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Flight Augmentation (FAC) (1) General (a) Functions The FAC performs the functions given below:  ? yaw damper ? rudder trim (manual and automatic) ? rudder travel limiting  ? monitoring of the flight envelope and computations of maneuvering speed ? achievement of yaw autopilot order using power loops of yaw damper and rudder trim. In addition the FAC 1 performs the BITE function of the AFS.  (b) Operating principles The FAC is a dual-dual type system for yaw damper, rudder trim and rudder travel functions. FACs 1 and 2 can be engaged at the same time through FAC 1 and FAC 2 pushbutton switches on the overhead panel. Only one system is active at a time : FAC 1 has priority, FAC 2 being in standby and synchronized on FAC 1 orders. An automatic changeover occurs on FAC 2 in case of disengagement or failure of FAC 1. Partial changeover function per function (yaw damper, rudder trim, RTL) is possible. When the aircraft electrical network is energized, the functions that follow will operate independently of the FAC pushbutton switches: ?monitoring of the flight envelope ? computation of maneuvering speed. The FMGCs and the PFDs receive these information signals as follow: ? FMGC 1 and Capt PFD normally use data from FAC 1 ? FMGC 2 and F/O PFD normally use data from FAC 2 In the event of failure, the FMGCs and the PFDs use the data from the active FAC.  (2) Yaw damper The yaw damper function provides: ? manual yaw stabilization. The ELACs compute the corresponding data and transmit them to the rudder surface via the servo loop of the yaw damper (FAC). ? alternate law for Dutch roll damping when the ELAC no longer computes normal yaw stabilization. ? Dutch roll damping (including turn coordination) when the autopilot is engaged in cruise only. ? engine failure recovery when the autopilot is engaged (the ELACs provide this function in manual flight).  (3) Rudder trim The rudder trim function provides: ? manual control via a rudder trim control switch located on the center pedestal. In addition the ELACs compute a command signal for rudder deflection (normal yaw damping law including recovery of engine failure) performed by the trim sub-system in manual flight. Reset of the rudder trim position is possible using a pushbutton switch located on the center pedestal. ? automatic control when the autopilot is engaged which provides the accomplishment of yaw autopilot command and the recovery of engine failure.

Each MLG has a MLG leg which includes an oleo-pneumatic shock absorber. The L/G retracts into the available space of its MLG bays. A two-piece side stay assembly holds the MLG in the extended position. A lock stay keeps the side stay assembly stable in the lock down postion.  The NLG includes on oleo-pneumatic shock absorber and retracts forward into a bay in the fuselage. A two-piece drag strut assembly with a lock stay locks the leg in the extended position. The leg includes a hydraulically-operated steering mechanism.  The hydraulic circuit has actuators that extend and retract the gears and open and close the doors. The green hydraulic system supplies the hydraulic power to operate the actuators.  The system has two electro-hydraulic valves that control the operation of the actuators. One electro-hydraulic valve controls the actuators for the L/G. The other controls the actuators for the doors.  The LGCIU makes the selections of the valve assembly to retract or extend the L/G and move the doors in a given sequence.  The normal braking system is used to decrease the speed of the aircraft when it moves on the ground. The Brake and Steering Control Unit (BSCU) controls the operation of normal braking. Each brake has two hydraulically operated pistons. When braking is necessary, hydraulic pressure is supplied from the Green main hydraulic system . The system has two modes of operation, manual and automatic, and gives automatic anti-skid protection in each mode.  The input signals from the brake pedals are proportional to the amount of pedal travel and supply braking independently to each MLG. For automatic braking three Pushbutton Switches (P/BSW) each set an automatic braking program (LO, MED or MAX) in the BSCU, which gives a different deceleration rate. Use of the pedals that supplies an input signal more than a specified value, cancels an automatic braking program.  The brake temperature system includes a temperature sensor at each brake which measures the temperature and sends the data to a brake temperature monitoring unit. The brake temperatures are shown automatically on the wheel page of the ECAM System Display (SD)

