Airbus 27 A300 A310 Flight Controls

MTT M540000 R3.3 01AUG01MTT M540000 R3.3 01AUG01For Training Purposes OnlyFor Training Purposes Only

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THIS PAGE INTENTIONALLY

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FLIGHT CONTROLSURFACES -

GENERAL


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A300/A310 FLIGHT CONTROL SURFACES

The control of the aircraft is achieved by:• the primary flight controls• the secondary flight controls

The primary flight controls ensure:

• ROLL CONTROL achieved on each wing by:- one aileron- five roll spoilers, upper wing surfaces No. 3 through No. 7.

• PITCH CONTROL achieved by two elevators hinged on thetrimmable horizontal stabilizer.

• PITCH TRIM CONTROL achieved by the trimmable horizontalstabilizer hinged on the aircraft structure.

• YAW CONTROL achieved by one rudder.

The secondary flight controls are the:

• FLAPS- three single slotted flaps on each wing

• LIFT AUGMENTATION devices on each wing- three slats- one Krueger flap- one notch flap—not applicable to A310

• SPEEDBRAKES No. 1 through No. 5 on the upper surface of eachwing

• GROUND SPOILERS No. 1 through No. 7 on the upper surface ofeach wing


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A300 Flight Control Surfaces


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Flight Compartment Controls and Indications

This illustration depicts all the controls and indications for the flightsurfaces located in the cockpit.

A. Servo Control Panel

B. Slats and Flaps Position Indicator

C. Pitch Trim and Yaw Damper Switch Panel

D. Flight Control Maintenance Test Panel

E. Right ECAM Display Unit

F. Pitch Trim Wheel

G. Speed Brake Control Panel

H. ECAM Display Control Panel

I. Aileron and Rudder Trim Switches

J. Left ECAM Display Unit

K. Master Warning and Caution Lights L/H

L. Master Warning and Caution Lights R/H


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Flight Compartment Controls and Indicating


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FLIGHT CONTROLS HYDRAULIC POWER SUPPLYThe flight controls are powered by the three independent hydraulicsystems; redundancy is such that with two hydraulic systems failed, the

remaining system can operate the aircraft within an acceptable range ofthe flight envelope.


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Flight Controls Hydraulic Power Supply


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FLIGHT CONTROLS - GREENHYDRAULIC POWER SUPPLY


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Flight Controls - Green Hydraulic Power Supply


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FLIGHT CONTROLS - BLUE HYDRAULIC POWER SUPPLY


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Flight Controls - Blue Hydraulic Power Supply


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FLIGHT CONTROLS - YELLOW HYDRAULIC POWER SUPPLY


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Flight Controls - Yellow Hydraulic Power Supply


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SERVO CONTROL P/B SWITCHES

1. SERVO CTL PUSHBUTTON SWITCHES

All these P/B switches are guarded. These P/B switches controlthe servo shut-off valves for the individual hydraulic circuits Blue,Green and Yellow.

• NORMAL (P/B SWITCH PRESSED-IN)Hydraulic power is supplied to the corresponding users assoon as pressure is available in the corresponding hydraulicsystem.

• OFF (P/B SWITCH RELEASED-OUT)The OFF light comes on White and the hydraulic powersupply to the corresponding users is shut off. The associatedJAM warning is inhibited and LO PR Amber illuminationconfirms the OFF selection.

• JAMWhen a P/B switch is pressed-in, the associated JAM lightcomes on Amber when a jamming is detected in the relatedhydraulic control valves of rudder, elevator, ailerons ortrimmable horizontal stabilizer. Illumination of a JAM light isaccompanied by ECAM activation. The jammed control isidentified on the Warning Display (left CRT).

2. B, G, Y LO PR LIGHTS

A light comes on Amber when the flight control supply pressure inthe corres ponding hydraulic system has dropped (below 1450 PSI)downstream of the servo control valve, or when the hydraulic supplyhas been shut off. Illumination of an Amber LO PR light isaccompanied by ECAM activation.

NOTE: The SERVO CTL P/B switch positions and associated warningsare repeated on the ECAM hydraulic system page.

- To prevent inadvertent complete deactivation of servo controls, onlytwo systems can be deactivated at a time by selection of SERVOCTL P/B switches to OFF. When the third P/B switch is selected toOFF all three systems are reactivated regardless of P/B switchsetting.


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SERVO CTL (Servo Control) Panel


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ROLL CONTROL

SECTION


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ROLL CONTROL

The roll control surfaces on each wing are:• One (1) aileron powered by 3 servo controls• 5 roll spoilers, each one powered by one (1) servo control.

The spoiler system is supplied from two normal bus bars (28 V DC and26 V AC). If the normal buses have been cut off before landing, power issupplied again to three spoiler groups by pressing the LAND RECOVERYP/B switch on the overhead panel.

From the two interconnected control wheels, the roll inputs aretransmitted to the ailerons by dual cables providing fail safe operation. Ineach wing the inputs are transmitted to a differential unit receivingadditional inputs from:• artificial feel unit• aileron droop unit• trim screw jack

In case of jamming in one control run, the interconnected spring strut canbe compressed to permit operation of the other control run to the otherwing. The pilot effort required on the wheel is between 34 lbs. and 90 lbs.Spoiler control is still available but downgraded. Each servo controllinkage on the aileron includes a spring rod to protect it against arunaway if an input lever on one jack remains in the open position.

The artificial feel is provided by a spring loaded rod. The, trim actuator iselectrically signaled by a control on the center pedestal. In order to

improve the aerodynamic characteristics, a droop signal coming from theslats control system moves the ailerons down 9.2° maximum when theslats are extended. During cruise, the operational limits for aileron trimare ±2°. The roll spoilers and speedbrakes are electrically signaled bytwo identical computers (EFCU-Electrical Flight Control Units) thatelaborate the roll orders by processing the signals coming from thecontrol wheel position transducers units.

Each computer is composed of two control units and two monitoringunits. Each unit controls or monitors one group of surfaces. Each group ismade of one or two pairs of servo controls: spoilers 2-3, spoilers 4-1,spoilers 5-7, spoilers 6. Thus, for a group of servo controls, thecorresponding control unit is in one computer and the monitoring unit is inthe other one. For the roll spoilers the control laws are such that they arenot usually used unless the control wheel is moved enough. An autopilotservo actuator is mounted adjacent to the RH wing rear cable quadrant. Itdrives the complete control via a detent lever which can be overridden bythe pilots.

INTERFACE WITH AUTOPILOT SYSTEM

An autopilot actuator is mounted adjacent to the right wing rear cablequadrant; it drives the complete control via a detent lever which can beoverridden by the pilots. Dynamometric rods are installed upstream of thecable tension regulators, they provide control signals to the control wheelsteering system.


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Roll Control - Mechanical Aileron Control


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AILERON SYSTEM - COMPONENTS DESCRIPTION

1. CABLE TENSION REGULATORSTwo tension regulators maintain a constant tension on the cables of28.13 ± 5.30 lbf. They are identical apart from the input leverposition. They incorporate provision for installation of a special toolused for installing the regulator on the aircraft.

2. SERVO CONTROL ACTUATING SPRING RODThe three ASA servo control actuating spring rods prevent runawayof the control system if an input lever jams on its servo control body.

3. CONTROL WHEEL INTERCONNECTING SPRING RODThe two control wheels are interconnected by a spring rod in order toallow one of the crew members to control half the surfaces in theevent of any single item jamming in the mechanical control system.

4. RODSPush-pull rods are adjustable or nonadjustable length, fitted withreplaceable ends.

5. CABLESThe flexible cables (Dia. 3.2 mm/0.126 in.) are made of zinc-coatedcarbon steel. The cable end fittings are equipped with barrels forquick installation and fool proofing; turnbuckles are cliplocked.Fairleads are of the roller type, for low friction purposes. The fairleadsupports allow passage of the cable end fittings. At bulkheads,cables are fed through pressure seals.

6. DYNAMOMETRIC RODSThe Flight Control Computer uses signals from the dynamometricrods to detect the Captain's and First Officer's loads on the controlwheels. There are two rods for the pitch axis and two for the roll axis.There is no rod in the yaw axis. The rods are placed in series in theFlight Control linkages.


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Aileron System - Components

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All Speed Aileron (ASA) - Mechanical Control

Each all speed aileron (ASA) is operated by three mechanicallycontrolled servo controls. The two interconnected control wheels drivetwo symmetrical control systems composed of levers, rods, cables andtension regulators routed along each side of the fuselage up to the inputlevers of the servo controls. A differential and droop unit is installed in thecontrol linkage upstream of the servo controls. The unit receives twoinputs. One is from the control wheels (pilots input), the other is a droopsignal from the slat control system which droops the all speed ailerons9.2° when the slats are extended in order to optimize aerodynamicefficiency of the wing.

When the droop signal is applied, all speed aileron deflection is notsimply modified by 9.2° throughout the travel range. Instead, response ofthe all speed ailerons to control wheel motion is modified so that themaximum up and down deflections remain close to those without droopinput. The droop signal also drives a differential mechanism between thetrim screwjack and the artificial feel unit. The mechanism pivots theartificial feel unit, thus allowing the spring rod to remain at neutral.


