flight controls

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Contents

1. PRIMARY FLIGHT CONTROLS.............................1-11.1 Ailerons.......................................1-1

1.1.1 Inboard and Outboard Ailerons............1-21.2 Elevators......................................1-21.3 Rudders........................................1-21.4 Spoilers.......................................1-3

2. TRIM CONTROLS.......................................2-5Typical Trim System.................................2-52.1 Rudder Trim System.............................2-52.2 Aileron Trim System............................2-52.3 TAILPLANE TRIM.................................2-5

2.3.1 Reasons for fitting to transport aircraft:2-6

2.3.2 Typical System Design....................2-62.3.3 Hydraulic Power Supply...................2-62.3.4 Position Indication Systems..............2-62.3.5 Tail-plane in Motion Warning.............2-6

3. ACTIVE LOAD CONTROLS................................3-14. HIGH LIFT DEVICES...................................4-1

4.1 Flaps..........................................4-14.2 FLAP CONTROL AND OPERATING SYSTEMS.............4-24.3 Asymmetry protection...........................4-9

5. LIFT DUMP & SPEED BRAKES............................5-16. SYSTEM OPERATION....................................6-1

6.1 Manual Operation...............................6-16.2 Powered Flight Controls (P.F.C.U’s)............6-2

6.2.1 Proportionality..........................6-26.2.2 Redundancy of hydraulic Supplies.........6-26.2.3 Tandem PFCU..............................6-36.2.4 Dual Assembly PFCU's.....................6-36.2.5 Duplicate/Triplicate PFCU's..............6-4

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6.2.6 Duplicated Control Surfaces..............6-56.2.7 Self Contained PFCU......................6-66.2.8 INPUT SYSTEMS............................6-66.2.9 High Speed Primary Controls..............6-7

7. FEEL, YAW DAMPER, MACH TRIM, RUDDER LIMITER, GUST LOCKS7-97.1 Artificial Feel................................7-97.2 Yaw damping....................................7-97.3 Mach trim......................................7-97.4 Typical System.................................7-10

7.4.1 Controller...............................7-107.4.2 Mach Trim Actuator.......................7-107.4.3 Operation................................7-10

7.5 Rudder limiting................................7-127.6 Gust locks.....................................7-12

8. BALANCING AND RIGGING...............................8-1

9. STALL WARNING AND PROTECTION........................9-19.1 Stall Warning Systems..........................9-1

9.1.1 Pneumatic Stall Warning System...........9-19.1.2 Electric Stall Warning System............9-2

9.2 Stall Protection System........................9-29.2.1 System Functions.........................9-29.2.2 System Components........................9-3

9.3 Actual Stall Protection System.................9-39.3.1 Incidence Probes.........................9-39.3.2 Nitrogen System..........................9-49.3.3 Automatic Ignition.......................9-49.3.4 Stall Warning............................9-49.3.5 Stall Identification.....................9-5


1. PRIMARY FLIGHT CONTROLSAircraft theory of flight has already been discussed in the module covering Theory of Flight. Aircraft are equipped withmoveable aerofoil surfaces that provide control in flight. Controls are normally divided into Primary and Secondary controls. The primary flight controls are: Ailerons Elevators Rudders SpoilersBecause of the need of aircraft to operate over extremely wide speed ranges and weights, it is necessary to have othersecondary or auxiliary controls. These consist of: Trim controls High Lift Devices Speed Brakes and Lift DumpNote: There is some variation of opinion as to whether spoilers are considered to be primary controls. The JAR 66 syllabus includes them as primary controls, so that is how

these notes will define them. Both types of controls are illustrated in the following diagram.

Fig. 1 Typical Aircraft Flight Controls

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1.1 AILERONS

Ailerons are primary flight controls that provide lateral roll control of the aircraft. They control aircraft movementabout the longitudinal axis. Ailerons are normally mounted on the trailing edge of the wing near to the wing tip.

1.1.1 INBOARD AND OUTBOARDAILERONS

Some large turbine aircraftemploy two sets of ailerons. Oneset are in the conventionalposition near the wing tip, theother set are in the mid-wingposition or outboard of theflaps. At low speeds both sets ofailerons operate to give maximumcontrol. Hydraulic isolate valveswill cut power to the outerailerons so that only the inboardailerons operate. If the outer ailerons are operated at high speeds, the stress on the wing tips may twist the leading edge of the wing downwards and produce “aileron reversal”.

1.2 ELEVATORSElevators are primary flight controls that control the movement of the aircraft about the lateral aircraft (pitch).Elevators are normally attached to hinges on the rear spar of the horizontal stabilizer. Fig 1 shows the typical location for elevators.

1.3 RUDDERS


The rudder is the flight control surface that controls aircraft movement about the vertical or normal axis. Ruddersfor small aircraft are normally single structural units operated by a single control system. Rudders for larger transport aircraft vary in basic structural and operational design. They may comprise two or more operational segments, each controlled by different operating systems to provide a level of redundancy.

1.4 SPOILERS

Spoilers are secondary control surfaces used to reduce or spoil the lift on a wing. They normally consist of multiple


flat panels located on the upper surface of the wings. The diagram below shows the more common configuration.

Operation of Spoilers on a Typical Aircraft

The spoilers lay flush with the upper surface of the wing and are hinged at the forward edge. When the spoilers are operated, the surface raises and reduces the lift. The spoilers may be used for different purposes.

1.4.1 FLIGHT SPOILERSFlight Spoilers are used in flight to reduce the amount of lift. If the pilot operates the controls left or right to roll the aircraft, the spoilers on the down-going wing move upward to aid rolling the aircraft. The movement of the spoilers is in proportion to the rate of roll required. On some aircraft, the spoilers are the primary flight control for rolling. If operating only as flight spoilers, only the surfaces on one wing will be raised at any one time. The flight spoilers are normally positioned outboard of the ground spoilers.

1.4.2 GROUND SPOILERS

Ground Spoilers are only used when the aircraft is on the ground. They operate with the flight spoilers to greatly reduce the lift on landing. The also reduce the drag after landing to slow down the aircraft. Ground spoilers will normally be deflected to their maximum position to give maximum drag on landing.