The ice-protection system of the wing leading-edge slats 3, 4, and 5 uses hot air from the engine bleed air system . The hot air causes the slats leading edge to become warm which prevents ice. Each engine supplies its related wing. The anti-ice valve 9DL(10DL) isolates its related wing anti-ice system from the engine bleed air system. The anti-ice valves are electro-pneumatically operated, and, for safety, spring-loaded to the closed position. It is possible to supply the two wings from only one engine bleed-air system. To do this, the crossbleed valve opens to connect the two pneumatic systems.  Lagged ducts connect the anti-ice valve to a telescopic duct at the slat 3. Each of the slats 3, 4 and 5 have a piccolo duct that supplies the hot air (bled from the engine) to the related slat leading edge. The slat 3 piccolo duct is connected to the telescopic duct. The slat 3, 4 and 5 piccolo ducts are connected together by flexible ducts. The bleed air in the slats is then released overboard through the holes in the bottom surface of the slat.  The operation of the anti-ice valve is controlled by the WING pushbutton switch (P/BSW) on the overhead panel 25VU  The engine air-intake ice protection system uses hot air from the engine HP compressor intermediate stage. The hot air goes through a solenoid operated, butterfly, anti-ice valve, to make the engine intake-lip hot. Each engine has its own protection system and operates independently from the aircraft pneumatic system. If, with the engines running, there is an electrical power failure, the anti-ice valve will automatically open. If there is an anti-ice valve failure (with the aircraft on the ground) it can be manually locked in the open or the closed position.  A restrictor orifice is fitted downstream of each anti-icing valve to control the flow of air bled from the engine. The restrictor also reduces the leakage of air if a supply duct is damaged.  he pitot probes have electrical heat elements in the main body and in the inner surface of the cavity in the mast. When the aircraft is on the ground, the Control and Monitoring Unit (CMU) automatically decreases the temperature level.

Cabin Air Distribution and Recirculation The cabin air distribution system supplies conditioned air to the cabin and to the cockpit. The Cockpit and cabin main supply ducts come from the mixer unit, which is the central element of the distribution system. The air flows out of the cabin through the lower sidewall-panels (dado panels) to the underfloor area. A part of the exhaust air of the cabin and the underfloor area is mixed with fresh air from the air conditioning packs in the mixer unit. This recirculated air is drawn by recirculation fans through recirculation filters to the mixer unit. The passenger cabin is divided into the FWD and AFT distribution zones which are supplied with temperature controlled and conditioned air. The conditioned air provided by the air conditioning packs 10HM (11HM) is supplied through a mixing unit installed under the cabin floor (in front of frame 36). Cabin air which has entered the underfloor area, is drawn through recirculation filters 4012HM (4013HM) by recirculation fans 14HG (15HG). The recirculation fans blow the air through check valves 4020HM (4021HM) to the mixing unit. The mixing unit mixes the conditioned air from the conditioning packs with the cabin recirculation air.  From the mixing unit, the air is supplied to the cabin zones through ducts with different diameters. Large diameter supply ducts are installed under the cabin floor along the left-hand and right-hand side. Smaller diameter riser ducts with integrated sound absorbers are connected to the main supply ducts. The riser ducts supply the air to the cabin zones through cabin air outlets installed below and above the hatracks.  Additional riser ducts are installed at the rear of the FWD passenger doors, and also forward of the AFT passenger doors. The cabin and door outlets are positioned to prevent draughts at seat-head level, and divide the air equally through the cabin zones. Most of the distribution ducts are made of resin and glassfiber laminate with metal sleeves bonded to each end for duct interconnection. Flexible bellows, which are made of silicone laminate and glassfiber, connect the ducts to each other. Clamps secure the flexible bellows. Insulation shells which are made of polyethylene foam or glass wool (covered with a Hypolon material) are installed around the ducts.

