Airbus 27 A300 A310 Flight Controls

ATA 27 Flight Controls

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FLIGHT CONTROL SURFACES 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 secondary flight controls are the:

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 the trimmable horizontal stabilizer.

•

PITCH TRIM CONTROL achieved by the trimmable horizontal stabilizer hinged on the aircraft structure.

•

YAW CONTROL achieved by one rudder.

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•

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 each wing

•

GROUND SPOILERS No. 1 through No. 7 on the upper surface of each 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 flight surfaces located in the cockpit.

G. Speed Brake Control Panel H. ECAM Display Control Panel

A. Servo Control Panel I.

Aileron and Rudder Trim Switches

J.

Left ECAM Display Unit

B. Slats and Flaps Position Indicator C. Pitch Trim and Yaw Damper Switch Panel K. Master Warning and Caution Lights L/H D. Flight Control Maintenance Test Panel L.

Master Warning and Caution Lights R/H

E. Right ECAM Display Unit F. Pitch Trim Wheel

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Flight Compartment Controls and Indicating

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

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

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Flight Controls Hydraulic Power Supply

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FLIGHT CONTROLS - GREEN HYDRAULIC 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

2.

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

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

NORMAL (P/B SWITCH PRESSED-IN) Hydraulic power is supplied to the corresponding users as soon as pressure is available in the corresponding hydraulic system.

•

OFF (P/B SWITCH RELEASED-OUT) The OFF light comes on White and the hydraulic power supply to the corresponding users is shut off. The associated JAM warning is inhibited and LO PR Amber illumination confirms the OFF selection.

•

JAM When a P/B switch is pressed-in, the associated JAM light comes on Amber when a jamming is detected in the related hydraulic control valves of rudder, elevator, ailerons or trimmable horizontal stabilizer. Illumination of a JAM light is accompanied by ECAM activation. The jammed control is identified on the Warning Display (left CRT).

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B, G, Y LO PR LIGHTS

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

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To prevent inadvertent complete deactivation of servo controls, only two systems can be deactivated at a time by selection of SERVO CTL P/B switches to OFF. When the third P/B switch is selected to OFF all three systems are reactivated regardless of P/B switch setting.

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SERVO CTL (Servo Control) Panel

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ROLL CONTROL SECTION

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

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 and 26 V AC). If the normal buses have been cut off before landing, power is supplied again to three spoiler groups by pressing the LAND RECOVERY P/B switch on the overhead panel.

Each computer is composed of two control units and two monitoring units. Each unit controls or monitors one group of surfaces. Each group is made 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, the corresponding control unit is in one computer and the monitoring unit is in the other one. For the roll spoilers the control laws are such that they are not usually used unless the control wheel is moved enough. An autopilot servo actuator is mounted adjacent to the RH wing rear cable quadrant. It drives the complete control via a detent lever which can be overridden by the pilots.

From the two interconnected control wheels, the roll inputs are transmitted to the ailerons by dual cables providing fail safe operation. In each wing the inputs are transmitted to a differential unit receiving additional inputs from: • artificial feel unit • aileron droop unit • trim screw jack In case of jamming in one control run, the interconnected spring strut can be compressed to permit operation of the other control run to the other wing. The pilot effort required on the wheel is between 34 lbs. and 90 lbs. Spoiler control is still available but downgraded. Each servo control linkage on the aileron includes a spring rod to protect it against a runaway if an input lever on one jack remains in the open position.

INTERFACE WITH AUTOPILOT SYSTEM An autopilot actuator is mounted adjacent to the right wing rear cable quadrant; it drives the complete control via a detent lever which can be overridden by the pilots. Dynamometric rods are installed upstream of the cable tension regulators, they provide control signals to the control wheel steering system.

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

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Roll Control - Mechanical Aileron Control

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AILERON SYSTEM - COMPONENTS DESCRIPTION 1.

CABLE TENSION REGULATORS Two tension regulators maintain a constant tension on the cables of 28.13 ± 5.30 lbf. They are identical apart from the input lever position. They incorporate provision for installation of a special tool used for installing the regulator on the aircraft.

2.

SERVO CONTROL ACTUATING SPRING ROD The three ASA servo control actuating spring rods prevent runaway of the control system if an input lever jams on its servo control body.

3.

CONTROL WHEEL INTERCONNECTING SPRING ROD The two control wheels are interconnected by a spring rod in order to allow one of the crew members to control half the surfaces in the event of any single item jamming in the mechanical control system.

4.

RODS Push-pull rods are adjustable or nonadjustable length, fitted with replaceable ends.

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5.

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

6.

DYNAMOMETRIC RODS The Flight Control Computer uses signals from the dynamometric rods to detect the Captain's and First Officer's loads on the control wheels. 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 the Flight Control linkages.