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All Speed Aileron - Mechanical Control


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Aileron Trim

Trim control is electrically signaled. An electrical actuator installed in themain gear W/W (center fuselage) drives two trim screwjacks viasprockets, chains and cables. The actuator is controlled from panel408VU located at the rear part of the center pedestal. Two switches onthis panel allow the crew to select constant speed displacement in theappropriate direction. Trim position is indicated on scales at the top of thecontrol columns when the wheels are released.

In each wing root, displacement of the trim screwjack drives the all speedaileron servo control input linkage through the artificial feel unit, whosespring rod remains at neutral. When the ailerons are drooped, the droopsignal drives a differential mechanism between the trim screwjack andthe artificial feel unit. The mechanism pivots the artificial feel unit, thusallowing the spring rod to remain at neutral and the unit is held in thisposition by the trim screwjack.


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Aileron Trim System


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Aileron Trim Components

A. AILERON TRIM ACTUATOR

The actuator is driven by a 28VDC electric motor through a reductiongear and a torque limiter. The motor is a permanent magnet motor withon-off control. A strong dynamic braking effect is obtained by shorting themotor windings as soon as they are no longer energized (no staticbraking on the actuator itself: trim irreversibility is provided by thescrewjacks downstream of the actuators).

• Rotary stops limit output shaft rotation within the range allowed bythe screws.

• When the electric motor is energized, it is protected by a torquelimiter when the stop limits are reached.

• A rigging pin is used to set the output shaft at mid angular travel(zero trim position and also zero reference for synchro transmittersettings).

B. ELECTRIC MOTOR CONTROL

The electric motor windings of aileron trim actuator 9CG areenergized through contacts of two adjacent three position switches(5CG) on control panel 408VU. The switches are spring loaded to thecenter position and must both be moved simultaneously in the samedirection for the windings to be energized. The switch tabs are notmechanically connected, to prevent run-away in the event ofmechanical jamming of one tab. The windings are shorted when thetwo switches are in the center position.


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Aileron Trim System - Components


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All Speed Aileron - Artificial Feel Unit

There are two identical artificial feel units, each installed immediatelydownstream of the all speed aileron servo control actuating spring rods.The units each include a spring rod and are held in position by the trimscrewjacks.

Their function is:• To maintain servo control input linkage in trim position in the event of

disconnection of the control linkage upstream of the servo controls.• To provide artificial feel loads proportional to control wheel

deflection.• To provide accurate return of the surfaces to neutral.


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All Speed Aileron - Artificial Feel Unit


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Aileron Trim/spoiler and Speed Brake Switches

1. AIL TRIM SWITCHES

Ailerons trim control is electrically powered. For safety purposes,both switches must be moved and held in the same direction (LWING or R WING) to energize the system. This action selects aconstant speed displacement in the corresponding direction. Fulltravel of about 7° of aileron in each direction is achieved at a speedof 0.4° per second.

2. AILERON TRIM SCALES

A scale representing 14° of aileron movement (7° in each direction)is engraved and painted on the top of each control column oppositea pointer painted on the control wheel. With the control wheelsreleased, the crew can thus read the actual aileron trim value.

3. SPLR & SPD BRK PUSHBUTTON SWITCHES

Each P/B switch is associated with one or two pairs of symmetricalupper wing surfaces.

• ON (P/B switch pressed-in):Corresponding control system is activated. Each time a system isactivated, or corresponding hydraulic system on, or the aircraftelectrical network is energized, a 2 second safety BITE test istriggered for the corresponding EFCU units (control and monitor).

• OFF/R (P/B switch released-out):The OFF/R light comes on White and the corresponding controlsystem is deactivated. If hydraulic pressure is available, theactuators are automatically held in the retracted position. Themonitoring circuits are reset by ECAM activation. This action isaccompanied by ECAM activation.

• FAULT:When a P/B switch is pressed-in, the associated FAULT lightcomes on Amber if a failure is detected by the monitoringcircuits, which then deactivate the control system. Illumination ofthe FAULT light is accompanied by ECAM activation.


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Aileron Trim/spoiler and Speed Brake Switches

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RUDDER CONTROL

SECTION


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Rudder System - Yaw Control

The rudder, operated by 3 mechanically controlled servo controls,receives pilot's inputs by a single cable run to a spring loaded artificialfeel unit connected to the trim screwjack. From this point up to the servocontrols, the commands are transmitted by dual rigid linkage, receivingadditional inputs from a rudder travel limiter, yaw damper and autopilotservoactuators. The artificial feel is provided by a spring loaded rod. Thetrim actuator is electrically signaled. It is driven by an electrical motor.During cruise, the operational limits for rudder trim are ±1.5°.

The rudder travel limiter reduces the pedal and rudder deflection from±30° at speed below 165 kt to ±5° at 308 kt and above. The orders aredelivered by two independent RUDDER TRAVEL channels, each oneincluded in a digital computer (Feel and Limitation Computer) receivinginputs from the DADCs (Digital Air Data Computers) and the SFCCs(Slats Flaps Control Computers). Each computer controls an electricalmotor driving a common electromechanical actuator coupled to variablestop lever. Only one channel is normally active. The other is in standby.A spring loaded rod positions the variable stop lever in the low speedposition in case of dual failure.

An. autopilot servo actuator is mounted adjacent to the artificial feel unitupstream of the variable stop lever. It drives the complete control via adetent lever which: car. be overridden by the pilot. Yaw dampercommands are transmitted via a differential unit canceling a feedback tothe pedals. A spring loaded rod on each servo control input avoids arunaway of the rudder in case of jamming of one input lever in the openposition. Levers are attached to each pedal, to provide brake inputswhen the pedals rotate around their pivots.

INTERFACE WITH AUTOPILOT SYSTEM

An autopilot actuator is mounted adjacent to the artificial feel and trimunit upstream of the variable stop lever; it drives the complete control viaa detent lever which can be overridden by the pilot. A yaw damperactuator, mounted between the artificial feel and trim unit and thevariable stop lever, drives the rear control via a differential linkage. Theyaw damper actuator signals are added to those of the pilots, up to themaximum travel allowed by the variable stop lever. The yaw damperactuator is fail-safe, so that disconnection of the control is extremelyimprobable.

INTERFACE WITH MAIN WHEEL BRAKING SYSTEM

Levers are attached to each pedal, to provide braking inputs when thepedals rotate about their axis.

INTERFACE WITH NOSE WHEEL STEERING

The nose wheel steering control is connected to the rudder controlthrough a hydraulic steering control coupler (engaged when the landinggear is extended) and a spring rod, the threshold of which is lower thanthe threshold of the rudder artificial feel and trim unit spring rod. Thespring rod prevents the nose wheel steering control from transmittinginputs to the rudder control.


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Rudder System - Yaw Control


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Rudder System - Rudder Trim Actuator

One 28VDC electric motor is fitted in the actuator, directly coupled to thereduction gear. It is energized when rudder travel is selected.

• The motor is a permanent magnet motor with on-off control. A strongdynamic braking effect is obtained by shorting the windings of themotor when it is de-energized (no static braking on the actuator itself:trim irreversibility is provided by the screwjack downstream of theactuator).

• Rotary stops limit output shaft rotation within the range allowed bythe screw.

• When the electric motor is energized, it is protected by a torquelimiter when the stops are reached (motor rotation is not stopped).

• The actuator includes a position transducer which delivers ruddertrim position signals to associated electrical circuits. The transduceris a special RVDT, of the same type as those installed in the twotransducer units used for electrical roll control. The electricalcharacteristics of the RVDT are monitored by associated computercircuits.

• A rigging pin is used to set the output shaft at mid angular travel(zero trim position and also zero reference for transducer setting).


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Rudder System - Rudder Control Input Components


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Rudder System Yaw Control - Rudder Trim Switches

1. RUD TRIM ROTARY SELECTORRudder trim control is electrically powered. The rotary selector isspringloaded to the neutral (center) position. The direction of ruddertrim travel depends on the direction of rotary selector (NOSE L orNOSE R). Full authority of rudder trim is about 21° in each direction.

2. RESET PUSHBUTTON SWITCH

It allows initiation of an automatic sequence controlled by the EFCUsto position the rudder trim at 0° ±0.2°.

• ON (P/B Switch pressed-in)The ON light comes on White. The switch is latched during thereset action and will release out automatically when reset isachieved.

• Normal (P/B switch released-out)Automatically or manually, the reset action is stopped and theON light goes off.

• FAULTThe light comes on Amber if a failure of the reset function isdetected or if the actuator position transducer fails.

3. RUD TRIM POSITION INDICATORA digital indicator displays rudder trim direction (L or R) and value (0°to 21°).


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Rudder System Yaw Control - Rudder Mechanical/Hydraulic

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Rudder System - Rudder Artificial Feel

A spring assembly located in the artificial feel and trim unit restores aresistance to pedal depression which is proportional to ruddermovement. A variable stop lever installed downstream of the servocontrols on the control linkage serves to reduce rudder deflection withrespect to pedal movement as the airspeed increases.