1.4.3 CONTROL OF SPOILERS

Spoilers will normally be controlled by the pilot through the normal roll controls or by the automatic flight control system (auto-pilot). They may also be operated automaticallyas part of an automatic landing system. On a typical aircraft (Boeing 757) the spoilers


2. TRIM CONTROLSThe majority of aircraft at some time during a flight, develop a tendency to deviate from a straight and level attitude. This may be caused by a fuel state change, a speed change, a change in position of the aircraft's load, or flap and undercarriage positions. The pilot can counter this tendency by continuously applying a correcting force tothe controls - an operation, which, if maintained for any length of time, would be both fatiguing and difficult to maintain. The tendency to deviate is therefore corrected bymaking minor trim adjustments to the control surfaces. Oncean aircraft has been trimmed back to a 'balanced' flight condition, no further effort is required by the pilot until further deviation develops.As fully powered flying controls are irreversible, i.e. all loads (reactions) are fed via mountings to structure, trim tabs would be ineffective.To overcome this, electric trim struts or actuators are usedwithin the input system. These actuators commonly reposition the "null" position of a self centring spring device to hold the control input system in a new neutral position. Thus the main control surface will be held deflected and the aircraft trimmed.

TYPICAL TRIM SYSTEM

The following is a typical trim system as used on a fully powered flight control system.

2.1 RUDDER TRIM SYSTEM

A typical rudder trim system for a powered system is shown in Fig 6. Trim commands from the trim switch causes the actuator to extend or retract, which rotates the feel and centring mechanism. This provides a new zero force pedal position corresponding to the trimmed rudder position. The trim switch is spring loaded to return to neutral. Both positive and negative elements of the circuit are switched to prevent a trim runaway should one set of switch contacts become short circuited. The trim indicator is driven electrically by a transmitter in the rudder trim actuator. The indicator shows up to 17 units of left or right trim. Each unit represents approximately one degree of rudder trim.


2.2 AILERON TRIM SYSTEM

A typical aileron trim system for a powered system is shownin Fig 7. Trim commands from the trim switches causes theactuator to extend or retract, which repositions the feeland centring mechanism null detent. The trim switches mustbe operated simultaneously to provide an electrical input tothe actuator, as both positive and negative elements of thecircuit are switched to prevent a trim runaway should oneset of switch contacts become short circuited. Theavailable aileron trim provides 15 degrees aileron travel inboth directions from neutral.

2.3 TAILPLANE TRIM

For trimming the aircraft longitudinally (about the lateral axis) the elevators are not trimmed. Instead the angle of incidence of the whole tailplane is altered. Raising the leading edge of the tailplane will increase lift over the tailplane which imparts a nose-down attitude to the aircraftor vice versa.This is done by mounting the forward end of the tailplane ona screwjack. Depending on the system the screwjack is rotated by two hydraulic or electric motors via a gearbox. Movement is induced by a lever in the flight deck which operates solenoid selector valves or an electric control circuit to operate the motors. Over-travel is prevented by micro-switch.

2.3.1 REASONS FOR FITTING TO TRANSPORT AIRCRAFT:

1. All aircraft benefit from having as large a range of useable centre of gravity as possible. This gives flexibility in cargo loading and allows for fuel usage ina swept wing.

2. Aircraft benefit from a wide speed range. Very simply, when an aircraft is trimmed at a particular speed, a reduction in speed calls for "up" elevator and an increase in speed calls for "down" elevator. This would cause extra drag.

3. The need to compensate for centre of pressure changes dueto slat/flap extension, gear extension.

4. To reduce trim drag to a minimum to give the optimum performance in cruise.


2.3.2 TYPICAL SYSTEM DESIGN

The tail-plane is pivoted at the rear of the centre section torsion box and attached to an actuator forward of the centre section. Operation of the actuator raises or lowers the leading edge of the tail-plane, altering the incidence angle..The actuator, comprises a re-circulating ball screw jack andnut assembly driven by two hydraulic motors with separate spur gear reduction trains.Friction brakes ensure that air loads cannot back-drive the actuator when the system is de-pressurised.The actuator is signalled from one of three sources:i) Auto-pilot servoii) Mach trim servoiii) Trim handwheel operation.A cable loop runs from the pedestal in the cockpit, under the cabin floor, and ends at a cable reduction gearing unit at the tailplane incidence actuator.

2.3.3 HYDRAULIC POWER SUPPLY

Each hydraulic motor is powered from a separate system. In the event of a single hydraulic system failure, a bypass valve permits that motor to "freewheel" when the system is de-pressurised.

2.3.4 POSITION INDICATION SYSTEMS

Geared indicator scales inboard of the cockpit hand-wheels present the demanded position of the tail-plane. This will be the actual tail-plane incidence with the hydraulic system(s) pressurized.Actual tail-plane position is continuously displayed on the pilot's instrument panel, signalled by a position transmitter operated by the tail-plane.External markings on the structure adjacent to the tail-plane give the approximate position of the tail-plane.

2.3.5 TAIL-PLANE IN MOTION WARNING

Some aircraft types have a tail-plane in motion warning system to alert the pilots of continuous motion of the tail-plane beyond a certain time period.


3. ACTIVE LOAD CONTROLS


4. HIGH LIFT DEVICESAerodynamic lift is determined by the shape and size of the main lifting surfaces of the aircraft. The shape of the normal wing is designed to give enough lift to support the aircraft in cruise. Consequently when the aircraft is flyingat lower speeds, during take-off and landing, it needs more lift. This extra lift is normally provided by the use of flaps and other high lift devices.

4.1 FLAPS

Flaps are widely used as a method of modifying lift. Most flaps are on the trailing edge of the wing, inboard of the ailerons. Some flaps are positioned on the leading edge and are designated “leading edge flaps”. The various types of trailing edge flap have already beenintroduced in earlier notes inaerodynamics. To recap, we have thefollowing types:Plain FlapsThis is simply a hinged portion of thetrailing edge, inboard of the ailerons. It can be lowered toincrease the camber of the wing. Lowering these will increase the lift coefficient and produce much more drag.

Split FlapsIn this type of flap, the trailingedge is split. Lowering the lower partincreases the lift coefficient morethan a plain flap and also greatlyincreases the drag.

Slotted FlapsThis acts like a plain flap exceptthere is a slot in front of the leadingedge of the flap. When the flap islowered, high-energy air from below thewing flows over the upper surface ofthe flap. This speeds up the airflowand delays separation to produce a higher lift coefficient than either a plain or split flap.Fowler flaps


These move rearwards and downwards and produce a much largerwing chord and area. The Fowler flap produces the greatest ratio of lift increase with drag.

Triple-Slotted Fowler FlapsThese are used extensively on moderntransport aircraft. They offer all theadvantages of both the Fowler flap andthe slotted flap.