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Aileron System - Components 4

1 5

1

3

2

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All Speed Aileron (ASA) - Mechanical Control Each all speed aileron (ASA) is operated by three mechanically controlled servo controls. The two interconnected control wheels drive two symmetrical control systems composed of levers, rods, cables and tension regulators routed along each side of the fuselage up to the input levers of the servo controls. A differential and droop unit is installed in the control linkage upstream of the servo controls. The unit receives two inputs. One is from the control wheels (pilots input), the other is a droop signal from the slat control system which droops the all speed ailerons 9.2° when the slats are extended in order to optimize aerodynamic efficiency of the wing.

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When the droop signal is applied, all speed aileron deflection is not simply modified by 9.2° throughout the travel range. Instead, response of the all speed ailerons to control wheel motion is modified so that the maximum up and down deflections remain close to those without droop input. The droop signal also drives a differential mechanism between the trim screwjack and the artificial feel unit. The mechanism pivots the artificial 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 the main gear W/W (center fuselage) drives two trim screwjacks via sprockets, chains and cables. The actuator is controlled from panel 408VU located at the rear part of the center pedestal. Two switches on this panel allow the crew to select constant speed displacement in the appropriate direction. Trim position is indicated on scales at the top of the control columns when the wheels are released.

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In each wing root, displacement of the trim screwjack drives the all speed aileron servo control input linkage through the artificial feel unit, whose spring rod remains at neutral. When the ailerons are drooped, the droop signal drives a differential mechanism between the trim screwjack and the artificial feel unit. The mechanism pivots the artificial feel unit, thus allowing the spring rod to remain at neutral and the unit is held in this position by the trim screwjack.

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Aileron Trim System

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Aileron Trim Components A. AILERON TRIM ACTUATOR

B. ELECTRIC MOTOR CONTROL

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

Rotary stops limit output shaft rotation within the range allowed by the screws. When the electric motor is energized, it is protected by a torque limiter 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 transmitter settings).

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The electric motor windings of aileron trim actuator 9CG are energized through contacts of two adjacent three position switches (5CG) on control panel 408VU. The switches are spring loaded to the center position and must both be moved simultaneously in the same direction for the windings to be energized. The switch tabs are not mechanically connected, to prevent run-away in the event of mechanical jamming of one tab. The windings are shorted when the two 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 immediately downstream of the all speed aileron servo control actuating spring rods. The units each include a spring rod and are held in position by the trim screwjacks.

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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 (L WING or R WING) to energize the system. This action selects a constant speed displacement in the corresponding direction. Full travel of about 7° of aileron in each direction is achieved at a speed of 0.4° per second.

2.

ON (P/B switch pressed-in): Corresponding control system is activated. Each time a system is activated, or corresponding hydraulic system on, or the aircraft electrical network is energized, a 2 second safety BITE test is triggered 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 control system is deactivated. If hydraulic pressure is available, the actuators are automatically held in the retracted position. The monitoring circuits are reset by ECAM activation. This action is accompanied by ECAM activation.

•

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

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 opposite a pointer painted on the control wheel. With the control wheels released, 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 symmetrical upper wing surfaces.

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Aileron Trim/spoiler and Speed Brake Switches

A

B

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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 artificial feel unit connected to the trim screwjack. From this point up to the servo controls, the commands are transmitted by dual rigid linkage, receiving additional inputs from a rudder travel limiter, yaw damper and autopilot servoactuators. The artificial feel is provided by a spring loaded rod. The trim actuator is electrically signaled. It is driven by an electrical motor. During cruise, the operational limits for rudder trim are ±1.5°.

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

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 are delivered by two independent RUDDER TRAVEL channels, each one included in a digital computer (Feel and Limitation Computer) receiving inputs from the DADCs (Digital Air Data Computers) and the SFCCs (Slats Flaps Control Computers). Each computer controls an electrical motor driving a common electromechanical actuator coupled to variable stop lever. Only one channel is normally active. The other is in standby. A spring loaded rod positions the variable stop lever in the low speed position in case of dual failure.

INTERFACE WITH MAIN WHEEL BRAKING SYSTEM Levers are attached to each pedal, to provide braking inputs when the pedals rotate about their axis. INTERFACE WITH NOSE WHEEL STEERING The nose wheel steering control is connected to the rudder control through a hydraulic steering control coupler (engaged when the landing gear is extended) and a spring rod, the threshold of which is lower than the threshold of the rudder artificial feel and trim unit spring rod. The spring rod prevents the nose wheel steering control from transmitting inputs to the rudder control.

An. autopilot servo actuator is mounted adjacent to the artificial feel unit upstream of the variable stop lever. It drives the complete control via a detent lever which: car. be overridden by the pilot. Yaw damper commands are transmitted via a differential unit canceling a feedback to the pedals. A spring loaded rod on each servo control input avoids a runaway of the rudder in case of jamming of one input lever in the open position. Levers are attached to each pedal, to provide brake inputs when the pedals rotate around their pivots.

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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 the reduction gear. It is energized when rudder travel is selected.