RUDDER ARTIFICIAL FEEL

An artificial feel and trim unit is installed adjacent to the rear cablequadrant. It consists of a trim screwjack and a fail-safe constant resistingload spring rod, held in neutral position by the trim screwjack.

Spring function is:

• To maintain the downstream linkage and the input lever of the servocontrols at neutral in the event of disconnection of the control linkageupstream of the artificial feel and trim unit

• To provide artificial feel loads• proportional to rudder deflection• To provide accurate centering of the surface at neutral in the

absence of a control input• To maintain the upstream controls at neutral, when signals are

provided to the servo controls by the yaw damper actuator.


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Rudder System - Rudder Artificial Feel


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Servo Controls

1. SAFETY VALVES

To preserve the Green system, safety valves are installed upstream ofthe following components:• Krueger selector solenoid valve (in case of engine failure)• Rudder servo control (in case of inflight collision)

2. SERVO CONTROLS JAMMING DETECTION

There is one jamming detection circuit for each hydraulic system. Ifjamming occurs the electronic circuitry inside the jamming detectioncontrol unit receives 28V directly from the jamming detection microswitchif a servo control is involved and from an intermediate logic if a THSactuator hydraulic motor is involved.

Jamming detection is associated with the mechanically driven controlvalves of the left and right all speed aileron, and left and right elevatorand rudder servo controls. It is also associated with the THS actuatorhydraulic motor control valves (for the Green and Yellow systems only).

NOTE: When a hydraulic system is selected OFF, the + 28V sent to thecorresponding jamming detection microswitches is cut off.


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Rudder System - Servo Controls - Components

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Rudder System Travel - System 1 and 2 Pushbutton Switches

1. RUD TRAVEL CONTROL PANEL1. The P/B switches control channels 1 and 2 of the Feel and

Limitation Computers (FLC) for rudder travel limiting.

• ON (P/B switch pressed-in):The corresponding system is engaged. Both systems may beengaged simultaneously, but only system 1 is effectivelyactive. If system 1 fails, it is automatically deactivated andsystem 2 becomes active.

• OFF/R (P/B switch released-out):The OFF/R light comes on White and the system involved isdisengaged. The monitoring circuits are reset by this action.This indication is accompanied by ECAM activation.

• FAULT:When a P/B switch is pressed-in, its FAULT light comes onAmber if a failure is detected in the respective system.Illumination of the Amber FAULT light is accompanied byECAM activation. Both FAULT lights remain illuminatedwhen the switches are released-out and the OFF/R lights areilluminated White. This constitutes a rudder disagree warning(The variable stop lever is not in low speed position withflaps extended 20° or more). Illumination of both FAULTlights is accompanied by ECAM activation.

2. YAW DAMPER LEVERS

• 1 (or 2): The lever is magnetically latched in active position andthe yaw damper 1 (or 2) is engaged. If a failure is detected, theYAW DAMPER 1 (or 2) lever trips to OFF.

• OFF: The respective yaw damper is disengaged. When oneYAW DAMPER lever trips to OFF, the associated yaw dampersystem disengages and the ECAM is activated. When both YAWDAMPER levers trip to OFF the yaw damper function is lost andthe SCAM is activated.


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Rudder System Travel and Yaw Damper Systems - Control Switches

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Rudder System - Rudder Travel Limiting

The rudder travel limiting system modifies control inputs to the servocontrols to vary rudder travel in relation to airspeed (Vc). Limitation issuch that the maximum deflection which can be achieved by the rudderremains lower than the deflection which would induce limit loads on thestructure, throughout the flight envelope.

The system is composed of:

• A variable stop unit consisting of an articulated lever operated by anelectromechanical actuator and a transducer unit detecting leverposition. These items are all mounted on a frame assembly locateddownstream of the differential between the AP and yaw damperactuators.

• Two control and monitoring computers designated FLC (Feel andLimitation Computer).

• One RUD TRAVEL control panel, one PITCH FEEL & RUDTRAVEL maintenance panel and five electrical power supply circuitbreakers.

1. VARIABLE STOP ACTUATOR DESCRIPTION• Two AC motors, supplied with 26V-400 Hz• A single reduction gear actuated by both motors, which are

rigidly connected

• A nut/screw system, driven by means of a torque limiter• Mechanical end-of-travel stops• A torque limiter provided to protect the reduction system from

any abrupt jamming of the output shaft, particularly when itreaches the mechanical stop.

2. TRANSDUCER UNITThe actuator is servo controlled and is monitored through atransducer unit driven by variable stop lever movement. Thetransducer unit, comprising two inductive transducers, is identical tothe one used in the spoiler control system.

3. SPRING - RETENTION RODIn the event of a rupture or disconnection of an actuator attachment,a retention rod limits actuator movement to prevent it from jammingthe variable stop lever. A spring returns the lever to the "low speed"position where full control deflection (+30) is possible.


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Rudder System - Rudder Travel Limiting

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Rudder Travel and Pitch Feel Systems - Feel and Limitation Computer (FLC) - General

This computer contains the circuitry required for two functions: ruddertravel limiting and pitch feel. The FLC is a digital computer comprisingtwo different computation channels:• Rudder travel limiting/pitch feel control channel• Rudder travel limiting/pitch feel monitor channel

Safety of the systems is ensured by:• control and monitor channel programs which are intentionally

different• monitoring of digital computations which are performed by control

and monitor channels with the same input data, achieved bycomparison between the results of both channels, by means ofanalog comparators

• power loop monitoring achieved by software means in each digitalchannel.

If any indicator is on, the test of either RUDDER TRAVEL LIMITINGsystem or PITCH ARTIFICIAL FEEL system will not operate.


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Rudder System and Pitch Feel - Feel and Limitation Computer (FLC 1/2)

FLC1/FIN 302CY1FLC1/FIN 302CY1FLC2/FIN 302CY2FLC2/FIN 302CY2


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ELEVATOR SYSTEM PITCH CONTROL

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Elevator System - Pitch Control

Pitch control is achieved by two elevators hinged on the horizontalstabilizer, each actuated by three servo controls controlled by a dualmechanical linkage through dynamometric rods, cable runs, an artificialfeel system linked to the cable run of the LH control column, and loadlimiting rods. In normal operation the two elevators are controlledtogether. In case of jamming in one control linkage during flight (take offexcluded), pitch control is provided by THS (Trimmable HorizontalStabilizer). If jamming occurs at take off, two uncoupling bellcranksenable the elevator on the other side to be controlled by one or bothpilots.

A pitch uncoupling unit (locking rod plus solenoid) prevents accidentalasymmetrical deflection of the elevators during flight and allowsuncoupling of the RH and LH control systems during take off (locked atspeeds lower than 30 kt or higher than 195 kt). Artificial feel is providedby the associated action of:• a double action spring loaded rod• a torsion bar driven by a variable gain mechanism which generates a

variable stiffness in the control. The variable gain mechanism isactuated by either of two electrohydraulic actuators. Each actuator iscontrolled by an independent PITCH FEEL channel, each oneincluded in a FLC (Feel and Limitation Computer).

PITCH FEEL systems are operative above 165 kt. Inputs are a functionof stabilizer position, airspeed and Mach number. In case of failure of twosystems, the mechanism returns to the sow speed position In each run,

downstream of the artificial feel system, a load limiting spring rod limitsthe efforts in the elevators control linkage. A spring loaded rod on eachservo control input avoids a runaway of the elevator in case of jammingof one input lever in the open position. An autopilot actuator is mountedadjacent to the LH elevator. It drives the control via a detent lever whichcan be overridden by the pilots.

Pitch trim is provided by adjustment of the horizontal stabilizer from +3°(nose down) to -14° (nose up). It is actuated by a fail safe ball screw jackdriven by two independent hydraulic motors supplied respectively byGreen and Yellow systems and coupled by a differential gear throughpressure-off brakes. Horizontal stabilizer adjustment may be initiated:• manually (AP disengaged) by trim wheels operation (mechanical

mode) or by action of the control wheel rocking levers (electricalmode).

• automatically by AP trim, mach trim or alpha (angle of attack) trimfunction.

Electrical and automatic trim signals are processed in two FAC (FlightAugmentation Computers) and control two electrical motors. Trim speedand trim authority depend on trim mode and aircraft configuration. Themotors drive the control linkage to the hydraulic valves which control thehydraulic motors. The manual trim wheel run is connected to the samelinkage. Stall warning is provided by a stick shaker (electrical motor)which is installed on each control column, and controlled by the FWC(Flight Warning Computer)


ATA 27ATA 27A300/A310A300/A310

27-27-5757

Elevator System - Pitch Control - Diagram


ATA 27ATA 27A300/A310A300/A310

27-27-5858

Elevator System - Elevator Mechanical Control System - General and Components

A. GENERAL

Each elevator is operated by three mechanically controlled servocontrols. The inputs from the control columns are transmitted to theelevators by dual control systems. Each system is routed along one sideof the fuselage. The left and right systems are interconnected at twopoints by detent bellcranks, one beneath the flight compartment floor, theother between the two elevators.

