x-15 flight manual

8/8/2019 X-15 Flight Manual

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T.O.1X-15-1(FORMERLYFHB-23-1)

AIR FORCE Kerr Li tho, Culver City, Cal 5Apr il 62-250 (Nor th American Aviat ion) 29 DECEMBER 1961


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T.O. 1X-15-1Reproduction for nonmilitary use of the information or illustrations contained in this publication is not per-mitted without specific approval of the issuing service. The policy for use of Classified Publications is establishedfor the Air Force in AFR 205-1.

INSERT LATEST REVISED PAGES. DESTROY SUPERSEDED PAGES.LIST OF EFFECTIVE PAGES NOTE: The port ion of the rext affected by the revisions is indicatedby a vertical line in the outer margins of the page.

TOTAL NUMBER OF PAGES IN THIS PUBLICATION IS 122, CONSISTING OF THE FOLLOWING:

PageNo. Issue

Title .OriginalA Originali . . . . . . . . . . . . . . . Original11Blank . . . . . . . . . . Originaliii thru iv . . . . . . . . . Original1-1thru 1-47 Original1-48Blank. . . . . . . . Original2-1thru 2-19 Original2-20 Blank ... . . . .Original .3-1thru 3-14. . . . . . . original"4-1 thru 4-11 qrigmal4-12Blank. . . . . . . . Origiriiu5-1 thru 6-15 Original5-16Blank. . . . . . . . Original6-1 thru 6-5 . . . . . . . Original6-6 Blank Original

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.The asterisk indicates pages revised, added, or deleted by the current revision.

ADDITIONAL COPIESOF THIS PUBLICATION MAY BE OBTAINED AS FOLLOWS:USAF ACTIVITIES - In accordance with Technical Order No. 00-5-2.

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USAF


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X15-1-0-5A

Vii

Section I DESCRIPTION 1-1Section II NORMAL PROCEDURES 2-1Section III EMERGENCY PROCEDURES 3-1ISection IV AUXILIARY EQUIPMENT 4-1

I Section V OPERATING LIMITATIONS 5-1Section VI FLIGHT CHARACTERISTICS 6-1I


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This utility Flight Manual is the result of extensive research and analysisof engineering data. It contains the necessary information ~or safe and_efficient operation of the X-15 Airplane. Information involving safetyof flight will be disseminated by means of the regular Safety of FlightSupplement program. You can determine the status of Safety of FlightSupplements by referring to the Safety of Flight Supplement Index, T. O.O-l-lA. The title page of the Flight Manual and title block of each Safetyof Flight Supplement should also be checked to determine the effect thatthese publications may have on existing Safety of Flight Supplements.The manual is divided into six separate sections, each containing itsown table of contents. The research program for which this airplanewas designed requires that each individual mission be precisely pre-planned. Consequently, standard performance data is not included inthis manual. The Flight Manual does not discuss in detail certain com-plex units installed in the airplane, nor does it necessarily containinformation on the use or operation of test equipment.

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T. O. lX-15-1 Section I

TABLE OF CONTENTS PAGEAirplane.. . . . . . . . . . . . . . . . . . . . . . . . 1-1Engine . . . . . . . . . . . . . . . . . . . . . . . . .. 1-3Propellant Supply System . . . . . . . . . . . . .. 1-15Engine and Propellant Control HeliumSystem.. . . . . . . . . . . . . . . . . . . . . . . .. 1-16Propellant Pressurization Helium System ... 1-16Auxiliary Power Units ................ 1-17

1-22lectrical Power Supply Systems . . . . . . . . .Hydraulic Power Supply Systems . . . . . . . . . 1-26Flight Control Systems . . . . . . . . . . . . . . . 1-27WingFlapSystem . . . . . . . . . . . . . . . . . .. 1-36Speed Brake System . . . . . . . . . . . . . . . . . 1-36AIRPLANE.The X-15 is a single-place research airplane, specifi-cally designed to obtain data onflight at extremely highaltitudes and speeds and on the physiological and psycho-logical effects of such flight conditions on the pilot.Built by North American Aviation, Inc, the airplane hasan inertial all-attitude (gyro-stabilized platform) flightdata system and is powered by one XLR99 liquid-propellant rocket engine. The 25-1/2 degree swept-back winghas hydraulically operated flaps on the inboardtrailing edge of each wing panel. All aerodynamiccontrol surfaces are actuated by irreversible hydraulicsystems. The horizontal stabilizer has a 15-degreecathedral. The two sections move simultaneously forpitch control, differentially for roll control, and incompound for pitch-roll control. The upper and lowervertical stabilizers are in two sections, a movableouter span for yaw control and a fixed section adjacentto the fuselage. The lower movable section (ventral)is jettisonable for landing. Each fixed section in-corporates a split-flap speed brake. For changes inairplane attitude relative to flight trajectory at altitudeswhere aerodynamic controls are relatively ineffective,the airplane incorporates a ballistic control system,wherein the metered release of gas through smallrockets in the nose and wing causes the airplane tomove about each axis as required. Two auxiliarypower units drive the airplane hydraulic pumps and acelectrical generators. Fuel for the rocket engine is

SECTION I

PAGELaunchSystem . . . . . . . . . . . . . . . . . . .. 1-36LandingGear System. . . . . . . . . . . . . . . .. 1-37Instruments . . . . . . . . . . . . . . . . . . . . .. 1-37Inertial All-attitude Flight Data System(Gyro-stabilized Platform) 1-38Instrumentation System . . . . . . . . . . . . . . . 1-40Indicator, Caution, and WarningLight System . . . . . . . . . . . . . . . . . . . . . 1-41Canopy . . . . . . . . . . . . . . . . . . . . . . . . . 1-41EjectionSeat . . . . . . . . . . . . . . . . . . . . . . 1-43Auxiliary Equipment . . . . . . . . . . . . . . . . . 1-46carried internally. The airplane is not designed fornormal ground take-off, but is air-launched by a B-52Airplane. The landing gear consists of a dual-wheelnose gear and two main landing skids. The gear islowered in flight by gravity and air loads.AIRPLANE DIMENSIONS.The over-all dimensions of the airplane (in-flight con-figuration with gear up and ventral retained) are asfollows:Length 49 feet 2 inchesSpan . . . . . . . . . . . . . . . . . . . . . 22 feet 4 inchesHeight . . . . . . . . . . . . . . . . . . . . . . 13 feet 1 inch

NOTEIn the landing configuration (landing grossweight and gear down, with specified nose tireand strut inflation and with ventral jettisoned),height is 11 feet 6 inches.

AIRPLANE GROSS WEIGHT.The approximate launch gross weight of the airplane(including full internal load and pilot) is 32,900 pounds.However, this can vary a few hundred pounds, dependingon the type of instrumentation carried.

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1. BALLNOSE2. BALLISTICCONTROLSYSTEM ROCKETS (8)3 . NOSE G EAR4. PIT OT HEADS. EJECTION SEAT6. HELIUM TANKS (2) FOR APU PROPELLANT PRESSURIZATION7. AUXIL IARY POW ER UNITS ( 2)8 . A IR CONDITIONING AND PRESSURIZ AT IO N SYSTEM

LIQUID NITRQGEN TANK9.8 -52 PYLON F ORWARD ATT ACH POINT10. L IQ UID O XYGEN VENT11. UPPER UHF ANTENNA12. L IQ UID O XYGEN T ANK ( ENGINE O XIDIZER)13. BALLlSfIC CONTRQL SYSTEM ROCKETS (2)14. HELIUM TANK (FOR ENGINE PROPELLANT PRESSURIZATION)IS. HELIUM TANK FOR PRESSURIZATION CONTROL, ENGINECONTROL, AND ENGINE HYDROGEN PEROXIDE PRESSURIZATION16. R IG HT -HAND W ING F LAP1 7. A MM ON IA T AN K18. 8 -S2 PYLON REARAT rACH POINT ( BO TH SIDES)

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19. HYDROGEN-PEROXIDE TANK (ENGINEWRBOPUMP PROPELLANT)20. MOVABLE UPPERVERTICAL STABILIZER21. UPPERSPEEDBRAKE22. XLR99-RM- l ENG INE23. HEL IUM T ANKS ( 2) F OR ENG INE PURGE24. MOVABLE HORIZONTAL STABILIZER2S. LOWER SPEEDBRAKE26. VENTRAL2 7. L AND ING GE AR S KI D l B OT H S IDE S)28. BALLISTIC CONTROL SYSTEM ROCKETS (2)29. NO.3 EQUIPMENT COMPARTMENT30. L IQ UID O XYGEN T ANK ( ENGINE O XIDIZER) F ILLER31. HYDRAULIC SYSTEM RESERVOIRS (2)32. HYDRO GEN- PERO XIDE T ANKS ( 2) APU ANDBALLISTIC CONTROL SYSTEMS PROPELLANT33. NO.2 EQUIPMENT COMPARTMENT34. LOWER UHF ANTENNA35. NO.1 EQUIPMENT COMPARTMENT X-I5-1-00_1B

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AffiPLANE SERIAL NUMBERS.The Air Force serial numbers for X-15 Airplanescovered by this manual are AF56-6670, -6671, and-6672.ENGINE.Thrust is provided by one XLR99 turbo rocket engine.It has a single thrust chamber, a two-stage, continuous-igniter starting system, a turbopump, and a gas gen-erator. Propellants are liquid oxygen and anhydrousammonia, supplied from the airplane propellant system.(See figure 1-6.) The engine is of variable-thrustdesign, capable of operating over the range of 50 to100 percent of full rated thrust. The gas generatordecomposes a monopropellant fuel, 90 percent hydrogenperoxide (H202), to provide a high-pressure gas mix-ture for driving the turbopump, which in turn drivesthe two centrifugal pumps that supply the propellants tothe engine. Upon discharge from the pumps, the pro-pellants are delivered to the two igniters and the thrustchamber where they are burned. At the first-stageigniter, the oxygen (in gaseous form) and ammonia aremixed and then ignited by three spark plugs. Liquidoxygen and ammonia, meanwhile, also are routed to thesecond-stage igniter. When the pressure created by thehot gases in the first-stage igniter actuates a pressureswitch, propellants are allowed to enter the second-stage igniter. Here, the propellants are mixed andignited by the incoming first-stage gases at greatlyincreased pressure. When a pressure switch in thesecond-stage igniter is actuated, propellants are allowedto enter the thrust chamber itself. They are againmixed and ignited by the gases coming from the second-stage igniter and build to the tremendous pressuresneeded for required thrust. The thrust chamber isan assembly of small welded, wire-wound tubes pre-formed as segments of the chamber. Before injectioninto the thrust chamber, the ammonia passes throughthese tubes to cool the chamber. Exhaust gases aredischarged through a venturi-shaped sonic nozzle.ENGINE COMPARTMENT.The engine compartment, in the extreme aft end of thefuselage, houses the tubular steel engine mount whichsupports the engine and turbopump. The engine com-partment is completely isolated from the airframe bya mono-fire-wall. A large access door is provided inthe forward end of the engine compartment for accessto the engine compartment from the hydrogen peroxidestorage tank area. The engine compartment alsohouses the instrumentation pick-offs, fire detectionsystem sensors, and helium release line. A fire sealcloses out the compartment and protects against theentry of exhaust gases and expelled propellants intothe engine compartment. For engine compartmentpurging, refer to "Engine Compartment Purging System"in this section.Engine Compartment Fire Detection System.A detection circuit is provided to detect and indicatea fire condition in the engine compartment. This