•

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

•

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

•

•

The actuator includes a position transducer which delivers rudder trim position signals to associated electrical circuits. The transducer is a special RVDT, of the same type as those installed in the two transducer units used for electrical roll control. The electrical characteristics of the RVDT are monitored by associated computer circuits.

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

•

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.

2.

RUD TRIM ROTARY SELECTOR Rudder trim control is electrically powered. The rotary selector is springloaded to the neutral (center) position. The direction of rudder trim travel depends on the direction of rotary selector (NOSE L or NOSE R). Full authority of rudder trim is about 21° in each direction. RESET PUSHBUTTON SWITCH It allows initiation of an automatic sequence controlled by the EFCUs to position the rudder trim at 0° ±0.2°. •

3.

•

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

•

FAULT The light comes on Amber if a failure of the reset function is detected or if the actuator position transducer fails.

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

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

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Rudder System Yaw Control - Rudder Mechanical/Hydraulic

3 1

2

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Rudder System - Rudder Artificial Feel A spring assembly located in the artificial feel and trim unit restores a resistance to pedal depression which is proportional to rudder movement. A variable stop lever installed downstream of the servo controls on the control linkage serves to reduce rudder deflection with respect to pedal movement as the airspeed increases.

Spring function is: • • • •

RUDDER ARTIFICIAL FEEL An artificial feel and trim unit is installed adjacent to the rear cable quadrant. It consists of a trim screwjack and a fail-safe constant resisting load spring rod, held in neutral position by the trim screwjack.

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•

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To maintain the downstream linkage and the input lever of the servo controls at neutral in the event of disconnection of the control linkage upstream 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

2.

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

SERVO CONTROLS JAMMING DETECTION

There is one jamming detection circuit for each hydraulic system. If jamming occurs the electronic circuitry inside the jamming detection control unit receives 28V directly from the jamming detection microswitch if a servo control is involved and from an intermediate logic if a THS actuator hydraulic motor is involved. Jamming detection is associated with the mechanically driven control valves of the left and right all speed aileron, and left and right elevator and rudder servo controls. It is also associated with the THS actuator hydraulic motor control valves (for the Green and Yellow systems only). NOTE: When a hydraulic system is selected OFF, the + 28V sent to the corresponding 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 PANEL 1. The P/B switches control channels 1 and 2 of the Feel and Limitation Computers (FLC) for rudder travel limiting. •

2.

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

•

OFF/R (P/B switch released-out): The OFF/R light comes on White and the system involved is disengaged. 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 on Amber if a failure is detected in the respective system. Illumination of the Amber FAULT light is accompanied by ECAM activation. Both FAULT lights remain illuminated when the switches are released-out and the OFF/R lights are illuminated White. This constitutes a rudder disagree warning (The variable stop lever is not in low speed position with flaps extended 20° or more). Illumination of both FAULT lights is accompanied by ECAM activation.

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YAW DAMPER LEVERS •

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

•

OFF: The respective yaw damper is disengaged. When one YAW DAMPER lever trips to OFF, the associated yaw damper system disengages and the ECAM is activated. When both YAW DAMPER levers trip to OFF the yaw damper function is lost and the SCAM is activated.

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Rudder System Travel and Yaw Damper Systems - Control Switches

B

A

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Rudder System - Rudder Travel Limiting • • •

The rudder travel limiting system modifies control inputs to the servo controls to vary rudder travel in relation to airspeed (Vc). Limitation is such that the maximum deflection which can be achieved by the rudder remains lower than the deflection which would induce limit loads on the structure, throughout the flight envelope. 2.

TRANSDUCER UNIT The actuator is servo controlled and is monitored through a transducer unit driven by variable stop lever movement. The transducer unit, comprising two inductive transducers, is identical to the one used in the spoiler control system.

3.

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

The system is composed of: •

• •

1.

A variable stop unit consisting of an articulated lever operated by an electromechanical actuator and a transducer unit detecting lever position. These items are all mounted on a frame assembly located downstream of the differential between the AP and yaw damper actuators. Two control and monitoring computers designated FLC (Feel and Limitation Computer). One RUD TRAVEL control panel, one PITCH FEEL & RUD TRAVEL maintenance panel and five electrical power supply circuit breakers.

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 it reaches the mechanical stop.

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

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Rudder System - Rudder Travel Limiting

2

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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: rudder travel limiting and pitch feel. The FLC is a digital computer comprising two 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 by comparison between the results of both channels, by means of analog comparators • power loop monitoring achieved by software means in each digital channel. If any indicator is on, the test of either RUDDER TRAVEL LIMITING system 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 302CY1 FLC2/FIN 302CY2

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ELEVATOR SYSTEM PITCH CONTROL SECTION

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Elevator System - Pitch Control Pitch control is achieved by two elevators hinged on the horizontal stabilizer, each actuated by three servo controls controlled by a dual mechanical linkage through dynamometric rods, cable runs, an artificial feel system linked to the cable run of the LH control column, and load limiting rods. In normal operation the two elevators are controlled together. In case of jamming in one control linkage during flight (take off excluded), pitch control is provided by THS (Trimmable Horizontal Stabilizer). If jamming occurs at take off, two uncoupling bellcranks enable the elevator on the other side to be controlled by one or both pilots.

downstream of the artificial feel system, a load limiting spring rod limits the efforts in the elevators control linkage. A spring loaded rod on each servo control input avoids a runaway of the elevator in case of jamming of one input lever in the open position. An autopilot actuator is mounted adjacent to the LH elevator. It drives the control via a detent lever which can 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 jack driven by two independent hydraulic motors supplied respectively by Green and Yellow systems and coupled by a differential gear through pressure-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 (electrical mode). • automatically by AP trim, mach trim or alpha (angle of attack) trim function.