B. COMPONENT DESCRIPTION

1. Cable tension regulators maintain a constant tension on the cables(49.50 ±9.23 lbf).

2. The servo control actuating spring rods:• Provide flexibility in the control for any asymmetrical deflection of

the elevators in ground gusts• Prevent runaway of the control system if an input lever jams on

its servo control body.

3. A load limiting spring rod in each system, downstream of the artificialfeel unit, limits the design loads.

4. Rods: Identical to those used in the aileron control system

5. Cables: Identical to those used in the aileron control system

6. Control column stops: Control column travel is limited in bothdirections by non adjustable stops.Elevator operational stops: Maximum input to the servo controls islimited by adjustable stops located at a lever, close to each elevator.Elevator travel stops: These are the stroke end stops(non-adjustable) of the servo controls, never reached in normaloperation.Elevator structural stops, when the servo controls are not installed:The elevators rest on structural down stops, designed for thatpurpose, which are not able to withstand any load other than theweight of the elevators.Adjustable levers: The length of a lever close to each elevator isadjustable in order to maintain maximum travel of the elevatorswithin the design limits.

NOTE: Rigging pin holes are provided at convenient places to facilitaterigging.


ATA 27ATA 27A300/A310A300/A310

27-27-5959

Elevator System - Elevator Mechanical Control


ATA 27ATA 27A300/A310A300/A310

27-27-6060

ARTIFICIAL FEEL (ELEVATOR)

The pitch artificial feel system creates load feel at the control columnwhich is variable with flight conditions, in order to reduce the variation offorce per g throughout the whole flight envelope. At high angle of attack,the system causes an increase in the load feel at the control columnresulting in aircraft return to permissible angle of attack configuration.

1. Pitch Artificial Feel Unit

The artificial feel unit is composed of:• spring box providing a force threshold• A torsion bar driven by a variable gain mechanism which

generates variable load feel.• Two electrohydraulic actuators, displacement of which produces

the kinematic gain variation.• One return spring box used to retract the two actuators to the

position corresponding to "low speed" load feel, in the event ofdouble hydraulic failure.

Pitch Artificial Feel Unit - Operation

The actuators act on the gain variation mechanism by means oflevers. Gain is imposed by the actuator having extended the furthest.In the event of jamming of the mechanism, a microswitch transmits apitch disagree warning signal. The artificial feel unit includes a "failsafe" part to avoid loss of the force threshold and feel load at thesame time.

2. Pitch Artificial Feel Actuator

Each actuator includes:• A biased servovalve which modulates pressure in the actuator

large chamber, the small chamber being permanently suppliedwith high pressure. In the event of an electrical failure, servo-valve current is nulled and its control valve is displaced so thatthe actuator is retracted.

• A solenoid valve, energized in normal operation• A bypass which connects the large chamber to return in order to

retract the actuator when the solenoid valve is de-energized. It istherefore redundant with respect to the servovalve bias.

• A position pickoff potentiometer.

3. Pitch Upcoupling Unit

Provides connection of LH and RH elevators from 0-30 knotsairspeed during takeoff roll.Above 30-195 knots, the LH and RH elevators are disconnected byADC 1/2 to allow either pilot to control the elevator (pitch) function. Incase of a jam in the elevator control system on the captain’s or firstofficer’s control panelsAbove 192 knots, both elevators will reconnect for full control of theLH and RH elevator system runs.During landing conditions, this process is repeated in the sameairspeed conditions.

4. Aft Detent Bellcrank

In the event of an elevator jam on the LH or RH elevators, the aftdetent bellcrack will release on the jammed side to prevent lockout ofthe elevator system.


ATA 27ATA 27A300/A310A300/A310

27-27-6161

Elevator System - Pitch Artificial Feel - Components/Location

1

3

2

4

Artificial Feel Unit


ATA 27ATA 27A300/A310A300/A310

27-27-6262

Elevator System - Pitch Uncoupling System - General and Components

The two elevator control channels can be uncoupled during the takeoffphase in the event of jamming at any point on the control systems, bymeans of two detent bellcranks; one installed between the controlcolumns, the other between the two elevators. A pitch uncoupling unit,comprising a solenoid and rod, prevents any inadvertent uncoupling ofthe two elevators after the takeoff phase in order to prevent asymmetricalloads being applied to the structural attachments of the trimmablehorizontal stabilizer. The uncoupling unit solenoid is energized if airspeedVc is higher than 30 kts and lower than 195 kts.

COMPONENT DESCRIPTION

1. SolenoidThe solenoid includes the following components:• A low resistance draw coil, allowing high intensity current to

provide a high draw force when the coil is energized.• A high resistance holding coil allowing low intensity current to

provide permanent operation capability of the solenoid.• Two end of stroke switches, one for direct draw coil energization,

one for test purposes.• A return spring, to lock the rod when the solenoid is

de-energized.

NOTE: The lower limit of 30 kts (minimum speed for which a Vc valuecan be obtained from ADCs) has been introduced to prevent permanentenergization of the solenoid and power contactor coil when the aircraftelectrical network is energized on the ground.


ATA 27ATA 27A300/A310A300/A310

27-27-6363

Elevator System - Pitch Uncoupling System - Schematic


ATA 27ATA 27A300/A310A300/A310

27-27-6464

Trimmable Horizontal Stabilizer System - Pitch Controls

Both pitch trim wheels provide mechanical control of the TrimmableHorizontal Stabilizer (THS). When a pitch trim control wheel is used tooverride the electrical command, it disengages the electric actuators andthe PITCH TRIM levers trip to OFF. The trim range is from 14° nose upto 3° nose down. Trim position is indicated in degrees on a scaleadjacent to each trim wheel which is painted Green over the normal takeoff range (2° DN. 2.5° UP).

On each control wheel a rocking lever for pitch trim control is installed.Up or down movement of the rocking levers activates the two electricactuators which control the hydraulic motors for horizontal stabilizeradjustment providing that at least one PITCH TRIM system is engagedand AP is OFF or in CWS mode. The rocking levers are spring loaded toneutral position. If both rocking levers are operated simultaneously, but inopposite position, trimming action stops. If trimming by means of therocking levers lasts for more than 1 sec., an aural warning is activated.

NOTE: The pitch trim rate is:• 0 . 9 ° /s when the speed is below 200 kts.• 0.17°/s when the speed is above 240 kts. It varies

linearly from 0.9°/s to 0.17°/s when the speed is between200 and 240 kts.


ATA 27ATA 27A300/A310A300/A310

27-27-6565

Trimmable Horizontal Stabilizer System - Pitch Controls


ATA 27ATA 27A300/A310A300/A310

27-27-6666

Elevator System - Elevator Surface Position Indicating

Position of the right elevator is indicated on the right SCAM display unit,with the hydraulic systems available for the servo controls. There is nospecial reference mark painted on the elevators, but on each side of theAPU tailcone, there is:

• an engraved reference plate which indicates the neutral position ofthe corresponding elevator.

• an engraved placard with the following inscription:VALID STABILIZER IN NEUTRAL POSITION.

1. ELEV AND STAB POSITION INDICATION

A White scale covering the full travel range is provided for elevatorand trimmable horizontal stabilizer position. An index indicating theactual position of the surfaces moves along each scale. In addition,each available hydraulic system on the THS is indicated by a Greensymbol (G,Y). In case of servo control low pressure detection, thecorresponding symbol becomes Amber.


ATA 27ATA 27A300/A310A300/A310

27-27-6767

Flight Controls System - RH ECAM Page - System Display

Elevator and Horizontal Stabilizer Position IndicationElevator and Horizontal Stabilizer Position Indication


ATA 27ATA 27A300/A310A300/A310

27-27-6868

Pitch Feel and Trim System - Control Switches and Levers

1. PITCH FEEL SYS 1 AND 2 PUSHBUTTON SWITCHES

The P/B switches control channels 1 and 2 of the Feel and LimitationComputers (FLC) for elevator control.

• ON (P/B switch pressed-in): The corresponding system isengaged. Both systems may be engaged simultaneously butonly one is effectively operating. If one system fails, it isautomatically deactivated and the other one continues tooperate.

• OFF/R (P/B switch released-out): The OFF/R light comes onWhite and the system involved is disengaged. The monitoringcircuits are reset by this action. This indication is accompaniedby SCAM activation.

• FAULT: When a P/B switch is pressed-in, the associated FAULTlight comes on Amber if a failure is detected in the correspondingsystem. Illumination of the Amber FAULT light is accompaniedby ECAM activation.

Both FAULT lights remaining illuminated when the P/B switchesare released-out and the OFF/R lights are illuminated White,constitutes a pitch disagree warning (The artificial feel unitoperates in high speed configuration when flaps are extended20° or more ) . Illumination of both FAULT lights is accompaniedby ECAM activation.

2. PITCH TRIM 1 AND 2 LEVERS

• 1 (or 2): The lever is magnetically latched in the active positionand the pitch trim 1 (or 2) is engaged. If a failure is detected, thecorresponding PITCH TRIM lever trips to OFF.