Section I

circuit is of the continuous-element type, which detectsexcessive temperatures anywhere along its length. Thesystem is powered by the battery bus and continuouslymonitors the resistance of the circuit. The resistanceof the material used in the circuit varies inversely withtemperature and total length of the sensing circuit.Whenever temperatures in the engine compartmentreach 1l00F (594C) or higher, the resistance of thesensing element falls below a preset value because ofthe excessive temperature, and the warning system isenergized. A placard-type warning light, a systemselector switch, and a test switch are in the cockpit.For emergency procedures in case of a fire-warningindication, refer to "Fire or Explosion" in Section m.Fire-warning Light. An abnormal rise in engine com-partment temperature is shown by a placard-typewarning light (70, figure 1-2), on the instrument panel.The light is powered by the primary dc bus and has ared plastic cap which shows the word "FIRE" when thelight is on. The light may be tested by a push-to-testswitch on the instrument panel right wing.Fire-warning Light Test Button. A fire-warning lightand detection circuit test button (30, figure 1-2) is onthe instrument panel right wing. The button is poweredby the primary dc bus. When the button is pressed, thefire-warning light should come on, verifying the con-tinuity of the detection circuit.Engine Compartment Purging System.The engine compartment can be purged by releasing aninert gas (helium) under pressure into the area toextinguish a fire or relieve an overheat condition.Three cubic feet of helium is stored in two sphericalcontainers under 3600 psi pressure. The containersare on either side of the engine compartment in theleft and right wing root fairing tunnels. Either auto-matic or manual release of helium can be selected bythe pilot. Because of the location of the two containersadjacent to the engine compartment, any high-temperature condition in the compartment will affectthese containers and create a potential explosion hazard.As the helium is released into the engine compartment,it inhibits any fire condition and at the same timeeliminates the explosion hazard.Helium Release Selector Switch. This three-positionswitch (69, figure 1-2), labeled "HE REL SW, " is onthe left side of the instrument panel. It permits thepilot to select the type of engine compartment purging(either automatic or manual) in case of a fire indication.The switch is powered by the battery bus. The AUTOposition sets up an entirely automatic sequence if a fireoccurs in the engine compartment, as indicated byillumination of the fire- warning light. The engine isshut down, and the helium from the two containersoutboard of the engine is jettisoned into the enginecompartment to inhibit the fire or overheat conditionand to prevent overpressurization of the containers byextreme temperature increase. The OFF position setsup the fire detection system for illuminating the fire-warning light only, in case a fire occurs. It will thenbe necessary either to move the switch to ON to releasethe helium into the engine compartment without affecting

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65 64 63 62 61 60 59 58 57 56 55 54 53 52 51 50 49 48 47 46 45 44 43 42 41 40 39 38 37 36 35 34 33 32 31 30 29X-15-1-00-5H

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1. IGNITION-READY CAUTION LIGHT2. ALTIMETER3. AIRSPEED INDICATOR4. ANGLE-Of-ATTACK INDICATOR5. ACCELEROMETER6. ATTITUDE INDICATOR7. AZIMUTH INDICATOR8. PITCH ANGLE SET CONTROL9. VERTICAL VELOCITY INDICATOR10. INERTIAL HEIGHT (ALTIMETER) INDICATOR11. NO.1 GENERATOR VOLTMETER12. RAS-OUT INDICATOR LIGHT13. INERTIAL SPEED (VELOCITY) INDICATOR14. NO.1 GENERATOR-OUT LIGHT15. NO.1 GENERATOR SWITCH16. EMERGENCY BATTERY SWITCH17. NO.2 GENERATOR SWITCH18. NO.2 GENERATOR-OUT LIGHT19. NO.2 GENERATOR VOLTMETER20. NO.2 APU HYDROGEN PEROXIDE OVERHEAT WARNING LIGHT21. NO.2 APU COMPARTMENT OVERHEAT CAUTION LIGHT22. NO.2 APU HYDROGEN PEROXIDE-LOW CAUTION LIGHT23. NO.2 APU SWITCH24. NO.2 HYDRAULIC TEMPERATURE GAGE25. CANOPY INTERNAL EMERGENCY JETTISON HANDLE26. STABLEPLATfORM SWITCH27. NOSE BALLISTIC ROCKET HEATER SWITCH28. VENTRAL ARMING SWITCH29. WINDSHIELD HEATER SWITCHES (2)30. fiRE-WARNING LIGHT TESTBUTTON31. INDICATOR, CAUTION, AND WARNING LIGHTSWITCH32. COCKPIT LIGHTING SWITCH33. CABIN PRESSURE ALTIMETER34. HYDRAULIC PRESSURE GAGE35. CABIN HELIUM SOURCE PRESSURE GAGE36. NO. 1 AND NO.2 BALLISTICCONTROL SWITCHES37. HYDROGEN PEROXIDE TANK PRESSURE GAGE38. APU BEARING TEMPERATURE GAGE39. NO.1 APU HYDROGEN PEROXIDE OVERHEAT WARNING LIGHT40. NO.1 APU COMPARTMENT OVERHEAT CAUTION LIGHT

In

41. MIXING CHAMBER TEMPERATURE GAGE42. APU SOURCE PRESSUREGAGE43. NO. 1 APU HYDROGEN PEROXIDE-LOW CAUTION LIGHT44. CLOCK45. NO.1 HYDRAULIC TEMPERATURE GAGE46. NO.1 APU SWITCH47. LIQUID OXYGEN BEARING TEMPERATURE GAGE48. RATE-Of-ROLL INDICATOR49. IGNITER IDLESWITCH50. H"O" COMPARTMENT-HOT CAUTION LIGHT51. CHAMBER AND STAGE 2 IGNITER PRESSURE GAGE52. TURBOPUMP IDLEBUTTON53. fUEL QUANTITY GAGE54. ENGINE PRIME SWITCH55. ENGINE PRECOOL SWITCH56. fUEL LINE-LOW CAUTION LIGHT57. PROPELLANT MANifOLD PRESSURE GAGE58. ENGINE RESETBUTTON59. ENGINE MASTER SWITCH60. PROPELLANT PUMP INLET PRESSURE GAGE61. H,O, TANK AND ENGINE CONTROL LINEPRESSURE GAGE62. PROPELLANT SOURCE PRESSURE GAGE63. PROPELLANT TANK PRESSURE GAGE64. H,O, SOURCE AND PURGE PRESSURE GAGE65. JETTISON STOP SWITCHES66. AUXILIARY LAUNCH SWITCH67. LANDING GEAR HANDLE68. VENTRAL JETTISON BUTTON69. HELIUM RELEASESELECTOR SWITCH70. f iRE-WARNING LIGHT71. ENGINE VIBRATION MALfUNCTION CAUTION LIGHT72. AMMONIA TANK PRESSURE-LOW CAUTION LIGHT73. TURBOPUMP OVERSPEED CAUTION LIGHT74. STAGE 2 IGNITION MALfUNCTION CAUTION LIGHT75. PROPELLANT EMERGENCY PRESSURIZATION SWITCH76. VALVE MALfUNCTION CAUTION LIGHT77. LIQUID OXYGEN TANK PRESSURE-LOW CAUTION LIGHT78. IDLEEND LIGHT79. NO-DROP CAUTION LIGHT

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1T.O. lX-15-1

COCKPIT LEFT SIDE

1 . UHF CONTROL PANEL2. PRESSURESUIT VENTILATION KNOB3. VENT SUIT HEATER SWITCH4. FACE MASK HEATER SWITCHS. INTERCOMMUNICATION SWITCH6. BALLISTIC CONTROL STICK7. JETTISON TRIM SWITCH8. WING flAP SWITCH

9. VENT. PRESSURIZATION, AND JrnlSON LEVER10. ENGINE THROTTLE11. UPPER SPEEDBRAKE HANDLE12. LOWER SPEED BRAKE HANDLE13. ANTENNA Se lECTOR SWITCH14 . TRIM CONTROL SWITCH1S. READY-TO-LAUNCH SWITCH

ine operation, or to move the switch to AUTO toase the helium and simultaneously shut down theine. Moving the switch to ON will release theum into the engine compartment whenever theery bus is energized. Once energized, the heliumase valve is locked in the jettison position and mustelectrically unlocked by ground personnel.INE TURBOPUMP.rbopump, mounted on the front of the engine, de-rs the propellants in the desired quantities and atproper pressures to the engine. The turbopumptains a gas generator, a two-stage, axial-flowne, and two centrifugal pumps on a common shaft.h pump supplies one of the propellants to the engine.monopropellant for driving the turbine is 90percentrogen peroxide (H2~). An electrohydraulic powero system governs turbine speed to the selecteder requirement.opump Propellant (H202) System.hydrogen peroxide monopropellant (H202) used toe the turbopump is contained in a 10-cubic-foot

Figure 1-3.

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I\X-15-1-00-6E

spherical supply tank (19, figure 1-1) with a capacityof 854 pounds (77. 5 US gallons). A swivel-type pickupfeed line allows positive feeding of the monopropellantregardless of airplane attitude. The system includesa combination vent, pressure relief, and tank pres-surization valve; a jettison valve; a hydrogen peroxidethrottle control metering valve; a safety valve; a shut-off valve; and a gas generator. This system is con-trolled by switches and a control lever in the cockpitand is put into operation whenever the engine startingsequence is begun. For a description of these controls,refer to "Engine Controls" in this section. When theengine is not operating, the tank is vented to atmosphereif the vent, pressurization, and jettison control leveris at VENT and control gas is available. The tank ispressurized with helium control gas, to feed the H2~to the gas generator, which provides steam power forturbopump operation. Tank pressure can be read froma gage in the cockpit. Refer to "H202 Tank and EngineControl Line Pressure Gage" in this section. Thesystem also includes a jettison feature that permits theH202 to be forcibly expelled overboard.