A pitch uncoupling unit (locking rod plus solenoid) prevents accidental asymmetrical deflection of the elevators during flight and allows uncoupling of the RH and LH control systems during take off (locked at speeds lower than 30 kt or higher than 195 kt). Artificial feel is provided by 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 is actuated by either of two electrohydraulic actuators. Each actuator is controlled by an independent PITCH FEEL channel, each one included in a FLC (Feel and Limitation Computer).

Electrical and automatic trim signals are processed in two FAC (Flight Augmentation Computers) and control two electrical motors. Trim speed and trim authority depend on trim mode and aircraft configuration. The motors drive the control linkage to the hydraulic valves which control the hydraulic motors. The manual trim wheel run is connected to the same linkage. 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)

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

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Elevator System - Pitch Control - Diagram

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Elevator System - Elevator Mechanical Control System - General and Components A. GENERAL

4.

Rods: Identical to those used in the aileron control system

Each elevator is operated by three mechanically controlled servo controls. The inputs from the control columns are transmitted to the elevators by dual control systems. Each system is routed along one side of the fuselage. The left and right systems are interconnected at two points by detent bellcranks, one beneath the flight compartment floor, the other between the two elevators.

5.

Cables: Identical to those used in the aileron control system

6.

Control column stops: Control column travel is limited in both directions by non adjustable stops. Elevator operational stops: Maximum input to the servo controls is limited 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 normal operation. Elevator structural stops, when the servo controls are not installed: The elevators rest on structural down stops, designed for that purpose, which are not able to withstand any load other than the weight of the elevators. Adjustable levers: The length of a lever close to each elevator is adjustable in order to maintain maximum travel of the elevators within the design limits.

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.

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

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

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Elevator System - Elevator Mechanical Control

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ARTIFICIAL FEEL (ELEVATOR) The pitch artificial feel system creates load feel at the control column which is variable with flight conditions, in order to reduce the variation of force per g throughout the whole flight envelope. At high angle of attack, the system causes an increase in the load feel at the control column resulting in aircraft return to permissible angle of attack configuration. 1.

2.

Pitch Artificial Feel Actuator Each actuator includes: • A biased servovalve which modulates pressure in the actuator large chamber, the small chamber being permanently supplied with high pressure. In the event of an electrical failure, servovalve current is nulled and its control valve is displaced so that the 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 is therefore redundant with respect to the servovalve bias. • A position pickoff potentiometer.

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 of double hydraulic failure.

3.

Pitch Upcoupling Unit Provides connection of LH and RH elevators from 0-30 knots airspeed during takeoff roll. Above 30-195 knots, the LH and RH elevators are disconnected by ADC 1/2 to allow either pilot to control the elevator (pitch) function. In case of a jam in the elevator control system on the captain’s or first officer’s control panels Above 192 knots, both elevators will reconnect for full control of the LH and RH elevator system runs. During landing conditions, this process is repeated in the same airspeed conditions.

Pitch Artificial Feel Unit - Operation The actuators act on the gain variation mechanism by means of levers. Gain is imposed by the actuator having extended the furthest. In the event of jamming of the mechanism, a microswitch transmits a pitch disagree warning signal. The artificial feel unit includes a "fail safe" part to avoid loss of the force threshold and feel load at the same time. 4.

Aft Detent Bellcrank In the event of an elevator jam on the LH or RH elevators, the aft detent bellcrack will release on the jammed side to prevent lockout of the elevator system.

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Elevator System - Pitch Artificial Feel - Components/Location

3 2 4

1 Artificial Feel Unit

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Elevator System - Pitch Uncoupling System - General and Components The two elevator control channels can be uncoupled during the takeoff phase in the event of jamming at any point on the control systems, by means of two detent bellcranks; one installed between the control columns, the other between the two elevators. A pitch uncoupling unit, comprising a solenoid and rod, prevents any inadvertent uncoupling of the two elevators after the takeoff phase in order to prevent asymmetrical loads being applied to the structural attachments of the trimmable horizontal stabilizer. The uncoupling unit solenoid is energized if airspeed Vc is higher than 30 kts and lower than 195 kts.

COMPONENT DESCRIPTION 1.