• OFF: The respective pitch trim is disengaged.

- When one PITCH TRIM lever trips to OFF, all electricalcontrol modes of the THS are lost and the ECAM isactivated.

- When both PITCH TRIM levers trip to OFF, all electricalcontrol modes of the THS are lost and the SCAM isactivated.

NOTE: Pitch trim disengages and the levers drop to OFF when trimreaches full nose up or full nose down position (mechanical stops).


ATA 27ATA 27A300/A310A300/A310

27-27-6969

Flight Deck Pitch Feel and Trim System - Control Switches and Levers

BA


ATA 27ATA 27A300/A310A300/A310

27-27-7070


LEFT BLANK


ATA 27ATA 27A300/A310A300/A310

27-27-7171

TRIMMABLE HORIZONTAL

STABILIZER (THS) -

SECTION


ATA 27ATA 27A300/A310A300/A310

27-27-7272

Trimmable Horizontal Stabilizer System - General

Pitch trim control is achieved by a trimmable horizontal stabilizer (THS)hinged on the rear part of the fuselage. The two elevators are hinged onthe THS. Their control systems are installed so that the elevators are inline with the THS when the control columns are released. The THS isdriven by an actuator including a fail-safe ball screwjack, the structuralattachments of which are also fail-safe. Normal control of the actuator iselectrical, via the automatic pitch trim system.

Stand-by controls are mechanical. The pilots can override electricalcontrol by the mechanical control system by applying sufficient force tothe control wheels. A torque limiter is mounted in each electric pitch trimactuator for that purpose. The torque limiters remain automaticallyreleased after their operation.

INTERFACE WITH AUTOMATIC PITCH TRIM SYSTEM

Electrical control is achieved by means of two electric pitch trimactuators, installed on the THS actuator. They drive the actuatormechanical input. The two pitch trim actuators are controlled by two flightaugmentation computers (FAC) which deliver manual electric trim,automatic trim, Mach trim and alpha trim signals. Manual electric trimsignals are provided by rocker switches mounted on the Captain's andFirst Officer's control wheel horns. Electric limit switches detect THS endof travel in aircraft nose up direction. The signals are used in theautomatic trim system to avoid automatic disconnection of this systemduring automatic landings.


ATA 27ATA 27A300/A310A300/A310

27-27-7373

Trimmable Horizontal Stabilizer System - Diagram


ATA 27ATA 27A300/A310A300/A310

27-27-7474

Trimmable Horizontal Stabilizer System - Hydraulic Actuation - Components/Location

The trimmable horizontal stabilizer is driven by an actuator whichincludes two hydraulic motors, each powered by a different hydraulicsystem.


ATA 27ATA 27A300/A310A300/A310

27-27-7575

Trimmable Horizontal Stabilizer System - Hydraulic Components/Location


ATA 27ATA 27A300/A310A300/A310

27-27-7676

Trimmable Horizontal Stabilizer System - Actuator - General

The THS actuator consists of a fail-safe ball screwjack actuated by twohydraulic motors coupled by a differential gear.

NORMAL OPERATION

Pressure-off brakes (9) are released. The ball screw is held by theno-back brake formed by items (14) (15) (16) (17). Rotation of input shaft(1), driven either by one of the electrical pitch trim actuators or by

the mechanical input, controls rotation of the ball screw through twoidentical control loops, including input and feedback gear trains,feedback differentials (4), control valves (5), hydraulic motors (8) andactuate a power gear train through power differential (10). The controlstroke is limited by the actuator input shaft stop (2). The structuralcomponents (ball screw and nut assembly, attachments to THS andfuselage) and the power gear train are duplicated, the secondary loadpath being normally unloaded.


ATA 27ATA 27A300/A310A300/A310

27-27-7777

Trimmable Horizontal Stabilizer System - Actuator - Schematic


ATA 27ATA 27A300/A310A300/A310

27-27-7878

WARNING LOGIC PITCH CONTROL SYSTEM

Depicted below are the various warnings displayed in the Flight Compartment in theevent the fault shown occurs in the Pitch Control System of the aircraft. Alsographically shown are the Flight Phases at which the warnings will or will not bedisplayed.


ATA 27ATA 27A300/A310A300/A310

27-27-7979

Warning Logic - Pitch Control System


ATA 27ATA 27A300/A310A300/A310

27-27-8080

Alpha Probes System - Stall Warning - General

• The dual stall warning system provides audio (cricket) and vibrating(stick shaker) warning in case of impending stall.

• The angle of attack is the governing parameter for stall warning,together with slat extension.

• The angle of attack is given by two alpha probes (one on each sideof the forward fuselage) which are electrically heated. Slat position istransmitted by two synchro-transmitters, one for each FWC.

• On each control column, a stick shaker is installed and controlled bythe stall warning generator included in FWC 1 or 2.

• Stall warnings are activated when angle of attack exceeds apredetermined value.

• slat retraction is inhibited.

• turn coordination of yaw damper is inhibited.

Stall warning is inhibited on the ground except during ALPHA PROBEStest.


ATA 27ATA 27A300/A310A300/A310

27-27-8181

Alpha Probes System - Stall Warning - Block Diagram

Captain

Captain

First Officer

First Officer


ATA 27ATA 27A300/A310A300/A310

27-27-8282

Flight Control System - Stick Shaker - General

One stick shaker is installed on each control column, and is controlled bythe Flight Warning Computers 1/2.


ATA 27ATA 27A300/A310A300/A310

27-27-8383

Flight Control System - Stick Shaker


ATA 27ATA 27A300/A310A300/A310

27-27-8484


LEFT BLANK


ATA 27ATA 27A300/A310A300/A310

27-27-8585

LEADING EDGE LIFT DEVICES

SECTION


ATA 27ATA 27A300/A310A300/A310

27-27-8686

Wing Leading Edge - Slat System - General

SLAT SYSTEM

There are three slat surfaces in each wing, the inboard, center andoutboard slats. They are guided on curved support tracks. The inboardslat has three tracks and the center and outboard have four each. Afolding nose on each inner slat folds to clear the engine pylon when theslats extend. The slats are actuated by ballscrew jacks, two for eachsurface. Two friction brakes, one at each end of the wing transmissionsystem provide system irreversibility. Attached to each friction brake is aposition pick-off unit (PPU) for asymmetry and system monitoring.

KRUEGER FLAP AND NOTCH FLAP

The Krueger flap and notch flap are provided to complete the wingleading edge profile when the slats are extended. The Krueger flap andnotch flap are operated by individual hydraulic actuators. Both arecontrolled by the SFCC and move to the extend and retract positionwhen the SFCC commands slat extension or retraction.

Slats and spoilers are numbered from inboard to outboard, each sideseparately.


ATA 27ATA 27A300/A310A300/A310

27-27-8787

A300 Wing Leading Edge - Slat System Components - Location

A300SHOWN/A310 DONOT HAVENOTCH (ORSLOT) FLAPS


ATA 27ATA 27A300/A310A300/A310

27-27-8888

Wing Leading Edge - Slat System - Hydraulic Operation - Diagram

The Power Control Unit (PCU) hydraulic supply is provided by the aircrafthydraulic systems. The slat system No. 1 is supplied by the Blue system,the slat system No. 2 is supplied by the Green system. If there is a singlesystem failure the system will still operate but at half speed.


ATA 27ATA 27A300/A310A300/A310

27-27-8989

Wing Leading Edge - Slat Hydraulic System - Schematic

** **

* NOTE: A300 ONLY/A310 DO NOT

HAVE NOTCH OR SLOT FLAPS


ATA 27ATA 27A300/A310A300/A310

27-27-9090

Wing Leading Edge - Slat Control System - General

The power is supplied to the ball screwjacks by a torque shaft driven by apower control unit and protected by a system torque limiter for eachwing. Each ball screwjack also has its own torque limiter. All these torquelimiters include a latched lockout indicator and in case of overload of ajack in the torque shaft system, they will freeze the system until a reverseselection is attempted.

As soon as an order is given, the corresponding computer of each motorsends signals to deliver the pressure to the motor, releases the involved

pressure-off brake and controls the sense and speed of movement.When the selected position is reached, the systems are de-energized,applying the pressure-off brakes and stopping the movement.

In case of hydraulic failure, the corresponding motor remains locked byits brake and the operating speed of the slats is reduced by half due tothe differential mechanism of the power unit gearbox. However, fulltorque is still available. Three slat positions can be selected (0°, 15°, 30°)by the five position control lever.


ATA 27ATA 27A300/A310A300/A310

27-27-9191

Wing Leading Edge - Slat System Control and Indicating - Schematic


HAVE NOTCH OR SLOT FLAPS*

ADC 1

ADC 2


ATA 27ATA 27A300/A310A300/A310

27-27-9292

Wing Leading Edge - Slat - Hydraulic Power Drive System - General

The slat drive system comprises the power control unit (PCU), atransverse torque shaft system and the screwjacks. In the PCU twoindependent hydraulic motors, one controlled by the Blue valve block andone controlled by the Green valve block, drive a summing gear. Theoutput is passed through torque shafts to a tee-gearbox which rotates themotor drive direction by 90°. A pressure-off brake is provided betweeneach motor and the summing gear to lock the transmission system whenthe slat system is static.