4.2 FLAP CONTROL AND OPERATING SYSTEMS

On small aircraft the flaps are operated using hydraulic jacks to operate a single flap on each mainplane. This arrangement is not suitable for use on larger aircraft due to the size of the airframe that requires that the flaps aremanufactured and mounted in "segments" along the trailing edge.

4.2.1 FLAP CONTROL UTILISING LINEAR HYDRAULIC ACTUATORS

The following system that may be regarded as a simple system, similarly uses linear hydraulic actuators for an aircraft that has three flap segments on each mainplane eachpositioned by a separate hydraulic actuator.Movement of each actuator is controlled by a servo valve (simiIar to that in a primary flight control unit). Controlis by flap lever/quadrant on the centre console. This is connected to the actuator servo valves by a duplicated system of control cables and pushrods.

4.2.2 GENERAL

The flap surfaces are operated through linkages by hydraulicactuators. The actuators respond simultaneously to the control-cable-relayed demands of a selector lever mounted onthe flight compartment centre console.The piston rod end of each actuator is structurally anchored; movement being confined to the unit body. A position control element (servo-valve) incorporated in the body is controlled by an attached operating lever that has


limited travel on each side of the neutral position. The lever is moved towards or away from the anchored piston rod end to retract or extend the actuator. Each actuator incorporates internal restrictors that control the rate of response and an internal mechanical lock that engages when the flaps are fully up. The lock is hydraulically released when a down selection is made.The control system consists of a duplicated input circuit, which through the medium of a spring strut, signals all six actuators. Beyond the spring strut the signal to the inner flap actuators is conveyed by a rod and lever system and to the mid and outer flap actuators by interconnected signalling cables.The purpose of the spring strut is to "store" control lever movement due to the actuators' restricted rate of travel.The adjacent ends of the mid and outer flap surfaces are connected by a link that allows sufficient free movement to accommodate normal variations of relative positions without the links being loaded. The links are incorporated as a safety feature and take effect to prevent an asymmetric flapcondition.The flap selector lever is afforded the following gated positions - 0º, 5º, 15º and 30º.

4.2.3 HYDRAULIC POWER

Hydraulic power for operation of the actuators is provided by main system pressure backed up by flap accumulator pressure, when the flight compartment selector lever is at any position other than fully up (0º). The accumulator stored pressure is released to the flap system when a solenoid valve is energised open via a micro-switch operatedby the selector lever. The 'back-up' pressure is introduceddownstream of a non return valve in the main system pressureline; thus maintenance of a selected down position is assumed, for a limited period in the event of a main system failure.


4.2.4 CONTROL INPUT CIRCUIT

From the flap selector lever on the centre console, the duplicated input cables are routed aft through the roof structure to a position immediately aft of the rear spar. Atthis point, the cables are directed through the roof skin terminating with a double quadrant assembly. A double acting spring strut is connected between an output lever on the quadrant and a series of levers and control rods. These:

Operate the position control elements (servo valves) onthe inner flap actuators and transmit actuator movement to the inner flap surfaces.

Provide an input to the left and right mid and out flapsignalling circuits.


The spring strut is incorporated to allow a selection to be made in one quick movement - the total input motion being absorbed by the spring strut and progressively released as all six actuators respond at their controlled rate of travel.Each of the left and right mid and outer flap signalling circuits consists of a pulley drum from which cables are routed outboard to quadrant assemblies at the mid and outer flap positions. Output levers on these quadrants are linkedby control rods to the position control element (servo-valve) operating levers in the appropriate actuator package assemblies.The left and right pulley drums are interconnected by two tie-rods to ensure symmetrical operation of the left and right wing flaps.System OperationImmediately a selection is made the total input motion is absorbed by the spring strut and progressively released as all six actuators respond at their limited rate of travel. When the spring strut returns to its pre-selection settled length - the rod which connects to the position control element operating lever on each actuator arrests. The actuators will then marginally run on until their now restrained element operating levers reach neutral positions.This simultaneously creates a hydraulic lock at all six actuators and hence arrests the surfaces in alignment at theselected position.Safety AspectsTwo main safety requirements must be met.One is that a control cable break will not mean loss of control of the flaps. System integrity is such that duplication of the input cables which allows for functioningin the event of loss of either circuit) will maintain control.The other is that an asymmetric deployment of the flaps is prevented. An asymmetric condition could happen in several ways and the following mechanisms are designed to prevent these.Controls jamming between an actuator and surface (input systems intact):-Should this occur during a programmed selection, the input system of the relevant actuator will arrest and in consequence will stop signalling of the remaining actuators which will then run on marginally until their now restrained


servo valve operating levers reach neutral positions - thus arresting all six surfaces in approximate alignment.Mechanical failure between an actuator and surface (which will not impede surface movement):-Should this occur at either of the inner flaps - the system will remain functional (full asymmetry between inner flaps can be adequately countered by aileron action).

Should this occur at a mid or outer flap - the link which interconnects the adjacent ends of these surfaces will take effect to allow full functioning of both surfaces from one actuator. Thus preventing an asymmetric condition that would be beyond the ailerons ability to counter.Loss of signalling (cable break) to a mid or outer flap actuator.Should loss of signalling to a mid or outer flap actuator occur and the 'free' actuator become hydraulically locked atany stage during a programmed operation - the interconnecting link will arrest the adjacent functional actuator and thus its intact signalling system. This will have the effect of simultaneously arresting the interconnected input circuits of the remaining actuators that then run on marginally, until their now restrained servo valve operating levers reach neutral positions - thus arresting all six surfaces in approximate alignment. The actuator arrested by the link will remain programmed to achieve intended travel in opposition to the locked adjacentsurface. For this reason and to prevent excessive structural overloading - the actuators incorporate internal relief valves.Loss of main system pressureMain system pressure is augmented by flap accumulator storedpressure via a solenoid valve when the FLAPS selector lever on the flight compartment centre console is at any position other than fully up (0'). The 'back up' pressure is introduced down-stream of a non return valve in the main system pressure line; thus maintenance of a selected down position is assured for a limited period in the event of a main system failure.Position IndicationFlap position is indicated on a twin pointer scale calibrated to 0', 5', 15' and 30' settings. The flap position is signalled by two transmitters that are driven from the flap hinge arms via control rods.