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1 . CANOPY INTERNAL HANDLE2. CABIN SOURCE HELIUM SHUTOFF VALVE SWITCH3. APU COOLING SWITCH4. CONSOLE STICK5. ALTERNATE CABIN PRESSURIZATION SWITCH6. CIRCUIT-BREAKER PANEL (TYPICAL)7 . PRESSURE COOLING LEVERB. BLOWER SWITCHES (2)

Figure 1-4.H202 Compartment-hot Light. An amber H202 com-partment-hot caution light (50, figure 1-2), on theinstrument panel, comes on when temperature in theupper area of the turbopump propellant compartmentreaches 538C (1000F) or when temperature in thelower area of the compartment reaches 427C (800F).When illuminated, the light reads "H2~ COMP HOT. "The light is powered by the primary dc bus and may betested through the indicator, caution, and warninglight test circuit.Turbopump Speed Control.An electrohydraulic servo system is used as an actuationand reference system between the turbine speed andH2~ flow. Its main components are a power package,a governor, a throttle synchro, a servo amplifier, agovernor actuator, and an H202 throttle control meteringvalve. Pressurized oil from an electrically drivenhydraulic pump is supplied to the governor and meteringvalve. When the engine throttle (throttle synchro) ismoved, the governor speed adjustment lever is set tothe desired position by the governor actuator. Thespeed of the turbopump is sensed by the governor, and

Section I

1

X-15-I-OO-7D

the hydraulic pressure balance between the governorand metering valve is adjusted to control peroxideflow into the gas generator. Decrease or increase ofthe turbopump speed from that required for the selectedthrust causes a hydraulic imbalance between governorand metering valve. As the governor reacts to restorethe hydraulic balance, hydraulic pressure to the meter-ing valve is increased or decreased, as necessary, toalter the rate of H202 flow to the gas generator andthus restore the turbopump to the desired speed.ENGINE PROPELLANT AND CONTROL SYSTEM.The two propellants, anhydrous ammonia and liquidoxygen, are routed from their respective fuel tanksto the main feed valves, which are operated by heliumpressure. From the main feed valves, the fuel isrouted to the turbopump. The fuels, pressurized by alow-pressure inert gas (helium), flow from the respec-tive tanks to the turbopump. Prime orifices allow pro-pellants to circulate and cool the engine and prime thepropellant pumps. The turbopump begins operationwhen the hydrogen peroxide supply upstream safetyvalve and downstream shutoff valves are opened. The

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CENTER PEDESTAL

X-15-1-00-I3F

Figure 1-5.

02 then flows to a gas generator, where it is con-rted to a high-pressure gas mixture of superheatedam and oxygen to drive the turbine wheel, which inn drives the propellant pumps. The propellants aresuppliedto the first-stage and second-stage ignitersd to the main thrust chamber. Mter priming is com-ed, the turbopump is operating, and the first-stageiter propellant valve is opened, the liquid oxygen tofirst-stage igniter is routed inside the turbine ex-st duct, whose hot gases heat the liquid oxygen andange it to a gaseous state. The gaseous oxygen andmonia then enter the first-stage igniter. Threerk plugs in the first-stage igniter fire the fuel andidizer mixture. When the pressure switch in thet-stage igniter is actuated, the second-stage ignitert valves open, allowing liquid oxygen and ammoniaflow into the second-stage igniter. First-stageiter flames ignite the second-stage fuel mixture.mbustion pressure in the second-stage igniter thenuates a switch which signals the main propellantlve to open. When the main propellant valve opens,l and liquid oxygen are injected into the main thrustamber, where they are ignited by second-stagemes. Before entering the main thrust chamber,

the ammonia is routed through the chamber tubes inorder to cool the main thrust chamber. Opening ofthe main propellant valve stops the flow of propellantsto the prime valves. Once engine operation has beeninitiated, thrust output is varied between 50%and 100%according to the throttle position selected by the pilot.The engine propellant control system is shown sche-matically in figure 1-7.ENGINECONTROLS.Throttle.The throttle (10, figure 1-3) controls thrust output of theengine. The throttle quadrant has three marked posi-tions: OFF, 50%, and 100%. The throttle controls anelectromechanical servo system, which includes asynchro transmitter attached to the throttle, a servoamplifier, and an actuator position transmitting syn-chro linked to the turbopump governor. Turbopumpand first- and second-stage igniter operation is accom-plished with the throttle at OFF (full aft and outboard).During the 30-second idle operation period with thethrottle OFF, the turbopump is automatically maintained

17 gI,1IJ ,.--.... .. 2 1. SAS CAUTIONLIGHTS3 2. YAR FUNCTION SWITCH16 3. YAW FUNCTION SWITCH15 4. YAW GAIN SELECTOR KNOB14 4 S. INSTRUMENTATION CONTROL PANEL13 6. COCKPITRAM-AIRKNOB

7. DCVOLTMETER5 8 . DC VOLTMETER SWITCH12 II ......---..1..i... 9. GROUND INTERPHONERECEPTACLEJ......" . 6 10. STABLE PLATFORM INSTRUMENT SWITCH11 _____ 11. RADAR BEACON SWITCH10 . '\", . 7 12. RAM-AIRLEVER13. ROLLAND YAR GAINSELECTORKNOB8 14. PITCHGAIN SELECTORKNOB15. SASTESTSWITCH9 -;:::: =,.........,'.';;.y::.-::._.,---._i';' --..


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T.O. lX-15-1at idle speed. Within 30 seconds after igniter idleoperation is begun, the throttle must be moved to50%to open the main propellant valves to the mainthrust chamber or the start sequence must be termi-nated. After main thrust chamber operation is begun,movement of the throttle between 50% and 100%willvary engine thrust accordingly.Vent, Pressurization, and Jettison Lever.This lever (9, figure 1-3) controls the pressurizationsystem selector valve. The valve is a manually con-trolled pneumatic selector valve. The lever has threepositions: VENT, PRESSURIZE, and JETTISON. Withthe lever at VENT, helium pressure (from the heliumpressure control system) is applied to all tank controlvalves in the propellant system. The pressurizationvalves close and the vent valves open, venting theH202, liquid oxygen, and ammonia tanks. In orderto obtain engine operation, the lever must be moved toPRESSURIZE, opening the propellant system pressuri-zation valves and closing the vent valves. This allowshelium to enter and pressurize the turbopump H202supply tank and the liquid oxygen and ammonia tanks.When the lever is positioned to JETTISON, helium pres-sure is applied to open three jettison valves and pres-surize the H202, liquid oxygen, and ammonia tanks.The three propellants will then begin to dump overboard.

NOTEThe propellants will not jettison if the jettisontest switches are OFF.

Jettison Trim Switch.This switch (7, figure 1-3) is on the left vertical sidepanel. It has three positions: NOSE UP, NOSE DOWN,and an unmarked, center off position. The switch ispowered by the 28-volt primary dc bus. With the vent,pressurization, and jettison lever at JETTISON, thejettison stop switches in the JETT position, and thisswitch at the unmarked off position, simultaneousjettisoning of H202, liquid oxygen, and ammonia willoccur. Moving this switch to NOSE UP (when nose-down trim is felt) stops the flow of the ammonia. Mov-ing the switch to NOSE DOWN (when nose-up trim isfelt) stops the flow of liquid oxygen. In either case,when the airplane returns to trim, the switch must bereleased and allowed to return to the unmarked offposition.Jettison Stop Switches.Three jettison stop switches (65, figure 1-2), on theinstrument panel left wing, have a STOP and a JETTposition. These switches, powered by the primarydc bus, are normally left in the STOP position untilthe prelaunch cruise portion of the flight. To performa test of the turbopump H202, liquid oxygen, and am-monia jettison system, the vent, pressurization, andjettison control lever should be placed at JETTISON.The systems then can be tested by placing the switchesto JETT. The jettison line of each system should thenemit a vaporous cloud. Flow will cease when theswitches are returned to the STOP position or whenthe vent, pressurization, and jettison control lever is

Section Imoved to PRESSURIZE. (See figure 1-15 for locationof jettison, drain, and bleed outlets.)Engine Master Switch.The engine master switch (59, figure 1-2), on the in-strument panel, is powered by the primary dc bus.With the switch at OFF, primary dc bus power forengine control and engine indicator lights is inter-rupted. With the switch at ARM, primary dc buspower is applied to the engine indicator lights andengine control switching units.Engine Reset Button.The engine reset button (58, figure 1-2), on the in-strument panel, is powered by the primary dc busthrough the engine master switch. For a normal en-gine start or if a malfunction causes automatic shut-down during any phase of operation, depressing thisbutton positions the engine control circuits to the armedposition. However, if the malfunction which causedshutdown persists, engine control circuits will not bearmed.Engine Precool Switch.The engine precool switch (55, figure 1-2), on the in-strument panel, is powered by the primary dc busthrough the engine master switch. It has two main-tained positions: PRECOOL and OFF (down). Withan engine start sequence initiated, moving the switchto PRECOOL opens the liquid oxygen main feed valveand precools the system up to the main propellant valve.The precooling flow dumps overboard through the engineliquid oxygen prime valve.

NOTEAbout 10 minutes is required to precool theengine liquid oxygen system. After precoolingis completed, the engine can be maintained ina precooled condition for an extended periodby the following schedule: engine precool switchat OFF for 20 minutes, then at PRECOOL for7-1/2 minutes, repeating this cycle as often asnecessary.

Engine Prime Switch.The engine prime switch (54, figure 1-2), on the instru-ment panel, has three positions: an unmarked, main-tained center position; a momentary PRIME position;and a maintained STOP PRIME position. With an enginestart sequence initiated, moving the switch momentarilyto PRIME opens the liquid oxygen and ammonia mainfeed valves and the turbopump H202 upstream safetyval ve and admits helium to the engine control and purgesystems. Approximately 30 seconds is required forpriming at high-flow rate, and when the engine precoolswitch is placed at OFF, prime continues at low-flowrate until an actual start stops the prime or until theengine prime switch is placed momentarily at STOPPRIME. Engine operation may be terminated duringany phase by moving this switch to STOP PRIME.

1-9


19/126

ion I T.O. 1X-15-1

1 AIRPLANE PROPELLANT AND CONTROL SYSTEM PROPELLANTSOURCEPRESSUREGAGEAIRPLANES EQUIPPED WITH XLR99 ENGINEHELIUM TANK HEL IUM F ILLER

-=- .N.O. -=- ..~....................................................._.........-.-.-.-.-.-.-.-.-.

~;~TOLIQUIDOXYGENMAIN FEEDVALVE

VENTJETTISONPOSITION VENT POSITION

VENTPRESSURIZATIONAND JmlSONLEVER

..N.O.---.;CONTROLSELECTORVALVE

(Vent and pressurizedposition shown. Openwhen at 1m/SON.)n

600 PSI

JmlSONSTOPSWITCHESJmlSON TRIMSWITCH

NOTEMove jettison trim switch to NOSE DOWN tostop liquid oxygen jettison. Move switch toNOSE UP to stop ammonia jettison.

X-15-1-48-5B

Figure 1-6


20/126

T. o. 1X-15-1 Section I

-..

N.O. 48 PSI

HELIUM TANK3600 PSI

.' -PROPEllANT EMERGENCY

PRESSURIZATIONSWITCH AND LIGHTS

HELIUMRelEASESWITCH

HELIUMFILLER

JEniSON............. -- ......... "

-Q-TOENGINECONTROL

EMERGENCYJrnl SON VALVE

ALTERNATESOURCE3600 PSI ..