The two transverse outputs drive the ballscrew jacks through torquelimiters, a series of torque shafts, steady bearings and gearboxes. Oneunidirectional friction brake is installed at each wing tip to provide systemirreversibility under compressive screwjack loads.


ATA 27ATA 27A300/A310A300/A310

27-27-9393

Wing Leading Edge - Slats Power Control Unit (PCU) Valve Block - Schematic

FWDFWD


ATA 27ATA 27A300/A310A300/A310

27-27-9494

Wing Slats and Flaps Systems - Indicating and Control - Description

A single control lever located on the center pedestal permits slat and flapcontrol. The lever has five gated positions. It is not possible to select anintermediate position (if the lever is held in between gates the systemdrives to the last demanded position and after 10 sec. all the slat and flapFAULT warnings illuminate). The slats and flaps are electrically signaledby two identical digital computers (Slats Flaps Control Computers). Eachone is composed of one slat control channel and one flap controlchannel.

SLATS (OR FLAPS) SYS 1 AND 2 FAULT LIGHTS

Each light comes on Amber when the associated hydraulic motor isinoperable. Both slats (or flaps) stop due to a system jam. In both cases,a reverse selection is possible. If system jam is released, the system willmove to the commanded position.

Both SLAT (or FLAP) FAULT lights and the associated Amber SLAT (orFLAP) lights on the Slat/Flap Position Indicator will come onsimultaneously if a mechanical failure is detected. In this case, thesystem is locked by the pressure-off brakes and there is no possibility ofrecovery in flight.

NOTE: If a SFCC is not installed, the two associated FAULT light (oneSLAT FAULT light and one FLAP FAULT light) will come on.

Illumination of these lights is associated with ECAM activation.

1. SLAT/FLAP POSITION INDICATOR STRIPSSlat and flap positions are shown by White strips moving up anddown associated scales. The corresponding VFE (speed limit) isplacarded opposite each normal position (indicated by a roundnumber).

2. SLAT AND FLAP LIGHTSCome on Amber when the associated system is blocked.

3. KRUEGER LIGHTComes on Amber if either KRUEGER flap is not in correct position10 sec. after a movement command. Illumination of KRUEGER lightis accompanied by ECAM activation.

4. SPD BRK LIGHTThe light comes on Blue when the speed brake control lever is not inRET position.

5. Flashes Blue when the slat lock function is activated (inhibition ofcomplete slat and KRUEGER flap retraction at high angle of attack).

Before selection of any position, the slat/flap control lever must be pulledup. A block is provided for positions 2 and 4 to prevent the lever movingstraight through.

NOTE: All slat and flap FAULT lights will illuminate if the control leverremains between two gated positions (after 10 sec.).

NOTE: All Slat and Flap FAULT lights will remain illuminated if Slat/FlapControl Computer (SFCC) number 1 and 2 are removed from ElectronicRack 90VU. The 28 Volt DC Interlock Relay in the system preventsaccidental dispatch of the aircraft if both SFCC 1/2 are removed from theelectronic rack or aircraft.


ATA 27ATA 27A300/A310A300/A310

27-27-9595

Wing Slats and Flaps Systems - Indication and Controls

A

B

C

0

15

15

15

30

0

0

15

20

40

FLAPSSLATS


ATA 27ATA 27A300/A310A300/A310

27-27-9696

A300/A310 Wing Leading Edge - Slat - Asymmetry Monitoring

Slat position is indicated to the flight crew by the SLATS vertical bardisplay on the slat/flap position indicator. The display is driven by inputsfrom the instrumentation Position Pick-off Unit (PPU) which also providesindependent slat position information for other systems.

Slat position discrete and digital data are provided for other systems bythe SFCCs.

ASYMMETRY AND POWER TRANSMISSION MONITORING( SLATS)

The Asymmetry PPUs enable the SFCCs to monitor the transmission forasymmetry and runaway conditions.

If an asymmetry or runaway condition is detected, the PCU operation isinhibited, preventing further movement of the transmission system.


ATA 27ATA 27A300/A310A300/A310

27-27-9797

A300/A310 Slat Asymmetry Position Pick-Off Unit and Adapter - Component Location

B

B

B


ATA 27ATA 27A300/A310A300/A310

27-27-9898

Wing Leading Edge - Krueger and Notch Flap Actuation

When the slats are extended, SFCC1 and SFCC2 send an extenddiscrete signal to a Krueger selector solenoid valve, which has an extendand retract solenoid. When the extend solenoid is energized, hydraulicpressure is passed to the Krueger and notch flap actuators and to the allspeed aileron system. The Krueger and notch flaps extend. The solenoidremains energized.

When the slats are retracting and have passed the 15° position, theretract solenoid is energized, the extend solenoid is deenergized and theKrueger and notch flaps are retracted.

KRUEGER AND NOTCH FLAP ACTUATION

The Krueger selector solenoid valve is located in the hydraulic bay atFR47. It is a three-position, four-port solenoid-operated shuttle valve.The shuttle valve is springloaded to center. It moves to the centerposition when both solenoids are deenergized. In this position, thehydraulic pressure input in A is shut off and ports B, C and D areinterconnected.

INTERFACE WITH THE AILERON SYSTEM

When the Krueger and notch flaps are supplied with pressure from theKrueger selector solenoid valve, so also is the droop actuator in the allspeed aileron system. When the Krueger and notch flaps are extended,the ailerons droop 9.2°. On retraction, the ailerons return to their normalpositions.

NOTE: The A310 Slat System does not have Notch or Slot Flap devices.


ATA 27ATA 27A300/A310A300/A310

27-27-9999

Wing Leading Edge - Krueger and Notch Flap Hydraulic Actuation - Schematic

A300 ONLY


ATA 27ATA 27A300/A310A300/A310

27-27-100100

Wing Leading Edge - Krueger and Notch Flap - Components

KRUEGER FLAP ACTUATOR

The Krueger flap actuator is a double-acting actuator with mechanicallocking of the piston assembly in the extended position (flap retracted)and hydraulic locking in the retracted position (flap extended). The pistonend is connected to the Krueger flap by a reverse link. The cylinder headis mounted to the structure by two mounting blocks which allow theactuator to pivot on the mounting during the operating cycle. Theactuator consists of a cylinder:. f head a cylinder and a valve block.

*NOTCH FLAP ACTUATOR / A300 ONLY

The notch flap actuator is a double-acting actuator and is hydraulicallylocked in the retracted position. The piston rod is connected by aneye-end to the notch flap and the cylinder is attached to the structure bya shaft hinge to allow some pivoting during the retraction and extensioncycle.

SAFETY VALVE

The safety valve is located in the pressure line from the Green hydraulicsystem to the solenoid selector valve. Its function is to prevent loss ofhydraulic fluid from the Green system should there be a major rupture inthe Krueger and notch flap actuating system.

*NOTE: The A310 Slat System does not have Notch or Slot Flapdevices.


ATA 27ATA 27A300/A310A300/A310

27-27-101101

Wing Leading Edge - Krueger and Notch Flap - Components

**


HAVE NOTCH OR SLOT FLAPS


ATA 27ATA 27A300/A310A300/A310

27-27-102102

Wing Leading Edge - Krueger and Notch Flap Control and Monitoring - General

In addition, to obtain better aerodynamic characteristics, a KRUEGERflap and a NOTCH flap are provided on each wing and are locatedbetween the inner slat and the fuselage.

The KRUEGER and NOTCH flaps are extended when the slat/flapcontrol lever is moved from position 1 to 2 and remain extended for allother selected positions. When slats 0° position is selected, theKRUEGER flaps fold up under the leading edge and the NOTCH flapretracts into the fuselage.

Each KRUEGER flap and AIL droop actuator and each NOTCH actuatorare supplied from a KRUEGER selector solenoid valve supplied by theGreen circuit and controlled by the slats control system.

The KRUEGER jacks are mechanically locked in retracted position andhydraulically locked in extended position.


ATA 27ATA 27A300/A310A300/A310

27-27-103103

Krueger/Notch Flap Control and Monitoring - Schematic

** * NOTE: A300 ONLY/A310

DO NOT HAVE NOTCH ORSLOT FLAPS


ATA 27ATA 27A300/A310A300/A310

27-27-104104

Wing Leading Edge - Slats/Krueger Flap Monitoring and Fault Warning System

The SFCC provides continuous monitoring of the slat system and theKrueger and notch position. Fault warnings are generated for those faultsrequiring pilot action or flight crew awareness. The faults are stored inthe SFCC including those which are purely maintenance data. Thewarnings are displayed on one or more of the following:

• SLATS SYS 1 FAULT or SLATS SYS 2 FAULT annunciator (19CV)on the overhead panel

• BITE DISPLAY/SFCC1 or SFCC2 annunciators (52CV and 53CV) onthe FLIGHT CONTROL section panel 471VU

• Left electronic centralized aircraft monitor (ECAM)• Fault indicator on the SFCC front panel.