Flap System Utilising Hydraulic Motors and Torque Tube DriveOn large aircraft it is more common for the flaps to be driven by twin hydraulic motors, each motor deriving its hydraulic supply from a different hydraulic system.Each motor is mounted on the same gearbox, such that drive from either or both motors will drive the gearbox.The gearbox is commonly located in the main gear bay. The drive is transmitted to the flap surfaces by a system of torque tubes, gearboxes and screw-jacks. The screw-jacks drive trolley assemblies along flap tracks mounted to the wing structure via support units. The flap segments are mounted onto the trolleys.System DescriptionThe flap system of each side of the aircraft comprises of flap sections supported and moved by six support/operating units. (Flap Tracks) The flaps are manually controlled by a lever on the central console to UP (0º), take off (20º), approach (35º) and landing (45º) positions. This manual control operates independent electro/hydraulic systems A andB, employed simultaneously to power the drive unit (gearbox)and their supplies are drawn from the aircraft electric and hydraulic systems bearing the same suffix letter.

Both systems normally operate together, but should ahydraulic system fail, or a fault develop which necessitatesselection of ISOLATE on one system, the flaps travel only athalf rate due to the design of the drive unit.

Drive Unit

The drive unit, comprises a gearbox and selector drumassembly, powered by two hydraulic motors. It rotates atorque shaft system that operates screw-jack and trolleymechanisms at each support/operating unit.

The drive unit is mounted to the rear of the wing rear spar member in the left main landing gear bay. It is powered by two hydraulic motor/lock valve assemblies; one supplied fromhydraulic system A and the other from system B. The motors drive a main shaft through a differential gear and a spur wheel reduction gearing. A gear driven selector drum operates micro-switches to arrest the flaps when they reach the selected position.


Flap Transmission System

Torque shafts extend outboard in each wing from either end of the drive unit main shaft. The sections of torque shaft couple via universal joints and serrated sleeve joints to bevel gearboxes and to intermediate bevel gearboxes. The bevel gearboxes and intermediate gearboxes are connected by serrated sleeve joints and universal joints to screw shaft assemblies located at each support unit. Flap trolleys fitted to each screw shaft engage via their rollers with trolley tracks fitted to the support units. These trolleys support the flap sections.

The flaps are hinged by pins to lugs on the flap trolleys. A torque link pivoted to each flap section carries a forwardflap trolley, the rollers of which engage with the cam trackon the support unit.

Hydraulic System

For redundancy the flaps are supplied by two independent hydraulic systems, which are identical. The following therefore describes one system only.

Hydraulic pressure is supplied to the flap selector valve via a flow control valve and isolating valve.Movement of the flap selector lever energises the appropriate solenoid selector valve to allow pressurised fluid to pass to the hydraulic motor through the lock valve. Return fluid from thehydraulic motor passes through the lock valve and flap selector valve back to the main system. The flow control valve controls the rate at which the flaps move. A throttlevalve slows down the flaps at all selected positions.

When the flaps reach the selected position the selector valve solenoid is de-energised, through the operation of theselector drum micro switches. The pressurised fluid is heldat the selector valve and the two service lines from the lock valve are connected together and into return. The lockvalve prevents the hydraulic motor from rotating.


Flap Control

Each separate flap operating hydraulic circuit is controlledby a separate 28 volt D.C. electrical system. Each supply is derived from a separate D.C. Bus Bar.

Each system is controlled by three micro-switches operated by control lever movement, these provide a circuit to the selector valve solenoids via six micro-switches operated by the drive unit selector drum.

Cams on the outer periphery of the selector drum operate oneswitch at both the normal, up and down limit positions of the flaps and two switches at the take-off (20') and approach (35') positions.

Over-Run Protection

If the micro-switches in the drive unit selector drum shouldmalfunction there is a probability that structural damage may occur as the flap trolleys reach the end of travel on their screw jacks. If a malfunction should happen, a set ofover-run micro-switches mounted on the flap support units, will be operated to interrupt the supply to the selector valve solenoid and prevent the trolleys bottoming on their screw jacks. These micro-switches are part of the complete control circuit and are operated by strikers on the flap support trolleys.

Asymmetry Protection

If a malfunction should occur in the flap transmission system causing one part to seize, great damage could occur as the drive system attempted to drive the flaps to their selected position. To prevent this, weak "fail safe" jointsare incorporated in various torque tubes that are designed to fail under a certain load.

However, this will allow the damaged portion of the system to stop and the remainder to continue travelling, so producing an asymmetric flap condition. To prevent this an


asymmetry protection circuit is incorporated in the control system.

This system uses an A.C. electrical supply and is controlledby four synchro's which are small devices mounted on and driven by screw shafts in board and out board of the flap systems. These are paired, Fig 34, and as they rotate send an alternating signal to an asymmetry control box. If the signals become out of phase with each other the over-travel/asymmetry isolate relay will be energised to lock-out the system.

Position Indication

Flap position indication is provided by a d.c ratio meter indicating system comprising two transmitters, driven from the outboard end of the left and right torque shaft systems and dual indicators positioned on the centre in the flight deck.

Maintenance of Flap SystemsLubrication

Because of the exposed position of most flap system components regular lubrication of hinge bolts, screw jacks, trolleys etc is required. When carrying out this task all excess grease must be removed to prevent the accumulation ofdirt or grit that may enter bearings etc.

Rigging

The flap operating system is a large complex system which will only work if all parts are in their correct relative positions at all times. To ensure this, whenever the systemis disturbed by a maintenance task it must be checked or re-rigged. Provision is built into the system for this.


4.3 ASYMMETRY PROTECTION


5. LIFT DUMP & SPEED BRAKES


6. SYSTEM OPERATION

6.1 MANUAL OPERATION


6.2 POWERED FLIGHT CONTROLS (P.F.C.U’S)

In large modern aircraft that fly at high speeds, the air loads on the f'lying control surfaces far exceed the abilityof the pilot to move them manually. To overcome this problem hydraulic pressure is used to move the control surfaces, a POWERED FLYING CONTROL UNIT or BOOSTER being used to convert hydraulic pressure into a force exerted on the control surface.In its simplest form, a P.F.C.U. consists of a hydraulic jack, the body of which is fixed to the aircraft structure and the ram, via a linkage to the control surface.To control the P.F.C.U. a servo valve (control valve) ismounted on the jack. The servo valve, which is connected tothe pilot's controls by a system of cables and/or pushrods,called the input system, directs fluid to either side of thejack piston and directs the fluid from the other side toreturn. This flow of fluid will displace the jack ram and asthis is connected to the control surface via an outputsystem of pushrods or cables, the control surface is moved.

