circuit of theAOA system?

An error voltage will existbetween the transmitter andreceiver potentiometers in thecircuitA servomotor drives the receiverpotentiometer to return thebridge circuit to nullBoth 1 and 2 aboveA servomotor drives thetransmitter potentiometer toreturn the bridge circuit tonull

If the red arrow on the pilot’s AOAindexer illuminates, which of thefollowing conditions exist?

1. The aircraft is nose low2. The aircraft is nose high3. The aircraft is at optimum AOA4. There is a failure in the system

On most naval aircraft, the stallwarning system is activated by whichof the following systems?

1. Pitot-static system2. Angle-of-attack system3. Air data computer system4. Angle-of-sideslip system

Which of the following groups ofgeneral characteristics is the mostdesirable for an instrumentgyroscope?

1. Light weight for small size andlow speed of rotation

2. Light weight, small size, andhigh speed of rotation

3. Heavy weight, large size, andhigh speed of rotation

4. Heavy weight for small size andhigh speed of rotation

For a gyro to have two degrees offreedom, the platform must have whattotal number of gimbals?

1. One2. Two3. Three4. Four

What are the two fundamentalproperties for gyroscopic action?

1. Stability and rigidity in space2. Precession and centrifugal force3. Stability and centrifugal force4. Rigidity in space and precession

A spinning gyro precesses whensubjected to a deflecting force.Which of the following actions willmake the gyro precess at a fasterrate?

1. A decrease in the speed of therotor

2. An increase in the force applied3. Both 1 and 2 above4. An increase in the speed of the

What action does the gyro horizonhave, relative to its case, toindicate aircraft attitude?

1. The case revolves about the gyro2. The case and gyro spin axis are

both free to move with respect tothe aircraft

3. The gyro spin axis revolves aboutthe case

4. The gyro and case are heldstationary, and the needles arefree to move

7-47. The sphere in the attitude indicator

may be centered by a control on theface of the indicator to correct forwhich of the following flightattitudes?

1. Yaw2. Roll3. Pitch4. Each of the above

In a turn-and-bank indicator, whatfactor(s) determine(s) the positionof the ball?

1. Natural forces2. Gyroscopic precession3. Earth’s magnetic lines of force4. Electrical tilting of the plate

on which the indicator mounts

In a flight-coordinated turn, theball of a turn-and-bank indicatorwill be in which of the followingpositions?

1. The center2. To the left showing a slip3. To the right showing a skid4. Either 2 or 3 above, depending

on the direction of the turn

7-50. As the aircraft turns, the turn-and-bank indicator frame moves to theside opposite the turn. However theindicator needle moves in thedirection of the turn for which ofthe following reasons?

Centrifugal force and gyroprecession are in opposition toeach otherGravity is greater thancentrifugal force in any aircraftattitude other than straight-and-

level flightThe needle is connected to theframe in such a manner to causethe two to move in oppositedirectionsThe centrifugal force applied tothe needle is in the directionopposite to the centrifugal forceapplied to the frame

Learning Objective: Recognize theoperating principles and features ofmiscellaneous flight instruments, toinclude accelerometer indicators,clocks, direct-reading magneticcompasses, standby attitudeindicators, and outside airtemperature indicators.

7-51. What instrument does the pilot use toreduce the possibility of damage tothe aircraft from excessive stress?

1. The2. The3. The4. The

turn-and-bank indicatoraccelerometerangle-of-attack indicatorvertical gyro indicator

What is the reading on theaccelerometer of an aircraftweighing 40,000 pounds and travelingin straight-and-level flight?

1. +2 g2. +1 g3. 0 g4. -1 g

An aircraft weighed 10,000 pounds and has a lift equal to 20,000Pounds. Which of the followingreadings will the accelerometershow?

1. -2 g2. 0 g3. +1 g4. +2 g

The standard Navy aircraft clock haswhat type of movement?

1. 12-hour, 8-day2. 24-hour, 8-day3. 12-hour, 12-day4. 24-hour, 12-day

What is the purpose of filling thebowl of an aircraft direct-readingcompass with a liquid?

1. To keep the surface of thecompass card smooth

2. To slow the movement o+ thecompass card

3. To compensate for pressurechanges

4. To magnify the compass

What is the purpose of theline on the direct-reading

lubbercompass?