Solenoid The 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 value can be obtained from ADCs) has been introduced to prevent permanent energization of the solenoid and power contactor coil when the aircraft electrical network is energized on the ground.

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Elevator System - Pitch Uncoupling System - Schematic

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Trimmable Horizontal Stabilizer System - Pitch Controls Both pitch trim wheels provide mechanical control of the Trimmable Horizontal Stabilizer (THS). When a pitch trim control wheel is used to override the electrical command, it disengages the electric actuators and the PITCH TRIM levers trip to OFF. The trim range is from 14° nose up to 3° nose down. Trim position is indicated in degrees on a scale adjacent to each trim wheel which is painted Green over the normal take off 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 electric actuators which control the hydraulic motors for horizontal stabilizer adjustment providing that at least one PITCH TRIM system is engaged and AP is OFF or in CWS mode. The rocking levers are spring loaded to neutral position. If both rocking levers are operated simultaneously, but in opposite position, trimming action stops. If trimming by means of the rocking 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 between 200 and 240 kts.

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Trimmable Horizontal Stabilizer System - Pitch Controls

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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 no special reference mark painted on the elevators, but on each side of the APU tailcone, there is: •

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

•

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

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1.

ELEV AND STAB POSITION INDICATION A White scale covering the full travel range is provided for elevator and trimmable horizontal stabilizer position. An index indicating the actual position of the surfaces moves along each scale. In addition, each available hydraulic system on the THS is indicated by a Green symbol (G,Y). In case of servo control low pressure detection, the corresponding symbol becomes Amber.

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Flight Controls System - RH ECAM Page - System Display

Elevator and Horizontal Stabilizer Position Indication

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Pitch Feel and Trim System - Control Switches and Levers 1.

PITCH FEEL SYS 1 AND 2 PUSHBUTTON SWITCHES

2.

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

•

•

ON (P/B switch pressed-in): The corresponding system is engaged. Both systems may be engaged simultaneously but only one is effectively operating. If one system fails, it is automatically deactivated and the other one continues to operate. OFF/R (P/B switch released-out): The OFF/R light comes on White and the system involved is disengaged. The monitoring circuits are reset by this action. This indication is accompanied by SCAM activation.

PITCH TRIM 1 AND 2 LEVERS •

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

•

OFF: The respective pitch trim is disengaged. -

When one PITCH TRIM lever trips to OFF, all electrical control modes of the THS are lost and the ECAM is activated.

-

When both PITCH TRIM levers trip to OFF, all electrical control modes of the THS are lost and the SCAM is activated.

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

FAULT: When a P/B switch is pressed-in, the associated FAULT light comes on Amber if a failure is detected in the corresponding system. Illumination of the Amber FAULT light is accompanied by ECAM activation. Both FAULT lights remaining illuminated when the P/B switches are released-out and the OFF/R lights are illuminated White, constitutes a pitch disagree warning (The artificial feel unit operates in high speed configuration when flaps are extended 20° or more ) . Illumination of both FAULT lights is accompanied by ECAM activation.

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

B

A

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TRIMMABLE HORIZONTAL STABILIZER (THS) SECTION

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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 on the THS. Their control systems are installed so that the elevators are in line with the THS when the control columns are released. The THS is driven by an actuator including a fail-safe ball screwjack, the structural attachments of which are also fail-safe. Normal control of the actuator is electrical, via the automatic pitch trim system.

INTERFACE WITH AUTOMATIC PITCH TRIM SYSTEM Electrical control is achieved by means of two electric pitch trim actuators, installed on the THS actuator. They drive the actuator mechanical input. The two pitch trim actuators are controlled by two flight augmentation computers (FAC) which deliver manual electric trim, automatic trim, Mach trim and alpha trim signals. Manual electric trim signals are provided by rocker switches mounted on the Captain's and First Officer's control wheel horns. Electric limit switches detect THS end of travel in aircraft nose up direction. The signals are used in the automatic trim system to avoid automatic disconnection of this system during automatic landings.

Stand-by controls are mechanical. The pilots can override electrical control by the mechanical control system by applying sufficient force to the control wheels. A torque limiter is mounted in each electric pitch trim actuator for that purpose. The torque limiters remain automatically released after their operation.

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Trimmable Horizontal Stabilizer System - Diagram

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Trimmable Horizontal Stabilizer System - Hydraulic Actuation - Components/Location The trimmable horizontal stabilizer is driven by an actuator which includes two hydraulic motors, each powered by a different hydraulic system.

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Trimmable Horizontal Stabilizer System - Hydraulic Components/Location

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Trimmable Horizontal Stabilizer System - Actuator - General The THS actuator consists of a fail-safe ball screwjack actuated by two hydraulic motors coupled by a differential gear. NORMAL OPERATION Pressure-off brakes (9) are released. The ball screw is held by the no-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

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the mechanical input, controls rotation of the ball screw through two identical control loops, including input and feedback gear trains, feedback differentials (4), control valves (5), hydraulic motors (8) and actuate a power gear train through power differential (10). The control stroke is limited by the actuator input shaft stop (2). The structural components (ball screw and nut assembly, attachments to THS and fuselage) and the power gear train are duplicated, the secondary load path being normally unloaded.