6.2.1 PROPORTIONALITY

To make the controls "proportional" (i.e. the degree of movement of the jack-ram and hence the control surface, should be proportional to the degree of movement of the pilot's controls), a "follow-up linkage" is used. This linkage connects the input system, through a series of levers to the output system in such a way that the movement of the output system (jack ram) tends to cancel the input once the desired position is reached and so output movement ceases. In effect the movement of the jack ram is always trying to re-centre the servo valve and stop fluid flow in the jack.


6.2.2 REDUNDANCY OF HYDRAULIC SUPPLIES

Hydraulically powered flight control units usually derive their hydraulic power from the aircraft hydraulic system. If a PFCU obtained hydraulic power from only one hydraulic supply, a failure of that hydraulic supply due to an engine shut down, loss of fluid due to a leak, or failure of a hydraulic pump. The result would be loss of powered control of the aircraft. The probability of hydraulic failure is toogreat to allow a system to rely on one hydraulic supply, so redundancy must be introduced into the flight control system.As in the previous notes on hydraulic systems, modern large multi-engine aircraft, are arranged such that the engine driven pumps (and the other types of pumps) supply two or more independent hydraulic power supply systems.The following are methods that use that arrangement of hydraulic redundancy to allow failure of one hydraulic supply and still maintain control of the aircraft.

6.2.3 TANDEM PFCU


These are similar to the arrangement shown. They consist of a single jack ram but with two pistons. These pistons are housed in two co-axial cylinders each of which receives pressure fluid from separate power supply circuits via theirown duplicated servo valves. The servo valves, which are controlled by the same input system, are carefully set up inthe overhaul workshop to ensure they work in unison. This

prevents the two hydraulic pistons working against each other. With this arrangement a loss of one hydraulic supplywill allow the relevant piston to "free stroke whilst the other piston operates the control surface.


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6.2.4 DUAL ASSEMBLY PFCU'S

These are similar to the tandem arrangement but two piston rams are located in cylinders mounted side by side with the piston rams connected to a common output lever that transmits the movement to the control surface. The arrangement for the input system, the duplicated servo

valves and hydraulic fluid supply are the same.


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6.2.5 DUPLICATE/TRIPLICATE PFCU'S

In this arrangement each control surface is operated by two or three separate PFCU'S. For hydraulic redundancy, each PFCU is powered from separate hydraulic supply circuits. Ifone supply system should fail, or if one PFCU should malfunction the effected PFCU can be switched off. In this event a bypass valve within the PFCU will open interconnecting both sides of the jack ram. Therefore, as the pilot moves the input and operates the serviceable PFCU'S, the control surface will move and, "drag" the unserviceable PFCU ram with it. The open bypass valve will allow fluid to transfer from one side of the ram to the other as the PFCU "free strokes". Thus control will be maintained by the serviceable PFCU's driving the control surface, and a hydraulic lock in the unserviceable PFCU is prevented.


6.2.6 DUPLICATED CONTROL SURFACES

In this arrangement each control surface (rudder is shown inthe diagram) is split into two or three independent sections. Each section is operated by its own PFCU. For hydraulic redundancy, each PFCU is powered from separate


Page 7

hydraulic supply circuits. If one supply system should fail, or if one PFCU should malfunction the effected PFCU can be switched of. In this event the PFCU and its control surface segment will be "blown back" to the neutral positionby aerodynamic loads and held by a lock. Thus control will be maintained by the serviceable PFCU's driving their respective segments of control surface.All PFCU's are controlled via a single input system to a common input lever connected to all PFCU servo valves. Therefore if one PFCU malfunctioned it could prevent the operation of the remaining serviceable PFCU'S. To prevent this the input to the servo valves from the common input lever is via compressible spring struts or spring boxes. Innormal operation these spring struts/boxes resist compression and allow full control of all PFCU'S. If a PFCUis unserviceable, pilots input will compress the spring strut to that PFCU but the remaining spring struts/boxes will resist compression and operate the PFCU servo valves normally.

6.2.7 SELF CONTAINED PFCU

A self contained PFCU consists of a jack-ram powered by its own dedicated integrally mounted hydraulic “generator" and hydraulic reservoir. The generator is a radial piston pump arrangement within a slip ring assembly. The slip ring position is control ' led by a servo valve piston arrangement. With the slip ring held concentric with the piston bank no movement of the pistons within the rotating piston bank is allowed and no fluid flow will result. If aninput moves the slip ring the rotating bank of pistons will be allowed to "stroke" and a flow to the PFCU piston will occur and the PFCU ram will move. Movement of the slip ringin the opposite direction will cause fluid flow to the otherside of the piston and the ram will move in the other direction. The piston bank is rotated by a drive from a 3 phase electric motor which derives its supply from the aircraft electrical system.To maintain redundancy this type of PFCU will be duplicated and each may drive a duplicate and independent (split) control surface as above. As its source of power is electrical, it is independent of the aircrafts hydraulic system, therefore even with total hydraulic failure, controlcan still be maintained. On malfunction of a PFCU, or loss of electric power to that PFCU, it will lose hydraulic pressure and "blow back" to a neutral position where an integral lock will hold it. In this event further inputs to


the servo valves are absorbed by spring-strut that allows unhindered operation of the remaining PFCU'S.To give redundancy of electrical power supply, each PFCU in a "set" (i.e. rudder) gets its power supply from a differentbus bar.

6.2.8 INPUT SYSTEMS

Generally the input system of the powered flying control system is mainly a cable system with the related quadrants, pulleys and fairleads with the connections to the control column and the PFCU input lever by push rods. To guard against loss of control due to cable breaks the cable systemis duplicated. All duplicated runs are routed separately through the aircraft to avoid one incident damaging both control runs. The cable systems meet at a common input lever to the PFCU'S.

6.2.9 HIGH SPEED PRIMARY CONTROLS

Primary controls are designed to give adequate control in all flight phases. The flight phase at which the control surfaces are least effective is during low speeds (landing).


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This is because of the reduced aerodynamic effect with low speed. This means that the size and range of movement of each control surface must be sufficient to maintain sufficient control authority. With the control system designed to give efficient control at low speed, there may be a problem at high speed. This is that at high speeds theincreased air-loads on the control surfaces will cause them to be too sensitive producing over control and possible lossof control or over-stressing of the airframe. To prevent this two systems may possibly be used.