TOAMMONIAMAIN fEED

VALVENOTE

When the vent, pressurization, and jettison control lever is in theVENr position, all vent valves are open, and pressurizationand jettison valves are closed. When the lever is in the PRES-SURIZED position, the pressurization valves are open, andthe jettison and vent valves are closed. When the lever is inthe JEr1lS0N position, the jettison and pressurization valvesare open, and the vent valves are closed. X-15-1-48-6B

1-11

_ AMMONIA CHECKVALVE - FILLEROR OVERFLOWDRAINLIQUID OXYGEN --1 SHUTOFFVALVE........ HELIUM (ELECTRICAllYACTUATED) PRESSURE REGULA TORHYDROGENPEROXIDE SHUTOFF VALVE==== NONPRESSURIZEDLINE (PNEUMATICALLYACTUATED) .r'...- elECTRICALCONNECTION SHUnLE VALVEa PRESSURETRANSMlnER 1.,.6-MECHANICALLINKAGE .. ,N.O. NORMAllY OPEN . PILOTVALVE AIIIIiI.i.. flOW RESTRICTORN.C. NORMAllY CLOSED (ElECTRICAllY ACTUATED) "I - -- flOW LIMITINGORIFICEROPELLANT TANK

PRESSURE GAGE


21/126

ction I

1T.O. 1X-15-1

FROM ALLMALFUNCTION

SIGNALSII

ENGINE PROPELLANT CONTROL SYSTEM FROMHYDROGENPEROXIDETANK

NOTELiquidoxygen, ammonia, and hydrogen peroxide are suppliedto the feed valves under pressure when the vent, pressuriza-t ion, and jet tison control lever is placed at PRESSURIZE.

H202 TANK ANDENGINECONTROLLINEPRESSUREGAGE"

II~O!O

TOTURBOPUMPLUBE PUMP

_ AMMONIA~ LIQUIDXYGEN

. HELIUM~ HYDROGENEROXIDE. . . . HYDRAULIC== STEAM" " "" < , GASEOUSOXYGEN- ELECTRICALONNECTION

N. O. NORMALLY OPENN. C. NORMALLY CLOSED

PILOT VALVE(ELECTRICALLY ACTUATED I

X-15-1-48-7A

:M:' 0v ~ -----

o

I I I\" .

TOSPARK PLUGS .N.C.23-SECOND

TIMER ".

ENERGIZED AFTER23 SECONDS ORAT LAUNCH

..C.

Figure 1-7

MECHANICALLINKAGESHUTOFF VALVIiIPNEUMATICALL Y ACTUATED IC PRESSURETRANSMITTERPRESSURE SWITCH

m SHUTTLE VALVEI VENTURI[I] ORIFICErD ACTUATOR

CHECK VALVE

.N.C., f_( ..I r:=j jH ----".

---0\C?OFIRESWITCH, f "-'"

.N.C.


22/126

T. O. 1X-15-1 Section I

MAIN FEEDSAFETY VALVEN.C.

t

It> '.N.C.

w__u..&.&MALFUNCTION

SIGNAL NOTENomenclature on caution lights shownilluminated for information only.-15-t-48-8C

1-13


23/126

ction I T. O. 1X-15-1

rbopump Idle Button.he turbopump idle button (52, figure 1-2), on the in-rument panel, is powered by the primary dc busrough the engine master switch and the engine over-eed reset button. With an engine start sequenceitiated and the prime phase completed, depressingis button for one second opens the turbopump H202wnstream shutoff valve, which starts turbopumperation. Whenpressure of the ammonia in the pro-llant manifold builds to approximately 210 psi, therbopump speed control system begins operation andintains the turbopump at idle speed.niter Idle Switch.his switch (49, figure 1-2), on the instrument panel,ceives power from the primary dc bus through thegine master switch and engine overspeed reset but-n. With an engine start sequence initiated and thegine turbopump operating at idle speed, moving thisitch from OFF (down) to IGNITER causes the follow-g sequence of actions: a 2-second helium purge isitiated, and the spark plugs are energized; the first-age igniter propellant valves open, and igniter andle timing starts; gaseous nitrogen flow (from therrier airplane) starts when first-stage igniter pres-re reaches a specified value; propellants flow to thecond-stage igniter; and second-stage ignition occurs.

I WARNING ~This phase of operation (igniter idle) is limitedto 30 seconds. Either igniter idle operationmust be terminated (by placing the engine primeswitch to STOP PRIME) or the launch accom-plished at the end of the 30-second idle period.GINE INDICATORS.opellant Tank Pressure Gage.is dual-indicating gage (63, figure 1-2), on the instru-ent panel, is powered by the 26-volt ac bus. The gagedjcates the two propellant tank pressures. The gage isaduated from 0 to 100 psi in increments of 5 psi. Oneinter of the gage has the letter "L, " indicating theuid oxygen tank pressure; the other pointer has thetter "A, " indicating the ammonia tank pressure.opellant Pump Inlet Pressure Gage.is dual-indicating gage (60, figure 1-2) is poweredthe 26-volt ac bus. It indicates liquid oxygen andmonia pressures at the engine turbopump inlets.is gage is graduated from 0 to 100 psi in increments25 psi. The pointer labeled "L" reads the liquidygen pressure; the pointer labeled "A" indicatesammonia pressure.quid Oxygen Bearing Temperature Gage.is gage (47, figure 1-2) is powered by the 26-voltbus and indicates the temperature of the bearingr the liquid oxygen centrifugal pump segment of the

engine turbopump. The gage is graduated from _600 Fto 2600 Fin 5-degree increments.Propellant Manifold Pressure Gage.This dual-indicating gage (57, figure 1-2), on the in-strument panel, is powered by the 26-volt ac bus. Thegage indicates, propellant pump discharge pressures.It is graduated from 0 to 2000 psi in increments of 50psi. One pointer is labeled "L" and indicates liquidoxygen pump discharge pressure; the other pointer Lslabeled "A" and indicates ammonia pump dischargepressure.Chamber and Stage 2 Igniter Pressure Gage.This dual-indicating gage (51, figure 1-2), on the in-strument panel, is powered by the 26-volt ac bus. Thegage is graduated from 0 to 1000 psi in increments of20 psi from 0 to 100, and 50 psi from 100 to 1000. Theshort hand indicates pressure in the second-stage ig-niter. The long hand indicates pressure in the mainthrust chamber.Ignition-ready Indicator Light.This green indicator light (1, figure 1-2), on the instru-ment panel, is powered by the primary dc bus throughthe engine master switch. When illuminated, it reads"IGN READY, " indicating that the engine electricalcircuits and purge gas network have been energized.In the normal starting sequence, this light will go outfor 2 seconds when the igniter idle switch is moved toIGNITER, then come on again. During all heliumpurges, this light will go out. This light may be testedby the indicator, caution, and warning light test circuit.No-drop Caution Light.This amber caution light (79, figure 1-2) on the instru-ment panel, is powered by the primary dc bus throughthe engine master switch. During a normal engine startsequence, this light will come on when 7 seconds re-mains in the igniter idle phase of operation. The light,whenilluminated, reads "NODROP" and serves to warnthe pilot to terminate igniter idle operation or to con-tinue on to the launch phase. This light may be testedby the indicator, caution, and warning light test circuit.Idle-end Caution Light.This amber caution light (78, figure 1-2), on the instru-ment panel, is powered by the primary dc bus throughthe engine master switch. This light will illuminate,reading "IDLE END, " when the 30-second igniter idlephase of engine operation is complete. Whenthis lightcomes on, engine shutdown must be accomplished oroperation continued into the main chamber phase (afterlaunch). This light may be tested by the indicator,caution, and warning light test circuit.Valve Malfunction Caution Light.This amber caution light (76, figure 1-2), on the instru-ment panel, is powered by the primary dc bus throughthe engine master switch. When illuminated, this lightreads "VALVE MAL." This light will come on when


24/126

T. O. 1X-15-1malfWlctionshutdownoccurs because the main or first-stage propellant valve is improperly positioned duringthe starting sequence. The light will also come on mo-mentarily whenever a malfWlction shutdown occurs withthe main chamber operating. This light may be testedby the indicator, caution, and warning light test circuit.Stage 2 Ignition Malfunction Caution Light.This amber caution light (74, figure 1-2), on the instru-ment panel, is powered by the primary dc bus throughthe engine master switch. Whenilluminated, this lightreads fIST2 IGNMAL." The light will come on whena malfunction shutdown occurs during the startingsequence because of failure of the second-stage igniterto reach operating pressure. The light will also comeon momentarily whenever a malfunction shutdown occurswith the main chamber operating. This light may betested by the indicator, caution, and warning light testcircuit.Vibration Malfunction Caution Light.This amber light (71, figure 1-2), on the instrumentpanel, comes on when the engine shuts downbecause ofexcessive vibration. Excessive engine vibration causesa shutdown signal to be transmitted from either of twosensors to the engine control box. The signal causesthe actuation of two malfW1ctionrelays in the control boxthat de-energize the prime, precool, and firing circuitsto shut downthe engine and turn on the light. The lightis powered by the primary dc bus. If the light comeson during powered flight, an engine restart may beattempted.Turbopump Overspeed Caution Light.This amber caution light (73, figure 1-2), on the instru-ment panel, is powered by the primary dc bus throughthe engine master switch. When illuminated, this lightreads "PUMP O'SPD." The light will come on if a mal-function shutdown occurs because of turbopump over-speed whi:h is not corrected by the turbopump governor.This light may be tested by the indicator, caution, andwarning light test circuit.

Fuel Quantity Gage.A fuel quantity gage (53, figure 1-2), on the instrumentpanel, is calibrated in percent of total fuel load. Thegage is graduated from 0 to 100 in increments of 5 andoperates on the principle that chamber pressure in theengine is essentially proportional to the flow of fuel andoxidizer to the engine. During engine operation, thepointer moves counterclockwise on the dial at a rateproportional to the engine thrust selected (which alsois proportional to the rate of fuel consumed). Beforean engine start is initiated, the pointer is adjusted (bythe center reset knob) to indicate the percent of totalfuel on board During engine operation, the gage indi-cates the percentage of fuel load remaining. The fuelquantity gage receives a signal from the fuel quantitycontrol Wlit, which is powered by the primary dc andNo. 2 primary ac busses.

Section INOTE

The fuel quantity indicating system can be ad-justed for an accurate indication at burnouttime by selection of the proper start point withthe reset knob. This start point is determinedby calibration of the system to the airplane inwhich it is installed.Fuel Line Low Caution Light.An amber "FUEL LINE LOW" caution light (56, figure1-2) is on the left side of the instrument panel. Thislight, powered by the primary de bus through the enginemaster switch, is actuated by a pressure switch in-stalled in the fuel (ammonia) line downstream of the mainsafety valve. If fuel pressure at the turbopump inletdrops to 32 (:1:2)psi, the light will come on. Illumina-tion of the light indicates that partial cavitation of thepump is likely to occur.