SELF TEST

Self test facilities are provided to:

• detect and indicate failure in redundant and dormant circuits• identify a failed line replaceable unit (LRU)

The self test can be initiated from the flight deck maintenance panel or byusing the BITE pushbutton switch on the SFCC front panel.


ATA 27ATA 27A300/A310A300/A310

27-27-105105

Fault Indicators and Flight Controls Test Panel - Krueger System

A300 Indicator Shown

1/FIN 21CV2/FIN 22CV

MAINT PANEL 471VU

A300/SOME A310


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Warning Logic Slat System

Depicted below are the various warnings displayed in the Flight Compartment in the event the faultshown occurs in the Slat System of the aircraft. So shown graphically are the flight phases at whichthe warnings will or will not be displayed.


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A300/A310 Slats System - Warning Logic


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LEFT BLANK


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WING FLAPS

SECTION


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A300 Wing Flap System - Description

Each wing has three flap sections. The three flaps are single slottedfowler type and are guided by two tracks fitted with ball screwjacks. Thepower is supplied to the ball screwjacks by a torque shaft driven by apower control unit and protected by a system torque limiter for eachwing. Each ball screwjack has its own torque limiter.

All these torque limiters include a latched lock-out indicator and, in caseof overload of a jack in the torque shaft system, they will freeze thesystem until a reverse selection is attempted. The Power Control Unit(PCU) consists of two independent hydraulic motors coupled to adifferential mechanical system through pressure-off brakes that insurethe system irreversibility. The two motors are supplied by differenthydraulic circuits (Green and Yellow).

As soon as an order is given, the corresponding computer of each motorsends signals to deliver the pressure to the motor, release the involvedpressure-off brakes and control the direction and speed of movement.When the selected position is reached, the systems are de-energized,applying the pressure-off brakes and stopping the movement.

In case of hydraulic leakage, the corresponding motor remains lockedand the operating speed of the slats is reduced by half due to thedifferential mechanism of the power unit gear box. However, full torque isstill available. Four flap positions can be selected (0°,15°,20°,40°) bymoving the slat flap control lever from position 2 to 5. Furthermore,between the inboard and center flaps, there is an aileron droop signalunit which commands the aileron to droop 9.2° maximum with slatsextension to 15°.

A load relief system is provided to minimize the design loads on the flapsupport structure and the flap jacks. Load relief function can only engagewhen the slat flap control lever is in gate 5. Load relief is activated withinthe flap channels of the two SFCCs by using Calibrated Air Speed (CAS)received from the two ADCs. Load relief logic is the following:

• if CAS >178 kt, Flaps retract from 33.5° to 24°

• if CAS <173 kt, Flaps extend from 24° to 33.5°


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A300 Wing Flap System - Components/Location


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Wing Flap System Hydraulics

The PCU hydraulic supply is provided by the aircraft hydraulic systems.One motor is supplied from the Green system and one is supplied fromthe Yellow system. If there is a single-system failure, one of the motorswill drive the system at half speed.


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Wing Flaps Hydraulic System - Schematic

POWER CONTROL UNIT (PCU)

INSTALLATION LOCATION -

FUSELAGE FRAME (FR) 54


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Wing Flap System - Power Control Unit (PCU)

The flap drive system consists of two identical hydromechanical systems,comprising two hydraulic motors, a differential gear and pressure-offbrakes, contained within a power control unit, (PCU) and a transversetorque shaft system driving the ballscrew jacks. Static and dynamicsystem irreversibility is provided by no-back friction brakes withinscrewjacks 2, 4, 5 and 6. Screwjack 3 incorporates only one no-backwhich operates only during flap extension under tensile loads. Screwjack1 is not provided with no-back friction brakes.


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Wing Flap System PCU - Schematic

FWDFWD


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Wing Flap PCU Components

The PCU is located in the hydraulic compartment and provides thedriving power to the flap system.

It consists of the following subassemblies:• Acceleration control valve• Valve blocks• Transfer tubes• Hydraulic motors• Pressure-off brakes• Differential gearbox• Intermediate gearbox• Position pickoff units (PPU)


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Wing Flap PCU - Components and Valve Block

FWDFWD


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Wing Flap Controls

The flap system is controlled by two identical slat flap control computers(SFCC1 and SFCC2). The SFCCs provide, in addition, monitoring andtest facilities.

Included in the monitoring is a flap relief function in which the flaps areautomatically retracted from a fully extended position should the aircraft

exceed the maximum flap extended speed (VFE) limitation for the flapconfiguration. If the control lever setting has not been altered, the flapswill automatically extend when the aircraft speed has been reducedsufficiently. The air speed data is provided by the air data computer(ADC).


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A300 Wing Flaps - Control System - Schematic


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Wing Flap Controls

The flap system is controlled by two identical slat flap control computers(SFCC1 and SFCC2). The SFCCs provide, in addition, monitoring andtest facilities.

Included in the monitoring is a flap relief function in which the flaps areautomatically retracted from a fully extended position should the aircraft

exceed the maximum flap extended speed (VFE) limitation for the flapconfiguration. If the control lever setting has not been altered, the flapswill automatically extend when the aircraft speed has been reducedsufficiently. The air speed data is provided by the air data computer(ADC).


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A310 Wing Flaps - Control System - Schematic


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Warning Logic Flap System

Depicted below are the various warnings displayed in the Flight Compartment in the event the faultshown occurs in the Flap Control System of the aircraft. Also graphically shown are the Flight Phasesat which the warnings will or will not be displayed.


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A300 Flap System - Warning Logic


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LEFT BLANK


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SPEEDBRAKES AND SPOILERS

SECTION


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Speed Brakes and Ground Spoilers

SPEEDBRAKES

There are two inner and three outer speedbrakes located on the uppersurface of each wing. The outer speedbrakes are also used as rollspoilers. They are selected by a lever situated on the center pedestal.Eleven positions can be selected from RET (retracted) to FULL (fullyextended) position.

Each speedbrake element is powered by one servo control whichreceives pressure from an electrohydraulic valve group controlled by thecorresponding units of the EFCUs (Electrical Flight Control Units) andalso used for roll spoilers control.

"SPEED BRAKES EXTENDED" indication is given on the ECAM MEMOpage.

GROUND SPOILERS

Speedbrakes and roll spoiler surfaces are used on the ground as groundspoilers. Deflection angles become 50° for all surfaces. They areautomatically extended when:• they are selected.• the aircraft is on the ground.

The ground spoilers are selected when the two following conditions arefulfilled:• SPEED BRAKE control lever pulled upwards (when it is in RET

position) or thrust reverser selected on one engine.• both engine throttle levers in idle position.

The "aircraft on ground signal" is sent:• during takeoff or landing when the two main landing gear aft wheel

speed is higher than 85 kt• at landing, only if the ground spoiler preselection has been made by

speedbrake control lever selection, when:- main gear bogie beam is in the ground position.- radio altitude is lower that 5 ft.

These signals are inhibited 3 seconds after first shock absorbercompression.

Automatic extension is also achieved for an aborted take off when theseconditions are fulfilled. Ground spoilers will remain extended duringbounces as long as both throttles are in the idle position and the SPEEDBRAKE control lever is pulled up.

Ground spoiler retraction after landing is achieved:• Either by pressing the SPEED BRAKE control lever down

(preselection cancellation)• or by moving one throttle lever out of the idle position.


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A300/A310 Speed Brakes and Ground Spoilers - Schematic


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Speed Brake Control Lever - Description

A control lever assembly located in the center pedestal enables the crewto select the position of the speed brake surfaces arid to preselect theground spoiler function.

The control assembly is composed of:• A control lever which drives the speed brake control transducer unit

via mechanical linkage.• A fixed quadrant with eleven notches which lock the SPEED BRAKE

control lever in the selected position.

Lever motion is guided by a slot. True slot prevents ground spoilerpreselection if the control lever is not in the RET position.

• A spring device which holds the control lever extended when theground spoiler function is preselected.

• A cam which drives the ground spoiler preselection microswitches.• A spring which holds the cam in the "ground spoiler not preselected"

position when the control lever is not pulled upwards.


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Speedbrake Control Lever - Selection Positions

SpeedbrakeSpoilers

RetractedSpeedbrake

SpoilersExtended

CAM

MIROSWITCHES

CONTROLLEVER

FIXEDQUADRANT

GUIDESLOT


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Speedbrake Controls

A. SPLR & SPD BRK PUSHBUTTON SWITCHES

1. Each P/B switch is associated with one or two pairs ofsymmetrical upper wing surfaces.

• ON (P/B switch pressed-in): The corresponding controlsystem is activated. Each time a system is activated, orcorresponding hydraulic system on, or the aircraft electricalnetwork is energized, a 2 sec. safety BITE test is triggeredfor the corresponding EFCU units (control and monitor).

• OFF/R (P/B switch released-out): The OFF/R light comes onWhite and the corresponding control system is deactivated. Ifhydraulic pressure is available, the actuators areautomatically held in the retracted position. The monitoringcircuits are reset by this action. This indication isaccompanied by ECAM activation.