1. To reduce card oscillations2. To lock the card into position3. To serve as a reference mark

when reading the card4. To enable the compass to be used

as a standby unit

7-57. The standby attitude indicator iscapable of displaying a maximum ofhow many degrees of roll?

1. 79°

2. 92°3. 180°4. 360°

Learning Objective: Recognize theoperating principles andcharacteristics of various engineinstrument systems, includingtachometer, temperature indicating,fuel flow, oil pressure, fuelpressure, oil temperature, exhaustnozzle indicating, and torquemeter.(This objective is continued inassignment 8.)

A jet engine’s tachometer isreferenced to a percentage of whatRPM?

1. Flight idle2. Cruise3. Takeoff4. Military

The indicator in a tachometer systemis designed to respond

1. frequency2. current3. voltage polarity4. voltage amplitude

to changes in

7-60. What is the principle differencebetween a two-pole and a four-poletachometer generator?

1. The way their outputs are fed tothe indicator

2 The way they mount on the engine3. The way their stators are

constructed4. The polarity of their rotors

IN (ANSWERING QUESTIONS 7-61 AND 7-62, REFERTO FIGURE 6-45 IN YOUR TEXT.

At high speeds, the armature of asynchronous motor is brought intosynchronism by which of thefollowing components?

1. Hysteresis disk2. Hysteresis disk and the

permanent magnet3. Drag disk4. Hysteresis and drag disks

What is the purpose of concentratingthethe

flux near the outside edge ofdrag disk?

To provide the maximum torque toweight ratioTo stabilize the armaturerotational speedTo eliminate the effects oftemperature changesTo minimize the amount offriction encountered by thearmature

Which of the following arecharacteristics of the verticalscale indicator?

1. Its small size and light weight2. Its light weight and ease of

maintenance3. Its compactness and easily

readable scale4. Its compactness and ease of

maintenance

7-64. Activation of the test switch in avertical scale indicating system willcause the RPM indicator to read whatpercentage of RPM?

1. 40%2. 60%3. 80%4. l00%

IN ANSWERING QUESTIONS 7-65 AND 7-66, REFERTO FIGURE 6-48 IN YOUR TEXT.

7-65. When the galvanometers in a Whetstonebridge thermometer indicates zero,which of the following circuitconditions exist?

The bulb current and currentthrough resistor Y are equalThe sum of the voltage dropacross X and Z equals the sum ofthe voltage across the bulb and YThe combined resistance of X plusY equals the resistance of Z,plus bulb resistanceEach of the above

7-66. Which of the following quantities isindicated on the meter scale?

1. Resistance of X2. Temperature of the bulb3. Voltage between A and B4. Current flow between A and B

7-67. Refer to figure 6-49 in your text.In normal operation, an unbalance inthe radiometer bridge circuit iscaused by a change of resistance inwhich of the following components?

1. R12. The thermometer bulb3. The centering potentiometer4. R2

7-68. Thermocouple indicators operate onwhich of the following principles?

1. Dissimilar metal junctionsproduce an electromotive forcewhen heated

2. Changes in temperature causechanges in junction adhesion,producing an electromotive force

3. Changes in temperature causechanges in junction size,producing an electromotive force

4. Different metals bend atdifferent rates when heated,producing an electromotive force

7-69. A total of how many dual-unitthermocouples are mounted in eachengine turbine inlet casing of theP-3 aircraft?

1. 122. 183. 244. 36

7-70. What is the purpose of routingengine turbine inlet temperaturethermocouple wiring through aseparate harness to the TITindicator?

1. To shield the thermocouples fromheat

2. To concentrate heat on thethermocouples

3. To protect the bridge circuit ofthe cold junction compensator

4. To prevent outside signals fromchanging the average voltageproduced by the thermocouples

Textbook Assignment: “Aircraft Instruments” and “Compass and Inertial Navigation Systems.”Pages 6-45 through 7-20.

Learning Objective (Continued):Recognize the operating principlesand characteristics of various engineinstrument systems, includingtachometer, temperature indicating,fuel flow, oil pressure, fuelpressure, oil temperature, exhaustnozzle indicating, and torquemeter.

Thermocouples in an engine exhaustsystem convert exhaust gastemperature into which of thefollowing units of measurement?

1. Millivolts2. Milliamperes3. Milliwatts4. Milliohms

IN ANSWERING QUESTIONS 8-2FIGURE 6-53 IN YOUR TEXT.