When the "FUEL LINE LOW" caution light ison, there is not sufficient cooling around themain chamber of the engine for temperatureprotection. Thrust settings above 50% may in-crease the temperature and cause damage tothe engine.If the light comes on before the engine is started, thestart will be aborted. (Refer to "Fuel Line PressureLow" in Section ill.) When this light has been illumi-nated, it will remain on until the engine master switchhas been placed in the OFF position. The light may betested by the indicator, caution, and warning light testswitch.

PROPELLANT SUPPLY SYSTEM.The propellant supply system consists of the liquidoxygen supply (oxidizer), anhydrous ammonia (fuel)supply, valves, and associated plumbing. The liquidoxygen and ammonia are fed Wlder a low inert gas pres-sure from the supply tanks to the turbopump for engineoperation. The helium supply systems, which furnishgas pressure for tank pressurization, pneumatic valveoperation, and system purging, are described in sepa-rate paragraphs in this section. See figure 1-16 forthe liquid oxygen and ammonia specifications.LIQUID OXYGEN TANK.The liquid oxygen supply is carried in a triple-compartmented tank (12, figure 1-1), just aft of theNo. 2 equipment compartment. The center sectionarea of the cylindrical tank is hollow and forms a casefor a gaseous helium high-pressure storage tank. Whenthe liquid oxygen tank is not Wlder pressurization, it isvented to atmosphere. The tank compartments arecheck-valve-vented. Each compartment feeds rear-ward toward the airplane center of gravity. The liquidoxygen is fed from the rear compartment under 48 psiof helium pressure to the turbopump or jettison linethrough a series of control valves. The total volume

1-15


25/126


the tank is 1034 US gallons; of this amount, 14 gallonsresidual at a liquid surface angle of 38 degrees, andgallons is vent and expansion space. The total usableuid oxygen is 1003 gallons. The tank is filled forght through the carrier airplane's supply system.e tank incorporates a liquid oxygen fluid level sensingitch that permits the tank to be topped off automat-ally whenever fluid drops below a predetermined level.r ground operational checks, the tank is servicedrough the receptacle mounted on the engine feed line.e tank filler is on the topside of the wing fairing tun-l forward of the left wing root leading edge.MONIA TANK.

e ammonia supply is carried in a triple-compart-ented cylindrical tank (17, figure 1-1), just aft ofe No.3 equipment compartment and ahead of therbopump hydrogen peroxide tank. The center sectionea is hollow and closed at both ends. The rear com-rtment center section is perforated to allow storageammonia within the center section area. The com-rtments are check-valve-vented. This aids in theessure feed of the fluid transfer from the rear tankmpartment forward toward the airplane center ofavity. The rear compartment empties first; then theddle compu1:ment empties into the front compartment,ith the ammonia fed from the front compartment underpsi of helium pressure to the turbopump or jettisone through a series of control valves. The total vol-e of the tank is approximately 1445 US gallons. Thek is ground-serviced only. The filler receptacle fore tank is on the underside of the right wing root fairing

GINE AND PROPELLANT CONTROL HELIUM

elium to pressurize the turbopump hydrogen peroxidepply tank and to supply pneumatic pressure for engined propellant control is contained in four sphericalnks. One tank (15, figure 1-1) is between the liquidygen and ammonia tanks. Two tanks (23, figure 1-1)e in the left and right wing root fairing tunnelstboard of the engine. These three tanks are inter-nnected, supplying 3600 psig pressure to twoessure-reducing regulators in parallel. The fourthk is just to the right of the turbopump H202 supplyk and supplies helium at 3600 psig to a singleessure-reducing regulator for emergency or second-y pneumatic control of the propellant jettison valves.is tank is interconnected with the other three tanksr filling purposes only. From the parallel pressure-ducing regulators of the main supply, helium at 575600 psig is supplied to the engine helium manifoldr operation of engine control valves, to the turbopump202 supply tank for tank pressurization, and to pro-llant control jettison and main feed valves. Two ofe tanks supply helium directly to the helium dumplve, for engine compartment purging. The dumplve is solenoid-operated and controlled by the heliumlease selector switch. For information on operationthis switch, refer to "Helium Release Selector Switch"this section. The helium to the engine helium mani-ld is in turn routed to a control gas valve and the two

gas regulators in the purge valve network at a pressureof 550 to 600 psig. The control gas valve is energizedduring the prime period and admits helium at a pressureof 550to 600 psig to the pilot valves for the prime valve,first-stage igniter start valve, second-stage igniterstart and shutoff valves, and main propellant valve.Helium at a pressure of 125 to 200 psig is routed fromthe two purge gas regulators and to the return side ofthe second-stage igniter start and.shutoff valves. An-other regulator supplies helium from the helium mani-fold at 7.5psig to the lubrication system accumulator,engine control box, and hydraulic power package.HELIUM RELEASE SELECTOR SWITCH.Refer to "Engine Compartment Purging System" in thissection.

H202 SOURCE AND PURGE PRESSURE GAGE.A dual-indicating H 0 source and purge pressure gage(64, figure 1-2), on\h~ instrument panel, is powered bythe 26-volt ac bus. This gage indicates the helium pres-sure available from three of the engine and propellanthelium system tanks. Needle 1 indicates pressure inthe large tank between the liquid oxygen and ammoniatanks. Needle 2 indicates pressure in the two smallertanks in the wing root fairing tunnels. The gage iscalibrated from 0 to 4000 psi in increments of 100 psi.Normally, the twopointers will indicate the same pres-sure. However, if there is a malfunction of the emer-gency jettison system helium supply or if helium isdumped into the engine compartment, the pointers willnot indicate the same pressure. There is no gage inthe cockpit which indicates pressure in the emergencyjettison system helium supply tank.H2~ TANKANDENGINECONTROLLINE PRESSUREGAGE.This dual-indicating gage (61, figure 1-2), on the in-strument panel, is powered by the 26-volt ac bus. Onepointer, labeled "C, " indicates engine control line(helium) pressure downstream of the two parallel pres-sure regulators. The other pointer, labeled "T, " indi-cates pressure in the turbopump H2~ supply tank. Thegage is calibrated from 0 to 1000 psi in increments of50 psi, except that the range 0 to 100 is in incrementsof 20 psi.PROPELLANT PRESSURIZATION HELIUM SYSTEM.The propellant pressurization helium system suppliesgas to pressurize the liquid oxygen and ammonia tanks.This helium is contained in the supply tank (14, figure1-1) within the center section of the liquid oxygen tankand is pressurized to 3600 psi. The helium flows tothe normally open pressure regulators of the liquidoxygen and ammonia supply tanks. The two regulatorsare actuated by helium pressure (from the engine andpropellant helium control system) to the closed positionwhen the vent, pressurization, and jettison control leveris at VENT. When the control lever is placed at PRES-SURIZE or JETTISON, the regulators open and heliumpressure flows to the liquid oxygen and ammonia tanks.


26/126

T. O. 1X-15-1The regulators reduce the helium pressure to 48 psibefore it enters the liquid oxygen and ammonia tanks.when the liquid oxygen and ammonia tanks are pres-surized, the propellants are forced through the feedlines to the main feed shutoff valves.PROPELLANT SOURCE PRESSURE GAGE.The propellant source pressure gage (62, figure 1-2),on the instrument panel, is powered by the 26-volt acbus. The gage indicates pressure in the cylindricalhelium tank for liquid oxygen and ammonia tank pres-surization. The gage is calibrated from 0 to 4000 psiin increments of 100 psi.PROPELLANT EMERGENCY PRESSURIZATIONSYSTEM.The propellant emergency pressurization system canbe used to pressurize either the liquid oxygen or theammonia tank in case of a failure in the normal pres-surization system. This will per m i t continued lowthrust engine operation or pro~llant jettisoning. Thee mer g en c y system can supply pressurizing gas toonly one propellant tank at a time. The emergencysystem uses he Ii u m from the three interconnectedtanks in the engine and propellant cmtrol helium sys-tem. The system includes a switch and two cautionlights.Propellant Emergency Pressurization Switch.This three-position switch (75, figure 1-2), on the in-strument panel, controls primary dc bus power to thetwo emergency pressurization system solenoid-operatedshut-off valves. With the switch at OFF, the valves arede-energized closed. The switch must be pulled straightout of a detent to move it from OFF to either of the otherpositions. With the switch at LOX, electrical power isapplied to open the shutoff valve which controls emer-gency helium pressure to the liquid oxygen tank. Withthe switch at NH3, electrical power is applied to openthe shutoff valve which controls emergency heliumpressure to the ammonia tank. All three switch posi-tions are maintained.Liquid Oxygen and Anunonia Tank Pressure-low CautionLights.These lights (72 and 77, figure 1-2), on the instrumentpanel, are powered by the primary dc bus. The liquidoxygen tank pressure-low caution light is labeled "LOX. "The ammonia tank pressure-low caution light is labeled"NH3'" (The nomenclature for the lights also serves asposition nomenclature for the propellant emergencypressurization switch.) Mter the vent, pressurization,and jettison lever is placed at PRESSURIZE, the relatedlight will come on when pressure in the affected tankdrops to 34 (:1:2)psi. If a light comes on during poweredflight, it may remain on even after emergency pressur-ization of the affected tank has been initiated, indicatingthat the affected tank pressure is not above 40 psi.

NOTEDuring the transitional period when the vent,pressurization, and jettison lever is movedfrom VENT to PRESSURIZE, the lights shouldcome on and remain on for approximately 6seconds (during build-up of pressure in thepropellant tanks).