The equipment and furnishings which are installed in the aircraft give comfort to passengers and the crew. The equipment in the cargo compartments is installed for handling and safety of the cargo. The emergency equipment is installed in the aircraft for the safety of the passengers and the crew. Rack Lower Section This section of the rack is divided into several compartments. Each compartment houses electrical generation equipment such as contactors, TRs, etc. associated with the corresponding systems:  ? system 1 AC compartment associated with engine 1 generator,  ? system 2 AC compartment associated with engine 2 generator,  ? APU system and ground power unit AC compartment associated with the APU generators and the ground power unit,  ? system 1 DC compartment,  ? system 2 DC compartment.  Rack Center Section This section of the rack houses the primary circuit breakers associated with the electrical generation systems installed in the lower section of the rack and some primary circuit breakers. These circuit breakers are grouped per system, their functional designation is given and they are geographically located by means of placards.  Rack Upper Section This section of the rack houses, split over two panels, all the distribution circuit breakers protecting the electrical lines supplying the systems. These circuit breakers are grouped per system, their functional designation is given and they are geographically located by means of placards.  Access to the various components Access to the AC and DC electrical power centers is gained by removing the cover plates. Access to the inside of the circuit breaker panels is gained by opening the hinged panels.

The pneumatic system gets compressed air from the pneumatic air sources and supplies the compressed air through ducts to the user systems. The ducts are installed in the fuselage, the belly fairing and the wings. The control of the pneumatic system is usually automatic. Bleed-Air Monitoring Computers (BMCs) control the automatic function.  The BMCs are installed in the avionics compartment. There is one BMC for each engine. You can also control the system manually. The pushbutton and selector switches on the overhead panel 30VU in the cockpit, control the manual function.  You can monitor the operation of the pneumatic system on the ECAM System Display (SD). The pushbutton switches on the overhead panel 30VU have no indication when the system operates correctly.  The aircraft engines, the APU or a ground air source can supply compressed air to the pneumatic system. The distribution system supplies the compressed air from the different sources to the user systems.  The aircraft engines are the primary source of compressed air in flight. The air is bled from the 5th (IP) or 9th stage (CFMI engines) of the engine High Pressure (HP) compressor. The engine bleed air is temperature and pressure controlled.  The Auxiliary Power Unit (APU) is the primary source of compressed air on the ground. The air is bled from the APU load-compressor module. You can also use the APU to supply bleed air to the user systems during flight. The APU can supply bleed air up to an altitude of 20000 ft. (6096 m).  A ground air source is an alternative to the APU for the supply of compressed air on the ground. There is one High Pressure (HP) ground connector installed on the aircraft. You can use it to let the ground air source supply compressed air to the pneumatic system.  The operation and control of the pneumatic system is usually automatic. The BMCs monitor and control the automatic operation. You can also control the system manually from the cockpit. The pushbutton and selector switches on the overhead panel 30VU control the manual function.

Critical Design Configuration Control Limitations (CDCCL)  Some procedures identify a fuel system item that is in a category known as a Critical Design Configuration Control Limitation (CDCCL).  This category is applicable to items that are identified as the possible source of ignition of fuel in a fuel tank.  You must keep CDCCL items in a serviceable condition. It is possible that damage, wear or changes to a CDCCL item can cause a fuel tank explosion. CDCCL items are identified by a WARNING in the procedures where they occur. When a procedure identifies a CDCCL item, it is a mandatory condition that you do the instruction correctly and accurately as the procedure tells you.  Description The fuel system:  ? keeps the fuel in the main fuel tanks and the center (transfer) tank, which are open to atmosphere through the vent surge tanks  ? controls and supplies the fuel in the correct quantities to the fuel tanks during refuel operations  ? supplies the fuel to the engines for thrust and to cool  ? supplies the fuel to the Auxiliary Power Unit (APU)  ? controls the movement of the fuel out of the center (transfer) tank  ? supplies the fuel to cool, the excess fuel decreases the temperature of the engines and the Integrated Drive Generators oil  ? gives indications in the cockpit of the usual system operation  ? gives indications in the cockpit of the failures that could cause an unusual condition. The valve of the crossfeed system is usually closed, and in this configuration it divides the main fuel pump system into two parts (one part for each engine). When the crossfeed valve is open, the two fuel supplies are connected together. Thus the two engines can be supplied with fuel from one of the wing tanks or from the two wing tanks. The crossfeed system is operated manually by the X FEED pushbutton switch 4QE. The X FEED pushbutton switch is installed in the cockpit on the fuel control panel 40VU.  The Engine/Warning Display (EWD) and the System Display (SD) (which together make the ECAM) give crossfeed system information to the crew. If specified failures occur: ? a warning is given on the EWD ? the FUEL page shows on the SD. When necessary, the crew can make a selection on the ECAM control panel to look at the FUEL page.