AND 8-3, REFER TO

8-2. In a fuel flow transmitter, theincoming fuel is directed againstwhat component?

1. Ring magnet assembly2. Synchro transmitter3. Hairspring4. Vane

What force moves component 4 of thefuel flow transmitter?

1. The rotation of component 32. The mechanical force of

component 23. The magnetic force of component

54. The electrical power of

component 8

When the input fuel pressure to thefuel flow transmitter becomesexcessively high, the instrument isbypassed by what transmittercomponent?

1. Relief valve2. Baffle valve3. Check valve4. Intake valve

The fuel flow transmitter convertsfuel flow into an electrical signal,which represents the fuel flow inpounds per hour. This signal istransmitted to what component in thesystem?

1. The transmitter synchro receiver2. The fuel control regulator3. The synchro receiver in the fuel

flow indicator4. The bypass valve solenoid in the

fuel control regulator

In the fuel flow transmitter, whatcomponent eliminates the swirlingmotion of the incoming fuel?

1. Spiral springs2. Straightening vanes3. Rotating impeller4. Drum magnets

Refer to figure 6-54, view B, in yourtext. What quantity of fuel isremaining in the fuel cells?

1. 99 pounds2. 999 pounds3. 9,990 pounds4. 99,900 pounds

In the fuel flow totalizer indicator,what component causes the pointer todeflect proportionally to the fuelbeing consumed?

1. The friction clutch of the motorshaft

2. The magnetic drum and cup linkage3. The mechanical gear train4. The direct coupling of the motor

to the shaft

In a synchro oil pressure system, themovement of the indicator needledepends on which of the followingconditions?

1. Oil pressure in the engine2. Sensing the action of the oil

pressure transmitter3. Voltages set up in the synchro

stators4. Each of the above

8-10. Refer to figure 6-60 in your text.What is the purpose of the vent onthe fuel pressure transmitter?

1. To allow the transmitter toaccurately measure thedifferential pressure betweenthe pump and the atmosphere

2. To ensure synchro cooling whenaircraft engines are running

3. To equalize pressure in thetransmitter with that in thepressure converter

4. To vent moisture in thetransmitter case due tocondensation

8-11. Refer to figure 6-61 in your text,Movement of the transmitterpotentiometer will have which of thefollowing results?

1. Removes 28 volts from one of thecoils in the indicator

2. Removes 28 volts from both coilsin the indicator

3. Moves the nozzle positionindicator to the OFF position

4. Applies a signal to the nozzleposition indicator

8-12. The torquemeter system operates onthe principle that engine loadingwill have which of the followingresults?

Produce a measurable change inengine RPMProduce a measurable couplerchange between the splines ofthe two shaftsCreate a slight twist, which isdetected by magnetic pickups andmeasured electronicallyCreates shaft twist, whichcauses two magnets to measurethe distance from two toothedflanges

Learning Objective: Recognize theoperating principles andcharacteristics of miscellaneousaircraft instrument systems,including fuel quantity, hydraulicpressure, and position indicatingsystems.

8-13. In what units of measurement docapacitive-type fuel quantity gaugesindicate the fuel quantity in a tank?

1. Gallons2. Pounds3. Cubic inches4. Percentages

8-14. Refer to figure 6-65 in your text.The capacitance of the fuel gaugecapacitor depends on which of thefollowing factors?

1. The area of the plates2. The distance between the plates3. The dielectric constant4. Each of the above

8-15. Refer to figure 6-65 in your text.Which of the following conditionsexists in the circuit as the fuelquantity increases?

Tank unit capacitance increases,tank unit leg current decreases,and transformer voltage andvoltage across the voltmeter arein phaseTank unit capacitance increases,tank unit leg current increases,and transformer voltage andvoltage across the voltmeter arein phaseTank unit capacitance decreases,tank unit leg current decreases,and transformer voltage andvoltage across the voltmeter areout of phaseTank unit capacitance decreases,tank unit leg current increases,and transformer voltage andvoltage across the voltmeter areout of phase

8-16. The indicator’s motor direction ofrotation is determined by acomparison of what characteristicsof two voltages?

1. Polarity2. Phase3. Amplitude4. Frequency

8-17. Refer to figure 6-67 in your text.As the fuel level varies, thebalance of the bridge is maintainedby the automatic adjustment of whatvariable resistor?