Section I

AUXILIARY POWER UNITS.The airplane is equipped with two auxiliary power units(7, figure 1-1) that are set side-by-side in a compart-ment in the forward fuselage. Each unit is a completelyautomatic, constant-speed, turbine drive machine thattransmits power to, and provides structural support for,an ac generator and a hydraulic pump. Propellant foreach auxiliary power unit is provided by an independentfeed system, using helium pressure to move the mono-propellant, hydrogen peroxide. The two auxiliary powerunits with their respective feed systems are identifiedas system No. 1 and system No.2. Their operation iscompletely independent of each other, and each furnishesone half of the power required. If one unit should fall,the other will provide sufficient electrical and hydraulicpower for limited flight capabilities. Each auxiliarypower unit is started and stopped by a switch in thecockpit. When an APU is turned on, a solenoid-typeshutoff valve is opened to allow hydrogen peroxide fromthe propellant feed system to flow into the unit. Thepropellant is routed first through a gear case for coolingpurposes (nitrogen gas is also introduced into the upperturbine bearing area for additional cooling) and then to amodulatingflow control valve. The flow control valve ismodulated to open or close to provide stabilizationthrough a speed control system consisting of a tachom-eter generator and a frequency detector. Any turbineoverspeed condition is sensed by an overspeed sensingelement in the speed control system which will auto-matically act to close the shutoff valve. Whenthe shut-off valve is closed, fuel flow stops and the unit shutsdown. The APU shutoff valve is fitted with a drain thatopens when the valve is closed, to relieve any excesspressure in the line downstream of the shutoff valve.After passing through the flow control valve, the hydro-gen peroxide enters a decomposition chamber containinga catalyst bed. This catalyst bed is made up of a seriesof silver and stainless-steel screens which act to de-compose the hydrogen peroxide into a high-pressure gasmixture of superheated steam and oxygen. The de-composition chamber is heated electrically from thecarrier airplane to ensure a fast start under "cold-soak" conditions in case of an emergency. The super-heated steam and oxygen mixture enters a nozzle boxin the turbine housing. Here, five nozzles convertpressure energy of the fluid into kinetic energy anddirect the flow of gas against a turbine wheel. Theturbine, acting through a reduction gear train, trans-mits power to the ac generator and hydraulic pump.The turbine wheel is housed within an exhaust casingwhich is designed to contain any buckets that mightseparate from the wheel during an overspeed operation.The exhaust casing collects spent gases that have passedthrough the turbine wheel and exhausts them overboard.A gear casing assembly contains the reduction gearing,accessory drive pads, cooling passages, provisions forlubrication, and a drive for the tachometer generator.A typical auxiliary power unit and its propellant feedsystem are shown schematically in figure 1-8. Forinformation on nitrogen cooling of the upper turbinebearing of each APU, refer to "APU Cooling Switch"in Section IV.APU SPEED CONTROL.The speed control for each auxiliary power unit providespositive speed control, starting and stopping, and

1-17


27/126


1 ."AUXILIARY POWER UNIT AND BALLISTIC

NOTE.Only the No 1 supply system and No 1 APUare shown; the No.2 supply system and No.2 APU are identical.HYDROGENPEROXIDE-LOWCAUTION LIGHT

NO.1 BALLISTICCONTROLSWITCH

.Nomenclature on indicator lights shownil luminated for information only.

FILLER VENT(Provides ventingduring filling) REFILLING

RECEPTACLEPRESSUREDIFFERENTIALSWITCH

.TO

COMMON LINEOVERBOARD

FILLER VENT(Provides ventingduring filling)HELIUM PRESSUREGAGE (SERVICING)

DIFFERENTIAL PRESSUREVALVE

POSITIVEEXPULSIONOUTLET

BAFFLE CYLINDER

PICKUPTUBE-

REFILLINGRECEPTACLE

BLADDER-

STORAGETANK

H,O, HOT

APU SOURCE PRESSUREGAGE(Commono bothsystems)APU HYDROGEN PEROXIDE TANK

PRESSURE GAGE(Common to both sys tems)

HYDROGEN PEROXIDE OVERHEATWARNING LIGHT

X-15-1-49-IA

Figure 1-8 (Sheet 1 of 3)


28/126

T. O. 1X-15-1 Section I

CONTROL AND REACTION AUGMENTATION SYSTEMS

NO.1APUSWITCHAPU BEARING

TEMPERATURE GAGE

I"",,m....."~r-APUOMPARTMENT

OVERHEATWA RN IN G L IG HT

APU COMP HOT

TOCOMMON L INEOVERBOARD.0

-CONTROLLER

TURBINE HYDRAULICPUMP APU-DRIVENACCESSORYCOMPARTMENT

-BLOWOUT

PLUG

IIml COOLINGNITROGENINLETAPU SHUTOFFVALVE

JEmSON ANDBALLISTICCONTROLVALVE

DRAIN, TOI) COMMONLINE, OVERBOARD

Q-TOBAlliSTICCONTROLANDREACTIONAUGMENTATIONSYSTEMS

HELIUM GASHYDROGEN PEROXIDESUPERHEATEDSTEAMAND OXYGENNITROGEN GASVENT AND JmlSONCHECKVALVE

[!]I. THERMOSWITCHPRESSURE RELIEF VALVEPRESSURE TRANSMITTERELECTRICAL CONNECTIONM EC HAN IC AL LI NKAG EX-I5-1-52-3C

Figure 1- 8 (Sheet 2 of 3) 1-19


29/126


1

IIIIIIII . ..I~~.~, ~ : ~ _~/__..I~~Ir ~ ~""'I ~ ~~;ROLI STICKII

~I~,1/. YAWAND PITCH'( METERINGVALVE1'----1

ROLLSOLENOIDVALVE 1/ TONO. I SYSTEMROll ROCKET1/ MOTORS

fROM NO. 1APUANDBAllIST~CONTROLSWITCHES

TO1/ NO. I SYSTEMYAW AND PITCH1/ ROCKET1/ MOTORS

JmlSON ANDBALLISTIC

CONTROL VALVE TOCOMMON LINEOVERBOARD

II II II0.0-.0-.0-fROM RAS SYSTEM(No. I systemonly)

QIT

RAS CONTROLPANELRA'

PITCH ROlL YAW ACCHr- ENGAGE ---, AUTO~)~)(~)~)L.:.. STANDBY - -= .. J OFF

RAS GYROACCELEROMETER

UNIT

RAS-OUT INDICATORLIGHT

LEfT-YAW ROCKET- NOSE-UP PITCHROCKET--- ~RIGHT YAW ROCKET

NOSE-DOWN PITCHROCKET

LEf T ROLLROCKET---RIGHT ROLLROCKET

X-15-1-52-6

Figure 1-8 (Sheet 3 of 3)


30/126

T. O. lX-15-1 Section Ioverspeed protection by regulating the flow of hydrogenperoxide to the decomposition chamber. The key com-ponent of each speed control is a controller which con-tains the necessary circuits for sensing unit operationthrough the frequency output of a tachometer generator.During normal operation, the frequency of the powergenerated by the tachometer generator, being propor-tional to the turbine speed, matches a preset frequencyof the controller. Any change in turbine speed due to achange in load, or from any other disturbance, causes aproportional change in frequency of the tachometer gen-erator. This frequency change is sensed by the control-ler, which in turn adjusts the opening of the flow controlvalve to bring turbine speed back to the normal operatinglevel. During normal operation, the speed of the unit isautomatically controlled to maintain 51,200 rpm by thespeed-sensing element of the controller. Should anoverspeed condition occur (56,000 rpm or greater), theoverspeed sensing element of the controller automat-ically acts to energize the solenoid-operated APU shutoffvalve to the closed position, thereby shutting off the flowof the propellant. The unit then decelerates and stops.It cannot be restarted until the APU switch is first cycledto the OFF position.APU SWITCHES.There are two APU switches (23 and 46, figure 1-2) onthe instrument panel, one for control of each auxiliarypower unit and its associated feed system. When eitherswitch is turned to ON, battery-bus power is used toopen the helium shutoff valve in the related propellantfeed system, allowing helium pressure to move thehydrogen peroxide through the feed lines. At the sametime, power is applied to the opening circuit of the re-lated APU shutoff valve. This permits the propellantto flow to the auxiliary power unit. Turning the switchto OFF closes the helium shutoff valve (if the ballisticcontrol swftch is at OFF), shutting off the helium supply.

NOTEIf the ballistic control switch is at ON, the OFFposition of the APU switch will not close thehelium shutoff valve.

At the same time, the APU shutoff valve closes, shuttingoff the flow of the propellant to the unit. The JETTposition, powered by the primary dc bus, is used if anemergency arises and it is desired to jettison the pro-pellant overboard. The switch is guarded to prevent itfrom being accidentally moved to the JETT position.Whenthe switch is turned to JETT, the following occurs:The helium shutoff valve in the feed system opens, orremains open if it is already so, allowing the helium tocontinue to force the hydrogen peroxide through the feedlines. Concurrently, the APU shutoff valve closes anda jettison and ballistic control valve in the feed systemopens to the jettison position. The jettison and ballisticcontrol valve serves both as a shutoff valve for propel-lant supply to the ballistic control system and as a pro-pellant jettison control. As this valve opens to thejettison port, the propellant is routed through a linethat dumps overboard at the aft end of the airplane.

APU COMPARTMENT OVERHEAT CAUTION LIGHTS.Two APU compartment amber overheat caution lights(21 and 40, figure 1-2) are adjacent to their relatedAPU switches on the instrument panel. The lights arepowered by the battery bus. A thermoswitch in eachAPU accessory drive compartment is set to energizethe light when the temperature in the compartment risesto approximately 5250 F. The lights read "APU COMPHOT" when on. If either light comes on, the relatedauxiliary power unit should be shut down immediately.APU AND BALLISTIC CONTROL PROPELLANT FEEDSYSTEMS.Two completely independent feed systems provide pro-pellant for the auxiliary power units and the ballisticcontrol system. (Refer to "Ballistic Control System"in this section.) System 1 is in the left side of the fuse-lage; system 2 is in the right side of the fuselage. Thesystems are identical. Helium gas under pressuremoves the monopropellant hydrogen peroxide to itsbasic function of providing fuel to these units at therequired flow rates and pressures. Each propellantfeed system includes a high-pressure, spherical, he-lium storage tank and a positive expulsion-type hydro-gen peroxide storage tank. Both tanks are below therelated auxiliary power unit in the forward fuselage.Helium and hydrogen peroxide filler valves and heliumhigh-pressure gages for ground servicing are in eachside of the fuselage side fairings. Helium and hydro-gen peroxide pressure gages common to both systemsare in the cockpit. When the ballistic control switch isturned ON, or when the APU switch is turned ON (or toJETT), a shutoff valve is opened to allow helium to flowfrom the storage tank. The helium tank contains enoughhelium to expel all the hydrogen peroxide in the hydrogenperoxide storage tank. Helium pressure is reduced from3600 psi at the tank to 550 psi as it passes through apressure regulator. A relief valve upstream of thepressure regulator prevents overpressurization dueto overcharging or pressure build-up during high-temperature conditions. From the pressure regulator,the helium passes through the shutoff valve and pres-surizes the hydrogen peroxide tank. The positive-expulsion type hydrogen peroxide tank contains a bafflecylinder, perforated to allow the propellant to flow to apickup tube inside the baffle cylinder. The inlet of thepickup tube is very close to the bottom of the tank toprevent it from being uncovered during normal flightattitudes when only approximately 20 percent of thepropellant supply remains in the tank. Between thebaffle cylinder and tank wall is a collapsible plasticbladder. The helium enters the tank between the walland the bladder where pressure on the bladder forcesthe hydrogen peroxide into the baffle cylinder throughthe pickup tube and out of the tank. A check valve up-stream of the tank prevents hydrogen peroxide frombacking into the helium system in case of a bladderfailure. When the tank is emptied to the extent thatthe bladder collapses against the baffle cylinder, thefeed pressure will drop off. This pressure drop cre-ates a pressure differential between the helium andhydrogen peroxide. When this pressure differentialincreases to approximately 35 psi, a differential pres-sure switch in the system actuates a low-level cautionlight in the cockpit. Pressure differential is also sensed