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Trimmable Horizontal Stabilizer System - Actuator - Schematic

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WARNING LOGIC PITCH CONTROL SYSTEM

Depicted below are the various warnings displayed in the Flight Compartment in the event the fault shown occurs in the Pitch Control System of the aircraft. Also graphically shown are the Flight Phases at which the warnings will or will not be displayed.

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Warning Logic - Pitch Control System

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Alpha Probes System - Stall Warning - General •

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

•

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

•

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

•

Stall warnings are activated when angle of attack exceeds a predetermined value.

•

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

•

slat retraction is inhibited.

•

turn coordination of yaw damper is inhibited.

Stall warning is inhibited on the ground except during ALPHA PROBES test.

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Alpha Probes System - Stall Warning - Block Diagram

Captain

Captain

First Officer First Officer

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Flight Control System - Stick Shaker - General One stick shaker is installed on each control column, and is controlled by the Flight Warning Computers 1/2.

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Flight Control System - Stick Shaker

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LEADING EDGE LIFT DEVICES SECTION

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Wing Leading Edge - Slat System - General SLAT SYSTEM

KRUEGER FLAP AND NOTCH FLAP

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

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The Krueger flap and notch flap are provided to complete the wing leading edge profile when the slats are extended. The Krueger flap and notch flap are operated by individual hydraulic actuators. Both are controlled by the SFCC and move to the extend and retract position when the SFCC commands slat extension or retraction. Slats and spoilers are numbered from inboard to outboard, each side separately.

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A300 Wing Leading Edge - Slat System Components - Location

A300 SHOWN/A310 DO NOT HAVE NOTCH (OR SLOT) FLAPS

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Wing Leading Edge - Slat System - Hydraulic Operation - Diagram The Power Control Unit (PCU) hydraulic supply is provided by the aircraft hydraulic 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 single system failure the system will still operate but at half speed.

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Wing Leading Edge - Slat Hydraulic System - Schematic

* NOTE: A300 ONLY/A310 DO NOT HAVE NOTCH OR SLOT FLAPS

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Wing Leading Edge - Slat Control System - General The power is supplied to the ball screwjacks by a torque shaft driven by a power control unit and protected by a system torque limiter for each wing. Each ball screwjack also has its own torque limiter. All these torque limiters include a latched lockout indicator and in case of overload of a jack in the torque shaft system, they will freeze the system until a reverse selection is attempted. As soon as an order is given, the corresponding computer of each motor sends signals to deliver the pressure to the motor, releases the involved

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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 by its brake and the operating speed of the slats is reduced by half due to the differential mechanism of the power unit gearbox. However, full torque is still available. Three slat positions can be selected (0°, 15°, 30°) by the five position control lever.

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Wing Leading Edge - Slat System Control and Indicating - Schematic

ADC 1 ADC 2

* NOTE: A300 ONLY/A310 DO NOT

*

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HAVE NOTCH OR SLOT FLAPS

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Wing Leading Edge - Slat - Hydraulic Power Drive System - General The slat drive system comprises the power control unit (PCU), a transverse torque shaft system and the screwjacks. In the PCU two independent hydraulic motors, one controlled by the Blue valve block and one controlled by the Green valve block, drive a summing gear. The output is passed through torque shafts to a tee-gearbox which rotates the motor drive direction by 90°. A pressure-off brake is provided between each motor and the summing gear to lock the transmission system when the slat system is static.

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

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

D FW

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Wing Slats and Flaps Systems - Indicating and Control - Description A single control lever located on the center pedestal permits slat and flap control. The lever has five gated positions. It is not possible to select an intermediate position (if the lever is held in between gates the system drives to the last demanded position and after 10 sec. all the slat and flap FAULT warnings illuminate). The slats and flaps are electrically signaled by two identical digital computers (Slats Flaps Control Computers). Each one is composed of one slat control channel and one flap control channel.

1.

SLAT/FLAP POSITION INDICATOR STRIPS Slat and flap positions are shown by White strips moving up and down associated scales. The corresponding VFE (speed limit) is placarded opposite each normal position (indicated by a round number).

2.

SLAT AND FLAP LIGHTS Come on Amber when the associated system is blocked.

SLATS (OR FLAPS) SYS 1 AND 2 FAULT LIGHTS

3.

Each light comes on Amber when the associated hydraulic motor is inoperable. 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 will move to the commanded position.

KRUEGER LIGHT Comes on Amber if either KRUEGER flap is not in correct position 10 sec. after a movement command. Illumination of KRUEGER light is accompanied by ECAM activation.

4.

SPD BRK LIGHT The light comes on Blue when the speed brake control lever is not in RET position.

5.