6.2.9.1 Geared Controls

In this system a single acting hydraulic jack may be fitted to an idler lever. The control rod is attached to this jackso that the radius of operation can be altered. Thus for a given angular movement of the idler lever, if the length of the jack is shortened, the linear movement of the control rod is reduced. This will maintain a constant range of movement at the pilots’ controls but reduce the range of movement of the control surface. Pressure at the jack is usually controlled by a pressure-modulating valve sensitive to a pressure transducer in the pitot system.

6.2.9.2 High Speed Control Surfaces (ailerons)

Normal, "low speed" ailerons are situated at the usual wing tips position to gain maximum authority due to the moment arm produced. But again at high speed their authority may be too great. In this system an additional set of "high speed" ailerons is also fitted at the wing root. Hydraulic isolate valves are incorporated in the control system such that at low speed the outer ailerons are functional, but at high speed, their hydraulic power is cut off and the high speed ailerons are powered to maintain roll control.The isolate valves are again controlled by pressure switchesin the pitot system.


7. FEEL, YAW DAMPER, MACH TRIM, RUDDER LIMITER, GUST LOCKS

7.1 ARTIFICIAL FEEL

7.2 YAW DAMPING

7.3 MACH TRIM

Modern transport aircraft are designed to cruise at high mach numbers, close to, or at the speed where shock waves may form on the wing. This is their "critical mach number".At this aircraft speed the formation of the shock waves causes shock induced separation and a movement of the centreof pressure forward. This produces a pitch up which must becountered.The Mach Trim System is provided to automatically maintain the correct aircraft pitch trim angle in relation to speed by varying the tail-plane trim. In achieving this function,the system maintains the same degree of longitudinal stability throughout the operational speed range of the aircraft.


7.4 TYPICAL SYSTEM

The mach trim system operates within the range from 0.68 IMN(Indicated Mach Number) to 0.84 IMN when the aircraft is above 9000 ft.The system operates in passive mode when the aircraft is flown with the auto-pilot engaged, but becomes active if theautopilot is disengaged.A mach trim activity light on the pilot's instrument panel flashes intermittently to indicate that a trimming demand exists. Illumination of the light for a sustained period indicates a runaway or seized actuator.A mach trim ON/OFF switch located in the cockpit permits a faulty system to be isolated.


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7.4.1 CONTROLLER

The controller is supplied with height and speed inputs fromthe aircraft pitot static system. The inputs are used to generate control signals that determine the direction and rate of rotation of the mach trim actuator. The controller also provides the 28 volts DC output to energize the clutch and connect the mach trim actuator to the tail-plane trim system.

7.4.2 MACH TRIM ACTUATOR

The actuator is located in the centre pedestal in the cockpit, and is connected by a chain drive to the manual tail-plane trim hand-wheels cross-shaft. A solenoid operated clutch connects the mach trim actuator to the drivesystem. The tail-plane auto-trim actuator operated by the auto-pilot system is also attached to the mach trim actuator. The control system ensures that only one actuatorcan be engaged at a time.

7.4.3 OPERATION

With the system selected ON and the auto-pilot disengaged, the mach trim actuator is clutched to the tail trim mechanism as soon as the aircraft power supplies are switched on.The system becomes active as soon as the aircraft flies above 9000 ft and its speed is within the Mach number range 0.68 IMN to 0.84 IMN.If the manual tail-plane trim hand-wheels are operated, the mach trim actuator is declutched to permit the tail-plane incidence to be changed and the clutch reengaged when the trim hand-wheels are released.



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7.5 RUDDER LIMITING

7.6 GUST LOCKS


8. BALANCING AND RIGGINGWhen applied to control systems the term 'rigging' is used to describe the practice of truing and checking the system to ensure that the flying controls operate correctly, hence,rigging a control system ensures that:The pilot's control is in the correct relationship to the relevant control surface.The control surface moves in the correct sense and to its designed maximum travel position in either direction.Friction in the system is within acceptable limits.

When Rigging is Carried OutThe rigging and adjustment of the system are carried out: At specified intervals as laid down in the relevant aircraftservicing publication.After disturbing any part of the control system, including the control system.

Checks Before RiggingBefore operating any flying control system in an aircraft, first check that there are no obstructions which could damage the control surface when it is moved. It is also important to display warning notices informing personnel of the possibility of movement of the control surface. Inform personnel working in the vicinity of a control system when you are about to operate it.

In rigging an aircraft control system it is sometimes necessary o level the aircraft both laterally and longitudinally – to put it into the rigging position, as described. The appropriate aircraft maintenance manual will state on what occasions, if any, this is necessary.Before starting to rig a flying control system it is advisable to ensure that all parts of the system and the control surfaces are serviceable. There is little merit in rigging a control system only to discover, subsequently, that some parts have to be replaced. Thus cables and tubes should automatically be examined for wear and corrosion, andother components for freedom of movement, security of attachment and so on. Replace components as necessary before continuing.


Establishing the Neutral SettingThe first action is to set the cockpit to neutral and to lock it in this position, using the equipment provided for the particular system. The rest of the control run is then adjusted to the neutral setting and locked in that position,often by using rigging pins. Generally speaking, control surfaces are in neutral when they are in line with the main surface to which they are attached. An exception to this iswhere the trailing edge of the aileron is set a specified amount below the mainplane trailing edge. This setting is known as aileron droop.Rigging PinsRigging pins are issued in sets, the type and number depending upon the aircraft and also upon the specific control run being rigged. The type, number and positions ofrigging pins in the aircraft's system are shown in diagrams of appropriate aircraft maintenance manual. The first pins, called the No. I or master pin, is fitted at the cockpit end of the control run and, in conjunction with the cockpit control neutral setting bar, secures that end of thesystem in neutral. Between these two items, there may be anadjustable link which has to be set at the correct length. By adjusting the control cable and tubes, holes in idler gears or levers can be made to align with corresponding holes in the airframe structure; rigging pins are then used to join these two holes, thereby positively locating and locking the control system in neutral (see Fig. 24). When all the rigging pins have been fitted in this way, that particular control run has been adjusted to, and locked in,neutral. This setting may be checked by using setting gauges.The next stage is to remove and then refit each rigging pin in turn to ensure that this can be done without strain. This indicates that the system has been set up satisfactorily, and that there is no backlash in the system;this is particularly important where the system is cable operated. Finally, it is vital to check that the complete set of rigging pins are removed from the aircraft on completion of the work.Control Surface Setting GaugesWhere control surface setting gauges are provided, they are used to check the neutral and maximum travel position of controll- able aerofoil surfaces. Each gauge is manufactured for use with one specific surface (Fig. 18). The gauge is firmly attached to a fixed part of the aircraft, next to the movable surface with which it is Mod 11.9 Flight Controls by www.amevoice.com Page 2

associated. With the controls set at neutral, the trailing edge of the control surface should coincide with the neutralmark on the gauge. Now move the control surface to the maximum travel position, in either direction, and see if thetrailing edge of the control surface coincides with the appropriate mark on the gauge. The control surface movementcan be quickly and easily adjusted with the gauge in position by restricting the mechanical stops.