1-21


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ction I T.O. 1X-15-1 -------.-a valve that opens at approximately 55 psi and allowslium to flow to the top of the baffle cylinder, expellingremaining hydrogen peroxide. Protection of thedrogen peroxide tank against rupture due to overpres-rization is provided by a pressure relief valve and awout plug. The relief valve is designed to open atproximately 650 psi (100 psi over normal tank pres-re). In case of a malfunction of the pressure relieflve or an abnormal rate of pressure increase, thewout plug will rupture at approximately 900 psi. Ifblowout plug should rupture, the affected systemll be deactivated by loss of the propellant through theg. The flow of both of these relief devices is routederboard through a vent and jettison line. A thermo-itch at the base of the tank energizes a warning lightthe cockpit if the hydrogen peroxide in the tank be-mes overheated. From the storage tank, the propel-t is routed into feed lines through shutoff valves toauxiliary power unit and the ballistic control system.mperature of the propellant at the APU inlet must be animum of 40 F during starting. To prevent freezingthe feed lines, warm air from the carrier airplane ismped into the compartment containing the propellantd system hydrogen peroxide components to maintainemperature of approximately 120 F. The system isigned to dump helium and hydrogen peroxide over-rd if an emergency arises.U and Ballistic Control Propellant Feed Systemntrols and Indicators.U Source Pressure Gage. A dual-movement heliumssure gage (42, figure 1-2), common to both propel-t feed systems, is on the right side of the instrumentel. The gage, labeled "SOURCE, " is calibrated innds per square inch and has two pointers, markedand "2" for system identification. The gage includeslip ring, with calibration markings of "F, " "3/4, "2," "1/4, " and "E." The slip ring is used to indicateamount of H2~ available to the APU's. The quantityH202. available from the tanks is proportional torce (helium) pressure. Therefore, the slip ringuld be rotated just before APU start so that the "F"rk is aligned with the No. 1 pointer; then the positionthe No. 1 pointer in relation to the slip ring calibra-s will indicate the amount of H2~ available for APUration once they are started. The pressure indicatingtem is powered by the 26-volt ac bus. Operatingssure will vary from 3600 down to 550 psi, dependingn the helium supply in the storage tank.U Hydrogen Peroxide Tank Pressure Gage. A dual-vement hydrogen peroxide pressure gage (37, figure), common to both propellant feed systems, is onright side of the instrument panel. The gage iseled "H2~" and is calibrated in pounds per squareh. The pointers are marked "1" and "2" for systemtification. The gage shows tank pressure in therogen peroxide storage tanks, sensed by a pressuresmitter in each feed system. Normal operatingsure is approximately 550 psi. The pressure indi-ng system is powered by the 26-volt ac bus.Bearing Temperature Gage. A dual-pointer APUring temperature gage (38, figure 1-2) is on the in-ment panel. The gage shows in degrees centigradetemperature of No. 1 and No. 2 APU upper turbine

bearings. The temperature indicating system is pow-ered by the No. 1 primary ac bus. The gage is cali-brated from zero to 200 in increments of 20 degrees.The left pointer indicates No. 1 APU upper turbinebearing temperature; the right pointer, No. 2 APUupper turbine bearing temperature.Hydrogen Peroxide-low Caution Lights. Each propellantfeed system has a low-level caution light (22 and 43,figure 1-2) that reads "H202 LOW" when on. The twoamber, placard-type lights, on the right side of the in-strument panel, are powered by the primary dc bus. Apressure-differential switch in each system becomesenergized, causing the related light to come on whenthe differential pressure between helium and hydrogenperoxide rises to approximately 35 psi. The light willcome on at this instant, when approximately 20 percentof the hydrogen peroxide supply is left in the storagetank. If either light comes on, extreme maneuversshould be avoided to prevent uncovering the inlet of thepickup tube in the tank, thus allowing helium to flow intothe hydrogen peroxide line.Hydrogen Peroxide Overheat Warning Lights. The No. 1and No. 2 hydrogen peroxide overheat warning lights (20and 39, figure 1-2) are on the right side of the instru-ment panel. The red, placard-type lights are poweredby the battery bus and read "H202 HOT" when on. Athermoswitch at the base of each system hydrogen per-oxide storage tank energizes the related light if thetemperature of the contents of the tank rises to approxi-mately 160 F. If either light comes on, the contents ofthe affected tank should be jettisoned. Concurrently, therelated auxiliary power unit will automatically shut down.APU Switches. Refer to "Auxiliary Power Units" in thissection.Ballistic Control System Switches. Refer to "BallisticControl System" in this section.ELECTRICAL POWER SUPPLY SYSTEMS.The airplane is equipped with an alternating-currentand a direct-current electrical power system. (Seefigure 1-9.) Power for the ac system is supplied bytwo alternator-type generators. The dc system normallyis powered from the ac system through two transformer-rectifiers. A 24-volt battery is available for use in anemergency to supply dc power to essential equipment.During ground operation, ac and de power can be sup-plied to the airplane by an external power source. Dur-ing captive flight, the carrier airplane can supply ac anddc power to the airplane. Both external power sourcessupplyminimum dc power for initial relay or valve oper-ation only. Large amounts of dc power then are suppliedfrom the ac system through the transformer-rectifiers.AC ELECTRICAL POWER DISTRIBUTION.Two ac generators supply 200/115-volt, 400-cycle,three-phase ac power to the two primary ac busses.Each generator is driven through a gear train by anauxiliary power unit. (Refer to "Auxiliary Power Units"in this section.) Two 26-volt ac busses are powered bythe No. 2 primary ac bus. Automatic frequency controlandvoltage regulation are provided for the ac generators.


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T.O. lX-15-1AC Generators.Each ac generator is driven through a gear train by anauxiliary power unit and supplies 200j115-volt, 400-cycle, three-phase ac power to its respective primaryac bus. Failure of an APU causes failure of the acgenerator it drives. Hone ac generator fails for anyreason, the other ac generator automatically suppliespower to both primary ac busses. H either generatordrops "off the line" because of a momentary malfunc-tion, it can be reset "onto the line. "Primary AC Busses.The No. 1 primary ac bus normally is powered by theNo. 1 ac generator; the No. 2 primary ac bus, by theNo.2 ac generator. However, if either generator fails,the remaining generator will power both primary acbusses. External power, on the ground or from thecarrier airplane, will power the primary ac busses,but only when neither ac generator is on.26-volt AC Busses.The two 26-volt ac busses are powered through twoparallel transformers by the No. 2 primary ac bus.The 26-volt busses are powered as long as either acgenerator is operating. In addition, when externalpower is applied to the airplane on the ground or fromthe carrier airplane (and both ac generators are off),the 26-volt busses are powered.DC ELECTRICAL POWER DISTRIBUTION.Direct-current power is distributed from the 28-voltprimary dc bus and the battery bus.28-volt Primary DC Bus.The 28-volt primary dc bus is powered by both primaryac busses through two transformer-rectifiers. Theprimary dc bus, in addition to powering certain equip-ment, normally powers the battery bus. Failure of oneac generator will not de-energize the primary dc bus.The primary dc bus also is energized when externalpower is applied on the ground or from the carrierairplane.Battery Bus.The battery bus normally is powered by the primary dcbus. However, the emergency battery can be connectedto the battery bus to provide emergency dc power. Inaddition, external power on the ground or from thecarrier airplane can be applied to the battery bus.Emergency Battery.A stand-by, 24-volt, emergency battery is availableto provide emergency power to the battery bus.ELECTRICALL Y OPERATED EQUIPMENT.See figure 1-9.

Section IEXTERNAL ELECTRICAL POWER RECEPTACLE.The external power receptacle is on the upper surfaceof the fuselage, aft of the canopy. (See figure 1-16.)Whenac and dc external power is applied to the airplaneby a ground unit, an adapter must be used. When ex-ternal power is applied from the carrier airplane, asingle plug-in unit in the carrier airplane pylon is used.cmCUIT BREAKERS AND FUSES.The electrical distribution circuits are protected bycircuit breakers and fuses on the electrical power panelin the No. 2 equipment compartment and on the circuit-breaker panel (6, figure 1-4) on the right console in thecoclqJit. All of the circuit break~rs on the right consolepanel are of the push-pull type. The circuit breakers inthe No. 2 equjpment compartment must be properly posi-tioned before carrier take-off, because they are notaccessible in flight.

H the two external power circuit breakers inthe No. 2 equipment compartment are not closedbefore carrier take-off, carrier airplane electri-cal power cannot be applied to the X-15 Airplane.ELECTRICAL POWER SUPPLY CONTROLS.No. 1 Generator Switch.A three-position switch (15, figure 1-2) on the instru-ment panel controls operation of the No. 1 ac generatorby means of battery bus power. The switch is spring-loaded from the RESET position to ON. When the switchis OFF, the generator is taken "off the line." When thegenerator is "off the line, " because the switch is at OFFor because of a momentary generator malfunction, theswitch must be moved to RESET momentarily to bringthe generator "on the line" and then released to ON tomaintain the generator "on the line. "

NOTENeither generator will operate unless the APUfor the respective generator is also operatingand driving the generator.

To bring the No. 1 generator "on the line" initially, theswitch must be moved from OFF to RESET momentarilyand then released to ON. It is not necessary to movethe switch to OFF when the No.1 APU is shut down,because the No. 1 generator underfrequency protectiverelay will have tripped the generator off.

NOTEH either generator is operating and "on the line, "most external power is automatically discon-nected. Power to the battery bus, certainheaters, ready-to-Iaunch light, liquid oxygenlevel probe, and stabilization of stable platformremains on.

1-23


33/126

ction I

1T.O. 1X-15-1

CDn_~__mm_m\_n


34/126

FROM STRAINGAGE POWER

DCVOLTMmRSWITCH

DCVOLTMETER

un ............................

..................TRANSFORMERRECTIFIERS

T.O. 1X-15-1

EMERGENCYBATTERY

FROMEXTERNAL POWER

RECEPTACLE

.APU COOLING SWITCH. AUXILIARY LAUNCH SWITCH.ENGINE FI RE WARNING L IGHT ANDTEST CIRCUIT.EXTERNALPOWER RELAY.HELIUM RELEASE SELECTOR SWITCH.NO.1 GENERATORSWITCH.NO.2 GENERATORSWITCH.NO.1 APU SWITCH.NO.2 APU SWITCH

Section I

.NO.1 APU FUEL JmlSON SWITCH.NO.2 APU FUEL JmlSON SWITCH.NO. 1 BAWSTIC CONTROL SWITCH.NO. 2 BALL IS TI C CONTROL SWITCH.NO.1 APU COMPARTMENT OVER-HEAT LIGHT.NO.2 APU COMPARTMENT OVER-HEAT LIGHT.NO. I APU H,O, OVERHEAT LIGHT. NO.2 APU HoOt OVERHEAT LIGHT

.ADF. ALTERNATE CABIN PRESSURE SWITCH. ALTERNATE LONGITUDINALTRIM SWITCH. ANGLE-OF-ATTACK INDICATOR. BATTERY BUS CONTROL RELAY.BLOWERSWITCHES.CABIN SOURCEHELIUMSHUTOFFVALVE SWITCH. COCKP IT FLOODLIGHT SWITCH. COCKP IT PRESSURE SAFETY VALVE.COMMAND RADIO. FACE MASK HEATER SWITCH.ENGINE INDICATORLIGHTS.ENGINEMASTERSWITCH.ENGINERESETBUTTON.ENGINEPRECOOLSWITCH.ENGINE PRIMESWITCH.ENGINEMASTERSWITCH. ENGINE PURGE SWITCHES (GROUND). ENGINE H,O , COMPARTMENT HOTLIGHT. FUEL LINE-LOW CAUTION LIGHT. FUEL QUANTITY INDICATOR.IGNITERIDLESWITCH.INDICATORLIGHTSTESTCIRCUITS.INTERPHONEAMPLIFIER(AIC-10).JmlSON STOPSWITCHES.JETTISONTRIMSWITCHES.NASA INSTRUMENTATION. LONGITUDINAL TRIM SWITCH.NO.1 APUH,O,LOW WARNINGLIGHT

CD No.1 generator control relay. Energized openwhen No. 1 generator is operating. CD External power relay. Energized closed whenexternal power is applied and both generatorsare oH.