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

Both SLAT (or FLAP) FAULT lights and the associated Amber SLAT (or FLAP) lights on the Slat/Flap Position Indicator will come on simultaneously if a mechanical failure is detected. In this case, the system is locked by the pressure-off brakes and there is no possibility of recovery in flight. NOTE: If a SFCC is not installed, the two associated FAULT light (one SLAT FAULT light and one FLAP FAULT light) will come on. Illumination of these lights is associated with ECAM activation.

Before selection of any position, the slat/flap control lever must be pulled up. A block is provided for positions 2 and 4 to prevent the lever moving straight through. NOTE: All slat and flap FAULT lights will illuminate if the control lever remains between two gated positions (after 10 sec.). NOTE: All Slat and Flap FAULT lights will remain illuminated if Slat/Flap Control Computer (SFCC) number 1 and 2 are removed from Electronic Rack 90VU. The 28 Volt DC Interlock Relay in the system prevents accidental dispatch of the aircraft if both SFCC 1/2 are removed from the electronic rack or aircraft.

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Wing Slats and Flaps Systems - Indication and Controls B A

C

0

0

15

0

SLATS

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FLAPS

15

15

15

20

30

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A300/A310 Wing Leading Edge - Slat - Asymmetry Monitoring Slat position is indicated to the flight crew by the SLATS vertical bar display on the slat/flap position indicator. The display is driven by inputs from the instrumentation Position Pick-off Unit (PPU) which also provides independent slat position information for other systems.

ASYMMETRY AND POWER TRANSMISSION MONITORING ( SLATS) The Asymmetry PPUs enable the SFCCs to monitor the transmission for asymmetry and runaway conditions.

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

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

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

B

B

B

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Wing Leading Edge - Krueger and Notch Flap Actuation When the slats are extended, SFCC1 and SFCC2 send an extend discrete signal to a Krueger selector solenoid valve, which has an extend and retract solenoid. When the extend solenoid is energized, hydraulic pressure is passed to the Krueger and notch flap actuators and to the all speed aileron system. The Krueger and notch flaps extend. The solenoid remains energized.

KRUEGER AND NOTCH FLAP ACTUATION The Krueger selector solenoid valve is located in the hydraulic bay at FR47. It is a three-position, four-port solenoid-operated shuttle valve. The shuttle valve is springloaded to center. It moves to the center position when both solenoids are deenergized. In this position, the hydraulic pressure input in A is shut off and ports B, C and D are interconnected.

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

INTERFACE WITH THE AILERON SYSTEM When the Krueger and notch flaps are supplied with pressure from the Krueger selector solenoid valve, so also is the droop actuator in the all speed aileron system. When the Krueger and notch flaps are extended, the ailerons droop 9.2°. On retraction, the ailerons return to their normal positions. NOTE: The A310 Slat System does not have Notch or Slot Flap devices.

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

A300 ONLY

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Wing Leading Edge - Krueger and Notch Flap - Components KRUEGER FLAP ACTUATOR The Krueger flap actuator is a double-acting actuator with mechanical locking of the piston assembly in the extended position (flap retracted) and hydraulic locking in the retracted position (flap extended). The piston end is connected to the Krueger flap by a reverse link. The cylinder head is mounted to the structure by two mounting blocks which allow the actuator to pivot on the mounting during the operating cycle. The actuator 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 hydraulically locked in the retracted position. The piston rod is connected by an eye-end to the notch flap and the cylinder is attached to the structure by a shaft hinge to allow some pivoting during the retraction and extension cycle. SAFETY VALVE The safety valve is located in the pressure line from the Green hydraulic system to the solenoid selector valve. Its function is to prevent loss of hydraulic fluid from the Green system should there be a major rupture in the Krueger and notch flap actuating system. *NOTE: The A310 Slat System does not have Notch or Slot Flap devices.

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Wing Leading Edge - Krueger and Notch Flap - Components

*

* NOTE: A300 ONLY/A310 DO NOT HAVE NOTCH OR SLOT FLAPS

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Wing Leading Edge - Krueger and Notch Flap Control and Monitoring - General In addition, to obtain better aerodynamic characteristics, a KRUEGER flap and a NOTCH flap are provided on each wing and are located between the inner slat and the fuselage.

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

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

The KRUEGER jacks are mechanically locked in retracted position and hydraulically locked in extended position.

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Krueger/Notch Flap Control and Monitoring - Schematic

* MTT M540000 R3.3 01AUG01 For Training Purposes Only

*

NOTE: A300 ONLY/A310 DO NOT HAVE NOTCH OR SLOT FLAPS

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Wing Leading Edge - Slats/Krueger Flap Monitoring and Fault Warning System The SFCC provides continuous monitoring of the slat system and the Krueger and notch position. Fault warnings are generated for those faults requiring pilot action or flight crew awareness. The faults are stored in the SFCC including those which are purely maintenance data. The warnings are displayed on one or more of the following: • • • •

SELF TEST Self test facilities are provided to: • •

SLATS SYS 1 FAULT or SLATS SYS 2 FAULT annunciator (19CV) on the overhead panel BITE DISPLAY/SFCC1 or SFCC2 annunciators (52CV and 53CV) on the FLIGHT CONTROL section panel 471VU Left electronic centralized aircraft monitor (ECAM) Fault indicator on the SFCC front panel.