Checking for Sense of MovementHaving established the neutral position of the control system, the next stage is to ensure that the control run being rigged operates the control surface in the correct sense. This is clearly vital; inadvertent cross-over of connections would reverse the control surface movement with possible disastrous results. The sense of operation can be readily checked by two tradesmen (see Fig. 19) - one at the control in the cockpit and the other at the control surface,if you are not sure of the relationship between control movement and the corresponding control surface movement.Mechanical StopsThe next check is to ensure that the control surface moves to its designed maximum travel position, in both directions, when moved by the cockpit control. The maximum travel of a primary control surface is limited in either direction by mechanical (limit) stops. These stops (see Fig. 20) are fitted to limit the control surface movement due to excessive travel. In a manual system, the limit stops are usually located near the control surface, and a second pair of stops, known as ' override stops ' , are fitted to limit the pilot's control movement should the mainstop fail. Override stops are adjusted to a specified clearance under normal operating conditions. In powered control systems, the mechanical stops are-located on the input (PFCU); usually they are located next to the pilot's control in 'the cockpit, thus limiting the control system movement from that position. During the rigging procedure, the main mechanical (limit) stops may need to be re-set to ensure that the control surface reaches, but does not exceed, its maximum travel position. The maximum travel position of a control surface can be checked in a variety ofways using the instruments detailed later. However, most modern aircraft use control surface setting gauges for this purpose.We have now rigged the control system and also checked that it operates in the correct sense; and we have set the limit Mod 11.9 Flight Controls by www.amevoice.com Page 3

stops to give the maximum required travel in both directions. The next stage is to check the system for resistance to movement from rest and also the force requiredto maintain the speed of movement when the control system isoperated.Checking for Static and Running FrictionThe resistance to movement of a control system may be due tolack of lubrication, misalignment, or slight faults in bearing surfaces. This resistance can be measured using a spring balance attached to the cockpit control; an example of this is shown in Fig. 21. The pull on the spring balanceis in the direction that the control would normally move. Note the reading on the spring balance when the control starts to move from rest. This force is known as the break-out force, and represents the amount of static friction in the system. Once the control system is moving, the force required to keep it moving is less than the break-out force.The spring balance indicates this reduced force, which represents the running friction. The amount of static and running friction permissible in any given aircraft control run must not exceed the limits laid down in the appropriate aircraft maintenance manual. Insufficient lubrication will,of course, increase the friction of any parts that rub together.


9. STALL WARNING AND PROTECTIONIf an aircraft is flown at a high angle of attack, lift willbe increased. However, if the angle of attack is increased to too great an angle, the airflow over the wings will separate and become turbulent. This will cause the lift to instantly fall to a very low value and the wing (or aircraft) is said to have stalled.The design of some aircraft will give an inherent indicationof an approaching stall condition. The airflow or wake leaving the wings will become progressively more turbulent as the stall is approached. This turbulent wake will strikethe airframe structure or tail-plane causing a condition known as "buffet". The pilot will normally recognise this as an indication of an impending stall and take appropriate action to prevent it, i.e. push the control column forward to reduce the angle of attack. Many aircraft do not have this inherent warning characteristic of buffet, therefore these aircraft require a system to warn the pilot of an impending stall. There are several stall-warning systems inuse.

9.1 STALL WARNING SYSTEMS

9.1.1 PNEUMATIC STALL WARNING SYSTEM


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This system is common on light aircraft. In this system a plenum chamber is mounted in the wing leading edge. This is covered and sealed by an adjustable plate that acts as part of the leading edge. The plate is adjusted so that in normal flight attitude a slot in the plate coincides with the stagnation point of the wing. The plenum chamber is connected by tube to a horn/reed assembly in the cabin.As the angle of attack is increased the slot in the adjustable plate effectively moves up from the stagnation point into an area of progressively lower air pressure. Theslot is so positioned that it reaches a low pressure area sufficient to draw air through the horn/reed assembly which will emit a noise and alert the pilot to an impending stall.

9.1.2 ELECTRIC STALL WARNING SYSTEM

This is typical of a system fitted to larger aircraft. Thisis a simple system that employs a micro-switch (transducer),operated by a vane. The transducer is mounted in the wing leading edge such that the operating vane is at the stagnation point during normal flight. Therefore no air-loads are imposed on the vane and it is not deflected from its null position.As the aircraft angle of attack increases the transducer operating vane effectively moves up and away from the stagnation point. The air-loads on the vane will increase until at a set angle of attack they overcome a spring


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pressure to deflect the vane and close the micro-switch contacts. This completes a circuit to illuminate a warning light and sound a warning horn. This should occur just prior to reaching the stall.These systems are found on relatively simple or small aircraft. Larger and more complex aircraft generally require a more sophisticated system that will do more than just warn of impending stall. This is termed a stall protection system.

9.2 STALL PROTECTION SYSTEM

9.2.1 SYSTEM FUNCTIONS

5. Stall Warning

As with the previous system this tells the pilot that he is approaching a stall condition.6. Stall IdentificationThis detects an imminent stall and automatically takes action to prevent the stall occurring, i.e. the stick is automatically pushed forward by the system. This may be achieved by a hydraulic or pneumatic jack acting on the elevator control system.7. Auto IgnitionIn some aircraft, particularly rear engine aircraft, disturbed airflow entering the intakes may cause the enginesto flame out near or at the stall. To prevent this an auto ignition circuit may be initiated on a stall warning/identification condition to prevent this.8. Flap/Slat/Krueger Flap ModulationAs flap, slat and Krueger flap position affect the stall angle the stall protection system may include the monitoringof their position and delay the initiation of stall warning.

9.2.2 SYSTEM COMPONENTS

1. Stall Warning SensorsThere are several designs in use. They may be mounted on the main-planes or side of the fuselage. They are normally duplicated, each providing a signal to a duplicated system.2. Stall Warning ComputerReceives signals from the sensors and initiates warnings or control movements.