• FAULT: When a P/B switch is pressed-in, the associatedFAULT light comes on Amber if a failure is detected by themonitoring circuits which then deactivate the control system.Illumination of the FAULT light is accompanied by SCAMactivation. The lever controls:- the position of the speedbrake eleven surface positions

from retracted (RET) to fully extended (FULL).- manual preselection of the ground spoiler function.

B. SPEEDBRAKE CONTROL LEVER

2. SPEEDBRAKE SELECTIONTo select the speedbrake surfaces to the required position, thepilot must press on the top of the lever and move it to thecorresponding notch. The control lever cannot be moved as longas the ground spoiler function is preselected.

3. GROUND SPOILER OPERATIONTo arm the ground spoiler function, the control lever must belifted when in retracted position (RET).


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Speedbrake Controls - Control Switches and Handle

AB


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LEFT BLANK


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MISCELLANEOUS FLIGHTCONTROL AREAS

SECTION


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Flight Control Surface Position Indicating - FLT CTL Page

The flight control page of the system display (right SCAM display unit) iscalled either by manual selection, or automatically when a warningoccurs on a system covered by this page.

1. B,G,Y SYMBOLSEach available system on the flight controls and trimmable horizontalstabilizer is indicated by a Green symbol. In case of servo control lowpressure detection, the corresponding symbols become Amber.

2. PRIMARY FLIGHT CONTROLS SURFACE POSITION DISPLAYA white scale covering the full travel range is provided for eachsurface (elevators, rudder, all speed ailerons and trimmablehorizontal stabilizer). An index indicating the actual position of thesurfaces moves along each scale.

3. ROLL SPOILER AND SPEEDBRAKE SURFACE POSITIONDISPLAYEach roll spoiler and speedbrake surface is represented as follows:

When the surface is completely retracted, it is symbolized by a smalldash.

When the surface is deflected more than 2°, a small arrow appearsabove the dash.

In normal conditions, these indications are Green. in case of failuredetected for a group of surface in the EFCUs, the numbersidentifying the surfaces involved appear below the dash.

In case of hydraulic failure when a surface is extended, the color ofthe corresponding symbol changes from Green to Amber and thenumber appears below the symbol.

On the ground, after landing, the arrow of any surface extendedflashes as long as speedbrake surfaces No. 1 and 2 are not fullyretracted.


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Flight Control Surface Position Indicating - FLT CTL Page

Aircraft banking to right


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Flight Control Surface Position Indicating - WHEEL Page

The RH ECAM WHEEL Page called up automatically at landingproviding that the Spoiler Preselection Conditions are valid on aircrafttouchdown on the runway (Spoiler display on bottom of WHEEL Page).

Each spoiler is represented in the same manner as on the FTL CTL(Fight Control) Page.

On the ground, the arrow corresponding to any spoiler extended afterlanding and the control system has not failed, flashes as long as Spoilers1 and 2 are not fully retracted.


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Flight Control Surface Position Indicating - WHEEL Page


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Description Of Control Surface Position Indicating For The Three Axes

On all three axes synchro transmitters send control surface positionsignals to the right ECAM display unit:• Right all speed aileron: 6CT• Left all speed aileron : 7CT• Rudder: 10CT• Trimmable horizontal stabilizer: synchro in position sensor: 11CT• Elevator (right): 12CT

Travel of the above surfaces is indicated by displacement of an index,symmetrically in the case of the ailerons. For the spoilers, pulses fromthe EFCU are displayed on the right ECAM display unit by illumination ofthe relevant spoiler arrows.


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Flight Control System - Control Surface Position Indicating Sensors - Location

• ASA (LH/RH Side) (2)• Rudder (1)• Elevators (1)


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A300/A310 Aircraft Takeoff (T/O) Configuration Test

When T/0 CONFIG TEST pushbutton is pressed and held, T/O powerapplication is simulated. This test will activate the appropriate warnings ifthe aircraft is not at take-off configuration.

The warnings are canceled when the pushbutton is released. FlightControls conditions for warning and resultant warnings are:

• HORIZONTAL STABILIZER is not in Take Off Configuration (>3°UP± 0.4 or > 2.3°DN ± 0.4). The Red T.O. CONFIG light comes onWLDP with associated CRC and ECAM activation.

• SLATS OR FLAPS are not in Take Of f Configuration. The Red T/OCONFIG light comes on WLDP with associated CRC and ECAMactivation.

• SPEEDBRAKES OR GROUND SPOILERS are extended. In thiscase, illumination of the Blue SPD BRK light on SFPI is accompaniedby the Red T/O CONFIG light on WLDP with associated CRC andECAM activation.

Pressing T/O CONFIG TEST pushbutton also monitors the followingsystems:• DOORS (when not closed)• LANDING GEAR (parking brake, brake temperature)• PROBE HEAT (Standby or CAPT or F/O probes heat off).


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A300/A310 Aircraft Takeoff Configuration Test


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Maintenance Panel - Flight Controls Test Panel Pushbuttons

Only used on the ground to test the jamming detection microswitches,with pressure for the respective circuit shut off.

1. B,G,Y TEST PUSHBUTTON SWITCHESAfter selecting the related SERVO CTL P/B switch to OFF (onoverhead panel), the TEST P/B switch for the respective circuit ismagnetically latched when pressed-in and the TEST light comes onWhite.

For jamming detection test, the controls involved must be movedrapidly. Successful test is indicated by flashing of the JAM light in therelated SERVO CTL P/B switch. If not successful, the fault isolationprocedure must be done on the face of the jamming detection controlbox.

After selecting the SERVO CTL P/B switch to normal, the TEST P/Bswitch is automatically released-out and the TEST light goes off.

2. TEST PUSHBUTTONSThe pushbuttons control the test of PITCH FEEL and RUD TRAVELelectrical systems and warning systems continuity. The test ispossible only if PITCH FEEL and RUD TRAVEL systems areengaged on the control panel (overhead panel).

Left pushbutton tests pitch feel and rudder travel system 1. Rightpushbutton tests pitch feel and rudder travel system 2. When a TESTpushbutton is pressed and held, the associated system mustdisengage and its FAULT light comes on Amber.

Successful test is indicated by White OK lights illumination.• upper lights for PITCH FEEL SYS 1 and 2• lower lights for RUD TRAVEL SYS 1 and 2.

3. OK LIGHTSThese lights illuminate White as long as the TEST pushbutton ispressed and held, to indicate a successful test.


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Maintenance Panel - Flight Controls Test Panel Pushbuttons

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Maintenance Panel - Flight Controls Test Panel - Controls and Indications

1. TEST SELECTOR• NORM FLT: Normal operating position, test circuits

disconnected, warnings canceled.• GND SPLR: Checks that no undue condition is permanently

achieved in the ground spoilers logic for the EFCU involved whenthe TEST P/B switch is pressed-in.- SEL LANE TEST 1 and 2 positions check the integrity of

selection lanes aircraft wirings.- EFCU TEST 3 and 4 positions check the integrity of each

EFCU logic. The corresponding FAULT lights on the SPLR &SPD BRK panel will go off. This test requires all hydraulicpower to be cut off to all flight controls to have the FAULTlights illuminated before test.

• PITCH CTL UNCOUPLING: Tests periodically, on ground, theelectrical circuits of the pitch uncoupling unit.- TEST 1 checks that the uncoupling unit rod is in the locked

position.- TEST 2 checks that the uncoupling unit moves to the

unlocking position, when the control solenoid is energized.• SLATS/FLAPS: Commands a BITE sequence for the relevant

SFCC (SYS 1 or SYS 2) when the TEST P/B switch is pressedin.

2. GND SPLR SEL LANE FAULT LIGHTThis light comes on White when a fault has been detected in TEST 1or TEST 2 positions of the test selector.

3. EFCU BITE DISPLAY LIGHTThis light comes on White when a fault has been detected by thecontinuous monitoring of each EFCU. More details of the failure aredisplayed of the face of the EFCUs.

4. PTT PUSHBUTTON SWITCHThis P/B switch activates the test of the system selected by the testselector. A TEST indication is integrated into the P/B switch.

• PTT: When pressed-in and held, the selected system is tested.

• TEST: The light comes on White when the test selector is set to asystem test position. It is extinguished when the test selector is inNORM FLT position.

5. SFCC 1 AND 2 BITE DISPLAY LIGHTSThese lights come on White when a fault has been detected by thecontinuous monitoring of the SFCCs even if the failure does notrequire crew action (no FAULT indication on the overhead panel).More details of the failure are displayed on the face of the SFCC's.

6. TEST RESULT OK LIGHTThis light comes on White when the test is successful.

NOTE: These lights (OK, FAULT, BITE, DISPLAY) will illuminateproviding that the ANN LTS switch is in READ position during the test.


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Maintenance Panel - Flight Controls Test Panel - Controls and Indications

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LEFT BLANK

ATA 27 - END

COURSE CODE - M540000

Airbus 27 A300 A310 Flight Controls

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