1. R1212. R1223. R1284. R129

If two or more fuel tank units areconnected in parallel, the effects ofwhat condition is minimized?

1. A decrease in fuel2. An increase in fuel3. Variations in aircraft attitude4. Irregularities in fuel tank shape

Which of the following conditionswill cause the dielectric constantand fuel density to deviate?

1. Weight and temperature of thefuel

2. Temperature and composition ofthe fuel

3. Composition of the fuel and shapeof the fuel tank

4. Temperature of the fuel and shapeof the fuel tank

What unit of measurement is indicatedby hydraulic pressure indicators inmost naval aircraft?

1. Pounds per square inch2. Pounds per square foot3. Pounds per second4. Pounds per minute

In a dc position-indicating system,what is the purpose of the coppercylinder in the indicator?

1. To aid the magnetic field2. To maintain equal coil spacing3. To dampen pointer oscillation4. To position the transmitter

brushes

Learning Objective: Identify thevarious procedures used to maintain,troubleshoot, test, interchange,protect, ship, and handle aircraftinstruments.

Which of the following proceduresshould you follow to ensure theaccuracy of aircraft instruments?

1. Replace them rather than adjustthem

2. Replace them at regularintervals

3. Periodically subject them tofunctional tests

4. Periodically readjust theadjustable mechanisms

What is indicated by a red mark onan instrument glass?

1. The normal operating range ofthe instrument

2. Glass slippage3. A critical limit4. A maximum limit

What is the purpose of the whiteindex mark painted at the bottomcenter of all instruments color-marked for operating ranges?

1. To set the pointer at zeroreference

2. To designate the 6 o’clockposition of the indicator formounting purposes

3. To be used as a guide formounting the lighting masks

4. To show whether or not the glasscover has moved after theoperating ranges have beenmarked

You should perform which of thefollowing checks during the dailyinspection of an aircraft’sinstrument panel?

1. For the proper operation ofcaging and setting knobs

2. For the proper operation oflights

3. For loose or cracked cover glass4. Each of the above

Which of the following statementsabout the lubrication of aircraftinstruments is correct?

Lubrication is limited to thatrequired for the bearingsSquadron personnel rarelylubricate any part of theinstrumentsLubrication is done bythe direction of AEsLubrication is limitedrequired by the shafts

To what publication should

AMs under

to that

you referfor detailed information on aircrafttubing and tubing repair?

1. NAVAIR 01-1A-82. NAVAIR 16-1-5403. NAVEDTRA 10349-E4. NAVEDTRA 10348-F

Rigid tubing in modern aircraft maybe made from which of the followingmetals/alloys?

1. Copper2. Aluminum alloy3. Corrosion-resistant steel4. Both 2 and 3 above

Where are identification markingslocated on rigid tubing used in navalaircraft?

1. Near the fittings and in eachcompartment

2. Midway between each pair offittings

3. Not more than 1 inch from and onboth sides of each fitting

8-30. When identification tape is used onrigid tubing in military aircraft,the function of the tubing isidentified by the

1. wording on the 3/4-inch-widetape

2. color, wording, and symbols on1-inch-wide tape

3. color of a 1/2-inch tape4. color and wording on a 1-inch

8-31. Refer to table 6-3 in your text. Ofthe following groupings, which oneconsists of material classified asphysically dangerous and should havePHDAN printed on identifying tape orlines carrying the material?

1. Carbon dioxide, gaseous oxygen~nitrogen gas, and freon

2. Carbon dioxide, liquid nitrogen,liquid oxygen, and alcohol

3. Carbon dioxide, liquid nitrogen,nitrogen gas, and JP-4 fuel

4. Carbon dioxide, liquid nitrogen,gaseous oxygen, and liquidpetroleum gas

8-32. What method is used to identify thespecifications of flexible tubinginstalled on naval aircraft?

1. The type of fittings2. Colored tape bands applied to

the tubing3. A code of dots and dashes

printed on the tubing4. Color-coded bands of paint

printed on the tubing

4. Anywhere on the fittings

Which of the following statementsabout the replacement of high- andlow-pressure flexible hose iscorrect?