3. Stick ShakerThe Main stall warning device. An electrically driven, out of balance rotor which shakes the control column when a stall warning condition is detected.4. Stick PusherA hydraulic or pneumatic ram which pushes the control columnforward when a stall identification condition is sensed. Itmay usually be over-ridden by higher than normal pilot force.5. Ground/Flight SensingTo prevent unwanted operation of the system on the ground a circuit through the landing gear weight switches disarms thestall protection system on the ground.6. TestA pre-flight test facility is built into the system.7. Mach SensingSpeeds over the aircraft critical mach number may cause highspeed stall or flame out. To prevent this an input to the computer from macs switches or the air data computer may be included to give a stall warning at high niach numbers.

9.3 ACTUAL STALL PROTECTION SYSTEM

The following is the description of an actual system used ona large passenger aircraft.

The stall protection system provides the following during the various phases of approach to the stall:1. Automatic ignition on all four engines.2. Stall warning by the operation of a stick shaker on each

control column.3. Stall identification by the sounding of a klaxon for each

system, allowed by operation of a ram to move the controlcolumns forward.

9.3.1 INCIDENCE PROBES

Four slotted conical probes, Fig 16, are mounted, two on either side of the forward fuselage, and project into the air stream. Each probe can rotate about its own axis through 50º in pitch, 4º of which are above fuselage datum. The probe detects the direction of airflow and transmits to the computer unit a voltage, picked off from potentiometers,


proportional to the angle between the airflow and the fuselage datum.When the aircraft angle of incidence is steady, pressure acts equally on the two probe slots, but as the angle of incidence changes, differential pressures are set up which, applied to the opposite sides of a paddle wheel, cause the wheel to rotate the probe until the pressures are again equal, i.e. the direction of flow bisects the angle between the slots.Ice protection for the probes is provided by heaters supplied from the No 1 and No 2 essential 28 volt dc supply.The left probe heaters are controlled by the first pilot's pressure head heater switch and the right probe heaters by the No 2 autopilot pressure head heater switch.The heater supplies are monitored by separate ammeters on the engineer's engine panel. When the aircraft is on the ground the current is limited by a resistor in series with the power supply.

9.3.2 NITROGEN SYSTEM

Nitrogen is stored at 1,500 ps,i in a reservoir. Nitrogen ispiped via a stop valve to a pressure reducing valve and non-return valve to a low pressure reservoir. Gauges monitoringthe high and low pressure are on the right sill panel and forward roof panel respectively. A relief valve in the low pressure line vents at 52 p.s.i. to prevent too great a pressure build-up in the system.Low pressure nitrogen is fed to solenoid valve A and from there through solenoid B to a control ram which operates on the control column linkage. A dump valve operated from a STALL DUMP VALVE lever on the centre console is coupled to this part of the circuit, and when the lever is set to DUMP,pressure in the line is released and prevents further operation of the stick pusher until the lever is reset.

9.3.3 AUTOMATIC IGNITION

Automatic ignition is signalled from the lower two of four Ferranti-type probes located on each side of the forward fuselage. It is switched on at a predetermined incidence, which is modified by slat position and mach number and remains on as long as the incidence is at or above this value. Indication of igniter operator is shown on the engine start panel. The system is brought into operation earlier whenever the slats are in or whenever 0.74M is exceeded. The system, which is physically shared with, but electrically isolated from the stall identification system, Mod 11.9 Flight Controls by www.amevoice.com Page 5

consists of two computer units, two mach switches and two angle of incidence probes. One of the two igniters on each engine is coupled to its associated computer, thus providinga completely duplicated and independent system.

9.3.4 STALL WARNING

The stall warning function, is provided by two duplicated systems, No 1 and No 2, each containing a computer unit, a lift rate modifier, an angle of incidence probe, and a stickshaker motor. Stall warning is signalled by the upper two of the four fuselage mounted probes. One probe is dedicated to shaker system. The warning is signalled at a predetermined incidence which is modified by a combination of flap position, slat position and rate of change of incidence. It remains in operation as long as the incidenceis at or above this value.One stick shaker is mounted on each control column and is connected to respective computer unit and lift rate modifier, thus providing duplicated and independent indication of stall warning.

9.3.5 STALL IDENTIFICATION

Stall identification is provided by two duplicated systems, No 1 and No 2, each containing a lift rate modifier a solenoid operated valve, interlock relay and delay unit, a warning horn and an angle of incidence probe which is sharedwith, but electrically isolated from, the auto-ignition system.Identification of a stall is signalled by the two fuselage mounted probes which signal auto-ignition. The signal occurs at a predetermined incidence set at a level that is always above the stall warning value. This predetermined incidence is modified by, a combination of flap positions, slat position and rate of change of incidence. The stall, identification system operates only if armed by a prior stall warning signal, and remains in operation as long as the incidence is at or above the modified level.The system reverts to normal operation once the stall warning signal is cancelled by resuming normal flight.When the stick shakers operate, a priority circuit receives signals from each of the computer units of the stall warningsystem. The first signal received is passed to the stall identification interlock relays to arm the solenoid valves circuit.


The signal from the stall identification probes is fed to the appropriate computer unit, and when the signal reaches aparticular value, the unit supplies a 28 volt dc output. Thevalue of the signal can be changed by combinations of the flap and slat position compensation. The signal is passed through the lift rate modifier so that a quick rate of change of the probe angle causes an advanced signal, provided that has been preceded for 0.7 seconds by a stick shaker signal.The computer unit output is passed through a priority circuit to the stall identification relay in the interlock circuit. Providing the sequence is correct, this completes the circuit to the solenoid valves that open to allow nitrogen to the rams that extend to move the control column forward. The warning horn in each system sounds when the respective stall warning and stall identification computer units both signal, which is simultaneous with control columnmovement.Both solenoid operated selector valves are opened by a stallidentification signal. The opening of each valve is indicated by the associated red light on the overhead panel,and the subsequent movement of the ram is indicated by the STALL IDENT amber light adjacent to the airspeed indicators on each pilot's panel also coming on.The system is pneumatically powered from an HP nitrogen bottle that feeds the stick pusher ram through a reducing valve, an LP reservoir and the two solenoid-operated selector valves. A gauge on the forward roof panel indicates the pressure in the low pressure reservoir and another on the right sill panel indicates the pressure in the HP bottle. Minimum HP pressure for flight is 500 psi. When pressure falls to 32 PSI, the LP red light on the forward roof panel comes on.


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