MTT M540000 R3.3 01AUG01 For Training Purposes Only

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 by using the BITE pushbutton switch on the SFCC front panel.

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Fault Indicators and Flight Controls Test Panel - Krueger System

MAINT PANEL 471VU A300/SOME A310

1/FIN 21CV 2/FIN 22CV A300 Indicator Shown

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Warning Logic Slat System

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

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A300/A310 Slats System - Warning Logic

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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 slotted fowler type and are guided by two tracks fitted with ball screwjacks. The power is supplied to the ball screwjacks by a torque shaft driven by a power control unit and protected by a system torque limiter for each wing. Each ball screwjack has its own torque limiter. All these torque limiters include a latched lock-out indicator and, in case of overload of a jack in the torque shaft system, they will freeze the system until a reverse selection is attempted. The Power Control Unit (PCU) consists of two independent hydraulic motors coupled to a differential mechanical system through pressure-off brakes that insure the system irreversibility. The two motors are supplied by different hydraulic circuits (Green and Yellow). As soon as an order is given, the corresponding computer of each motor sends signals to deliver the pressure to the motor, release the involved pressure-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.

MTT M540000 R3.3 01AUG01 For Training Purposes Only

In case of hydraulic leakage, the corresponding motor remains locked and the operating speed of the slats is reduced by half due to the differential mechanism of the power unit gear box. However, full torque is still available. Four flap positions can be selected (0°,15°,20°,40°) by moving the slat flap control lever from position 2 to 5. Furthermore, between the inboard and center flaps, there is an aileron droop signal unit which commands the aileron to droop 9.2° maximum with slats extension to 15°. A load relief system is provided to minimize the design loads on the flap support structure and the flap jacks. Load relief function can only engage when the slat flap control lever is in gate 5. Load relief is activated within the 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 3°UP ± 0.4 or > 2.3°DN ± 0.4). The Red T.O. CONFIG light comes on WLDP with associated CRC and ECAM activation.

MTT M540000 R3.3 01AUG01 For Training Purposes Only

Pressing T/O CONFIG TEST pushbutton also monitors the following systems: • 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.

2.

B,G,Y TEST PUSHBUTTON SWITCHES After selecting the related SERVO CTL P/B switch to OFF (on overhead panel), the TEST P/B switch for the respective circuit is magnetically latched when pressed-in and the TEST light comes on White.

Left pushbutton tests pitch feel and rudder travel system 1. Right pushbutton tests pitch feel and rudder travel system 2. When a TEST pushbutton is pressed and held, the associated system must disengage and its FAULT light comes on Amber.

For jamming detection test, the controls involved must be moved rapidly. Successful test is indicated by flashing of the JAM light in the related SERVO CTL P/B switch. If not successful, the fault isolation procedure must be done on the face of the jamming detection control box. After selecting the SERVO CTL P/B switch to normal, the TEST P/B switch is automatically released-out and the TEST light goes off.

MTT M540000 R3.3 01AUG01 For Training Purposes Only

TEST PUSHBUTTONS The pushbuttons control the test of PITCH FEEL and RUD TRAVEL electrical systems and warning systems continuity. The test is possible only if PITCH FEEL and RUD TRAVEL systems are engaged on the control panel (overhead panel).

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.

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OK LIGHTS These lights illuminate White as long as the TEST pushbutton is pressed and held, to indicate a successful test.

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Maintenance Panel - Flight Controls Test Panel Pushbuttons 2 3

1

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Maintenance Panel - Flight Controls Test Panel - Controls and Indications 1.

2.

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 when the 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 hydraulic power to be cut off to all flight controls to have the FAULT lights illuminated before test. • PITCH CTL UNCOUPLING: Tests periodically, on ground, the electrical 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 pressed in. GND SPLR SEL LANE FAULT LIGHT This light comes on White when a fault has been detected in TEST 1 or TEST 2 positions of the test selector.

MTT M540000 R3.3 01AUG01 For Training Purposes Only

3.

EFCU BITE DISPLAY LIGHT This light comes on White when a fault has been detected by the continuous monitoring of each EFCU. More details of the failure are displayed of the face of the EFCUs.

4.

PTT PUSHBUTTON SWITCH This P/B switch activates the test of the system selected by the test selector. 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 a system test position. It is extinguished when the test selector is in NORM FLT position.

5.

SFCC 1 AND 2 BITE DISPLAY LIGHTS These lights come on White when a fault has been detected by the continuous monitoring of the SFCCs even if the failure does not require 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 LIGHT This light comes on White when the test is successful.

NOTE: These lights (OK, FAULT, BITE, DISPLAY) will illuminate providing 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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1

2

3

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