 (3) Normal System Operation  (a) Push the MASTER POWER pushbutton switch on the SCU to supply power to the system. To set one of the VTR's as source, push the SOURCE SEL/SEL membrane switch until the SOURCE SEL/VTR AUX indicator identifies the VTR. (b) To supply power to the HMDU push the ALL ZONES membrane switch in the ZONE SELECTION section. The switch integrated light comes on. Push the POWER/ON membrane switch in the ZONE OPERATION section. Its switch integrated light comes on. All HMDU in the cabin extend.  (c) To supply video sound use the PA/VOL arrow up or arrow down membrane switches and select the number "5" on the PA indicator in the ZONE OPERATION section.  (d) Insert cassettes into the VTR's and push the related PLAY pushbutton switches. The HMDU show the selected video program in the colored mode. (e) The video program is heard through the PCUs with a connected headset. The table below shows the PCU-channel/video-program sound channel relationship for both VTR's.

The power plant is installed in the rear fuselage aft of the passenger compartment. It occupies the tailcone, the rearmost portion of the fuselage. The tailcone has been fitted with a fireproof compartment to house the APU. Peripheral systems are installed both in the APU compartment as well as to the front and rear of it. Doors permit access to all components of the power plant.  The central component of the power plant is the APU. It is installed in a fireproof compartment between FR80 and FR84 . The access doors 315AL and 316AR, which together are the lower compartment wall, permit installation, servicing and removal of the APU.   The APU is positioned in the tailcone structure by 7 suspension rods. These are attached to one bracket each on the roof of the APU compartment between FR80 and FR83. Three vibration isolators, installed between the suspension rods and the APU, reduce the transmission of vibration and shocks from the APU to the aircraft structure and vice versa. Brackets attach the vibration isolators to the APU.  For fire protection purposes, the APU is encapsulated within the A/C. This has been accomplished by providing a dedicated compartment within the tailcone. The APU and most power plant systems are installed within this compartment. Only the exhaust system and the air intake extend beyond the compartment, thus the affected fire wall cutouts incorporate additional sealing.  Various APU components, at which waste fluids (fuel and oil) may occur, incorporate provisions to collect these. In addition, the exhaust system is exposed to rain water and cleaning fluids. For safety and operational reasons, all these fluids are collected separately and drained overboard through the APU compartment drain mast.  The APU compartment itself is designed such to permit all fluids occurring to gather at the lowest point, from which they are discharged overboard through the drain mast.  Various APU components, at which waste fluids (fuel and oil) may occur, incorporate provisions to collect these. In addition, the exhaust system is exposed to rain water and cleaning fluids. For safety and operational reasons, all these fluids are collected separately and drained overboard via the APU compartment drain mast