Low-pressure hose may befabricated locally, and high-pressure hose must be obtainedthrough supplyBoth types of hose must beobtained through supplyHigh-pressure hose may befabricated locally, and low-pressure hose must be obtainedthrough supplyBoth types of hose may befabricated locally

What action should you take toprevent flexible hose from beingpulled loose from the engine due toengine movement during hoseinstallment?

1. Use cushion clips to hold thehose

2. Use support mounts every 24inches

3. Allow slack between the last hosesupport and the connection to theengine

4. Properly shroud the hose

Refer to table 6-5 in your text. Thepointer on the fuel flow indicatorswings to the side of the dial and asqueal is heard from the instrument.Which of the following conditionscould cause this malfunction?

1. A short in the system2. Power leads reversed to the

stator3. Both 1 and 2 above4. An open stator lead

When it is necessary to substituteone aircraft instrument for another,which of the following precautionsshould you take?

1. Make sure that the instrumentshave the same stock number

2. Make sure that the instrumentshave the same manufacturer’spart number

3. Make sure that the instrumentshave the same stock number andmanufacturer’s part number

4. Flake sure the instruments aremade by the same manufacturer

What method should you use todispose of containers in whichinstruments are received?

1. Return them to supply2. Route them to salvage for final

disposition3. Destroy them or label them NOT

FOR REUSE FOR INSTRUMENTS4. Retain them for future use in

packaging instruments

Learning Objective: Recognize thenavigation-related terms anddefinitions that are basic tocompass and inertial navigationsystems.

Which of the following statementsdescribes the purpose of airnavigation equipment?

1. It is used to measure distanceand time

2. It is uses to determine aircraftposition

3. It is used to determine thedirection an aircraft must flyto reach its destination

IN ANSWERING QUESTIONS 8-39 THROUGH 8-42,SELECT FROM COLUMN B THE TERM FOR EACHDEFINITION LISTED IN COLUMN A. ALL ANSWERSARE USED.

A. DEFINITION B. TERM

The intended 1. Coursehorizontal directionof travel 2. Heading

The horizontal 3. Bearingdirection of oneterrestrial point 4. Distancefrom another

The horizontal direction in which anaircraft is pointed

The separation between two pointsmeasured in some scaler quantity

8-43. Planes that pass through the earthperpendicular to the earth’srotational axis and intersect withthe earth’s surface to form circlesare known as

1. parallels2. perpendiculars3. meridians4. prime meridians

8-44. Relative to the earth’s surface, ifan aircraft were at latitude 37° Sand longitude 83° E, it would be atwhich of the following positions?

1. 37° south of Greenwich, England,and 83° east of the equator

2. 83° east arid 37° south ofGreenwich, England

3. 83° east of Greenwich, England,and 37° south of the equator

4. 37° south and 83° east of theequator

8-45. Convert the following coordinatesfrom decimal form to degree/minutes/seconds form:

Latitude--47.7° NLongitude--131.45° E

1. 47°60'4" N and 131°45' E,respectively

2. 42°36' N and 48°36' E,respectively

3. 47°42' N and 131°27' E,

respectively4. 48°15’ N and 131°75’ E,

respectively

IN ANSWERING QUESTIONS 8-46 THROUGH 8-49,SELECT FROM COLUMN B THE TERM FOR EACHDEFINITION LISTED IN COLUMN A. ALL ANSWERSARE USED.

A. DEFINITION B. TERM

An irregular line 1. Variationconnecting pointson a map of the 2. Agonicearth, indicating linewhere a compasspoints to true 3. Isogonicnorth line

An irregular line 4. Deviationconnecting pointson a map of the earth, indicatingwhere there is the same variationfrom true north

The angular difference between thedirections of true north andmagnetic north at a particularlocation

The angular difference between thedirection of the earth’s magneticfield and the compass reading due tonearby electromagnetic influences

8-50. If variation is 6° west and deviationis 1° west, the compass error isequal to

1. -6° + 1° = 5° east

2. -6° - l ° = 7° east3. 6° + 1° = 7° west

4. 6° - 1° = 5° west

8-51. If variation is (-) 3° and deviationis (+) 4°, the compass error is equalto

1. (-)3° + (+)4° = (+)1° or 1° east2. (-)3° - (+)4° = (-)7° or 7° east3. (-)3° + (+)4° = (+)1° or 1° west4. (-)3° - (+)4° = (-)7° or 7° west

8-52. The difference between the directionof the earth’s magnetic field and thehorizontal at any location on theearth’s surface is known as the

1. magnetic dip2. magnetic variation3. surface variation4. surface deviation

8-53. A line on a map that connects allplaces having equal dip angles isknown as an

1. agonic line2. aclinic line3. isobaric line4. isoclinic line

8-54. The aircraft navigator is plottingthe present position by usingaircraft course and speed, last knownposition, elapsed time, and anychanges in speed and course since thelast known position. What type ofnavigation is the navigator using?