.NO.2 APU H,O, LOW WARNING LIGHT.NO .1 GENERATOR-OUT WARNING LIGHT.NO .2 G ENERATOR-oU T WARNING LIGHT.NO. 1 HYDRAUL IC TEMPERATURE GAGE. NO. 2 HYDRAUL IC TEMPERATURE GAGE. P ILOT 'S VENTILATED SUIT N, HEATERSWITCH. PROPELLANT EMERGENCY PRESSURIZA nONSWITCH. L IQUID OXYGEN AND AMMONIAEMERGENCY PRESSURIZATIONWARNINGLIGHTS.RASCONTROL. RASWARNINGUGHT.RATE-OF-ROLLINDICATOR.SAS TESTSWITCH.STABILITYAUGMENTATIONSYSTEM.STABLEPLATFORM.TURBOPUMPIDLEBUTTON.TURN-AND-SLIPINDICATOR

. VENTRALARMING SWITCH.VENTRALJETTISONBUTTON.VENTSUIT HEATERSWITCH.VIBRATIONMALRJNCTION CAUTIONLIGHT.WINDSHIELD HEATER SWITCHES.WING FLAPSWITCH

No.2 generator control relay. Energized openwhen No. 2 generator is operating. o Battery bus control relay. Energized closedwhen either generator is operating.

X-15-1-54-ZG

1-25


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ection I T. O. 1X-15-1

o. 2 Generator Switch.itchpositions and operation of this switch (17, figure-2), on the instrument panel, are identical to thoseor the No. 1 generator control switch, except that,viously, the No. 2 switch controls the No. 2 generator.ergency Battery Switch.

his guarded, two-position switch (16, figure 1-2), one instrument panel, controls connection of the emer-ency battery to the battery bus. Normally, the switchs in the guarded OFF position. Raising the guard andoving the switch to ON connects the emergency batteryirectly to the battery bus.Voltmeter Switch.

his three-position switch (8, figure 1-5), on the centeredestal, allows strain gage or primary dc bus voltageo be monitored. When the switch is at OFF, the dcoltmeter is disconnected. MOving the switch to STRAINAGE connects the dc voltmeter to the strain gageower. When the switch is at BUS, the voltmeter isonnected to the primary dc bus.ECTRICAL POWER SUPPLY SYSTEM INDICATORS.o.1 Generator Voltmeter.n ac voltmeter (11, figure 1-2), on the instrumentanel, shows the No. 1 generator voltage. The instru-ent has a range of 0 to 250 volts, graduated in incre-ents of 20 volts. The voltmeter reads line-to-line,ather than line-to-neutral or line-to-ground. Thus,e reading under normal conditions should be 200 voltsnd would check availability of two phases rather thannly one phase. (When both primary ac busses arenergized by one generator or by external power, bothenerator voltmeters should indicate the same voltage).

NOTEIf either generator fails, both the No. 1 andNo. 2 generator voltmeters will show the out-put of the remaining generator..2 Generator Voltmeter.n ac voltmeter (19, figure 1-2), on the instrumentanel, shows the No. 2 generator voltage. The in-trument has a range of 0 to 250 volts, graduated increments of 20 volts. The voltmeter reads line-to-ne, rather than line-to-neutral or line-to-ground.hus, the reading under normal conditions should be00 volts and would check availability of two phasesather than only one phase. (When both primary acusses are energized by one generator or by externalower, both generator voltmeters should indicate theme voltage.)

NOTEIf either generator fails, both the No. 1 andNo. 2 generator voltmeters will show the out-put of the remaining generator.

No. I Generator-out Light.Whenever the No. 1 generator drops "off the line, " theamber No. 1 generator-out caution light (14, figure1-2), on the instrument panel, comes on and reads"GEN OUT." The light is powered by the 28-volt pri-mary dc bus.No. 2 Generator-out Lip;ht.Whenever the No.2 generator drops "off the line, " theamber No. 2 generator-out caution light (18, figure1-2), on the instrument panel, comes on and reads"GEN OUT." The light is powered by the 28-volt pri-mary dc bus.DC Voltmeter.The dc voltmeter (7, figure 1-5), on the center pedestal,has a range of 0 to 30 volts, graduated in increments ofone volt. This voltmeter will indicate dc battery bus andstrain gage power voltage. The voltage reading selectionis through the dc voltmeter switch.HYDRAULIC POWER SUPPLY SYSTEMS.The airplane has a No. 1 and a No. 2 hydraulic system.(See figure 1-10.) Airplanes AF56-6670 and -6671 alsohave an SAS (stability augmentation system) emergencyhydraulic system. The No. 1 and No. 2 systems areindependent, but operate simultaneously to supply hy-draulic pressure to all hydraulically operated systemsof the airplane. Fluid is supplied to each hydraulicsystem by a reservoir, and pressure for each systemis maintained by a variable high-displacement pumpand a constant low-displacement pump. Each pump isdriven through a gear train from the APU. (Refer to"Auxiliary Power Units" in this section.) The hydraulicsystem No.1 pump is driven by APU No.1; the systemNo.2 pump is driven by APU No.2. The hydraulicsystems supply power for operation of the aerodynamicflight control system, speed brakes, and wing flaps.Dual, tandem hydraulic actuators are used, so thatfailure or shutdown of one hydraulic system will stillpermit the other hydraulic system to operate the variousunits. Each hydraulic actuator is capable of holdinghalf of the maximum design hinge moment during single-system operation, which is adequate for control andlanding of the airplane. The SASp itch-roll servo cylin-ders are powered by the No. 2 hydraulic system. TheSAS emergency hydraulic system will provide pressureto the pitch-roll servo cylinders in case of No.2 systemfailure or shutdown of APU No.2. The stability aug-mentation system yaw servo cylinder is powered onlyby hydraulic system No. 1. The No. 1 or No. 2 hy-draulic system is automatically in operation wheneverits respective APU is operating.SAS EMERGENCY HYDRAULIC SYSTEM.The SAS emergency hydraulic system (figure 1-10), onAirplanes AF56-6670 and -6671, provides hydraulicpower for the SAS pitch-roll servo cylinders in case offailure of the No.2 hydraulic system or its related APU.The 3000 psi system consists of a hydraulic motor,powered by the No. 1 hydraulic system, which drivesa variable-displacement, constant-pressure hydraulic


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T.O. lX-15-1pump. A pressure-operated selector valve directs No.1 hydraulic system pressure to the hydraulic motor incase of No. 2 system failure. The emergency systemhas its own reservoir.HYDRAULIC PRESSURE GAGE.A dual-movement, synchro-type gage (34, figure 1-2),on the instrument panel, indicates pressure in the No. 1and No. 2 hydraulic systems. There are two pointers,numbered 1 and 2, for indicating respective systempressure. The gage has a range of 0 to 4000 psi inincrements of 100 psi. The pressure indicating systemreceives power from the 26-volt ac bus.HYDRAULIC TEMPERATURE GAGES.Boththe No. 1 and No. 2 hydraulic systems have a gage(24 and 45, figure 1-2) which indicates hydraulic fluidtemperature. The gages are calibrated in degrees cen-tigrade through the range of _1000C to 3000C in incre-ments of 200 C. The temperature indicating systemsreceive power from the primary dc bus.FLIGHT CONTROL SYSTEMS.The airplane has two control systems. The aerodynamicflight control system (figure 1-10) consists of a mechan-ical system including hydraulically actuated controlsurfaces for use at altitudes where these surfaces areeffective for maneuvering the airplane. The ballisticcontrol system (figure 1-8) is used to control the air-plane attitude at altitudes where the aerodynamic sur-faces are relatively ineffective. The ballistic controlsystem uses a monopropellant which is released throughrockets at high velocity to rotate the airplane about itspitch, roll, and yaw axes as required for re-entry, andto correct oscillation.AERODYNAMIC FLIGHT CONTROL SYSTEM.The aerodynamic flight control system incorporateshydraulically actuated yaw and pitch-roll control sur-faces. The irreversible characteristics of the hydrau-lic system hold the control surfaces against any forcesthat do not originate from pilot control movement andprevent these forces from being transmitted back to thepilot controls. Thus, aerodynamic loads of any kindcannot reach the pilot through the controls. An artifi-cial-feel system is built into the control system tosimulate feel at the pilot controls. In-flight trimmingin pitch is accomplished by changing the neutral (no-load) position of the artificial-feel system and reposi-tioning the control sticks. Yaw control is provided bymovable upper and ventral vertical stabilizers. Theleft and right horizontal stabilizers provide pitch androll control, simultaneous operation for pitch control,and differential operation for roll control. On AirplanesAF56-6670 and -6671, an assist to the aerodynamicdamping in pitch, roll, and yaw is provided by a stabilityaugmentation system (SAS).Flight Control Hydraulic Systems.When both hydraulic systems fail, the aerodynamiccontrol surfaces will remain in the position at whichfailure occurred. However, the surfaces may be moved

Section I

in the direction in which they are driven by aerodynamicloads by repositioning of the pilot controls in the direc-tion to streamline the surface. The pedals, center stick,and console stick are mechanically linked to the controlvalves on their respective actuators. Movement of apilot control results in corresponding movement of itsactuator control valve. As the actuator moves, thecontrol valve is repositioned to a neutral position sothat flow to the actuator is shut off. The pressure re-maining in the actuator serves to hold the control sur-face in the desired position. Control cable rig tensionis maintained throughout a wide temperature and deflec-tion range by thermal expansion and contraction tensionregulators.Artificial-feel and Trim Systems.The artificial-feel system gives a sense of control feelto the pilot under all flight conditions where the aero-dynamic controls are used. Aerodynamic stick andrudder pedal forces are simulated by spring-loadedbungees in the control system. The bungees apply loadsto the pilot controls in proportion to stick or pedal move-ment, but the resultant feel has no relation to actual airloads. A nonlinear stick-to-stabilizer displacementratio is incorporated in the pitch control linkage to mini-mize sensitivity. Pitch trim is obtained by shifting theneutral "no-load" position of the fee

x-15 flight manual

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