8-55. Accelerometers in inertialnavigation equipment are mounted ona platform that remainsperpendicular to the earth’sgravitational field at all times tosense

1. gravity changes only2. vehicle accelerations only3. gravity changes and vehicle

accelerations

Learning Objective: Recognize theoperating principles and features ofcompass systems, the attitudereference system, and associatedsensors and indicators.

QUESTIONS 8-56 AND 8-57 RELATE TO A COMPASSSYSTEM THAT PROVIDES AIRCRAFT HEADINGINFORMATION IN ELECTRICAL SIGNAL FORM.

8-56. In the compass system, what unitsenses the direction of the fluxlines of the earth’s magnetic field?

1. The gyro amplifier2. The flux valve3. The magnetic amplifier4. The direct-reading compass

8-57. A displacement gyro provides whichof the following electrical signals?

1. Azimuth and pitch only2. Pitch and roll only3. Roll and azimuth only4. Azimuth, pitch, and roll

1. Mapping2. Pilotage3. Inertial4. Dead reckoning

QUESTIONS 8-58 THROUGH 8-60 RELATE TOVERTICAL GYROSCOPE OPERATION. IN ANSWERINGTHESE QUESTIONS, SELECT FROM COLUMN B THECONTROL TRANSMITTER THAT INITIATES EACH OFTHE SIGNALS LISTED IN COLUMN A. ALL ANSWERSARE NOT USED.

A. SIGNALS B. CONTROLTRANSMITTERS

8-58. Roll signals to 1. Outer rollthe AFCS control controlamplifier transmitter

8-59. Roll signals to 2. Pitch theindicators control

transmitter8-60. Pitch signals to

the indicators 3. Roll controltransmitter

4. Pitch servocontroltransmitter

8-61. The directional gyro pitch gimbal ismaintained perpendicular to thesurface of the earth by a motor-generator that is driven by theamplified output of

1. microswitches2. a generator dampened by a

pendulum-type weight3. the vertical gyro’s pitch servo

control transmitter4. electrolytic switches mounted on

the directional gyro’s pitchgimbal

8-62. The aircraft horizontal situationindicator will provide which of thefollowing items of information?

1. The selected course to anelectronic ground station

2. The aircraft’s deviation from aselected course

3. The pictorial bearing and thedistance in nautical miles to anelectronic ground station

8-63. The distance counter of the bearing-distance-heading indicator displayswhich of the following information?

1. Distance to base2. Distance to target3. Distance to a ground electronic

station4. Each of the above

IN ANSWERING QUESTIONS 8-64 THROUGH 8-67,SELECT FROM COLUMN B THE MODES THAT SHOULDBE USED FOR THE CONDITIONS LISTED IN COLUMNA. ALL ANSWERS ARE USED.

A. CONDITIONS B. MODES

Used in areas where 1. SLAVEDthe earth’s magneticfield is appreciably 2. FREEdistorted

3. COMPASSUsed in normalconditions

Used when the displacement gyro’ssignals are not reliable

Used when the flux valve signal isunreliable

8-68. What component provides the timing,switching, and voltages used forerection, monitoring, and levelingthe AHRS displacement gyroscopes?

1. Compass system controller2. Amplifier power supply3. Compass adapter compensator4. Displacement gyro motor-

generator

8-69. The compass adapter compensatorprovides correct heading informationto the attitude indicator when it isoperating in which of the followingmodes?

1. FREE2. COMPASS3. SLAVED4. Each of the above

8-70. The latitude degrees control on thecompass controller provides a signalto the directional gyroscope forcompensation of

1. mechanical drift2. gyroscope rigidity3. apparent drift4. magnetic variation

8-71. In which of the following modes ofoperation will the directional gyroprovide only azimuth information tothe indicator?

1. FREE2. COMPASS3. SLAVED4. Each of the above

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