Definition  There are two types of fire protection:  ? the active fire protection,which enables to detect the start of a fire or smoke, localized and neutralized quickly, the passive fire protection which is obtained through design precautions at each aircraft compartment level.  NOTE: This chapter describes only the active fire protection.  Active Fire Protection The various purposes of the active fire protection system are:  to detect and extinguish any fire in each engine nacelle and in the Auxiliary Power Unit (APU) compartment,  to protect the engine pylon against any torching flame from the combustion chamber,  to detect smoke in the avionics compartment, to detect smoke and to extinguish fire in lavatories, to detect any leakage from hot air ducts, to extinguish fire in passenger/crew compartments and other accessible areas in flight,  A fire can be due to excessive overheat or flammable fluid leaks and can endanger the aircraft safety. Thermo-sensitive elements detect fire or overheat conditions. They trigger the fire warnings by means of the Fire Detection Unit (FDU) when the temperature reaches the threshold of the monitored area of the engine. The ENG/APU FIRE panel (1WD) includes the controls, indications and a test function for each engine. In the section of this panel related to the engine fire detection, two functions are available:  a FIRE warning which comes on red on the ENG 1(2) FIRE pushbutton switch after a positive fire detection, a manual test capability of the system.  The components of the fire detection system of the engine are:  two fire detection loops installed in parallel in the fire zones and connected to a Fire Detection Unit (FDU). Each fire detection loop comprises three detectors connected in parallel, one FDU per engine, which processes signals from the detectors, an ENG/APU FIRE panel on the overhead panel with per engine: one ENG FIRE pushbutton switch, one TEST pushbutton switch, associated warnings and indications.  The lavatory fire-extinguishing system is installed in each lavatory service cabinet. The lavatory fire-extinguishing system discharges its extinguishant automatically when heat activates it. Any fire in the waste is kept within the confines of the metal waste-paper bin. An inert gas floods the lavatory service cabinet and extinguishes the fire. Each lavatory fire-extinguishing system has an extinguisher bottle which is self-actuated.

A. Inspection Coverage It is not possible to lay down precise details of the inspection procedure to be adopted after every incident because of the wide variations of weight, speed, nature and direction of loads, weather conditions and component failure patterns. Therefore these inspections have been written for the worst possible case. In order to gain an indication of the severity of the incident and to facilitate rapid location of primary damage, it is essential that full information is obtained from the flight crew. Therefore prior to starting an inspection, consult the crew and ascertain details of:  ? Weather conditions  ? Aircraft speed and flight attitude  ? Aircraft weight and fuel distribution  ? In landing mode, if touchdown was straight, drifting, wing low, nose high or low ? If any noise of impact or indicative of structure or component failure was heard  ? Relevant instrument indications. Printout of the maintenance data recorder tape will provide valuable additional data and indication of system malfunction.  B. Inspection Sequence To permit simultaneous inspection of several areas of the aircraft, the inspection has been divided into a number of "Packages". For example: "Inspection after Hail Impact" requires on Phase 1 a complete airframe external check. Therefore this is divided into four major zones: fuselage, wings, nacelles/pylons and stabilizers. These major zones are further divided into smaller zones or major components. The small zones are then broken down into items, which can be individually signed off as inspection is completed. The inspections are divided into three phases:  ? Phase 1 is a general inspection for primary damage and indication of remote damage and is mainly external. ? Phase 2 is a more detailed inspection and is mainly internal. Some component removal may be called up.  ? Phase 3 is a very detailed inspection involving component removal and strip down.  If the Phase 1 inspection reveals no damage, no further examination is necessary. If Phase 1 reveals damage then Phase 2 must be accomplished. If Phase 2 reveals damage then Phase 3 must be accomplished. If you find damage during Phase 1 or Phase 2 inspections, replace or repair the damaged component (Refer to the applicable manuals (SRM, AMM, CMM)) and make sure that the adjacent area has no damage.

VHF (VDR) transceiver – Operation  (1) General operation The VHF transceiver ensures its primary functions through:  The receiver The power amplifier The synthesizer The microprocessor.  The antenna switch: Routes the input signal from the antenna to the receiver in the receive mode Connects the power amplifier RF output to the antenna in the transmit mode.  The transceiver has two serial inputs: a port A serial input and a port B serial input. It can therefore be controlled through either input depending on the status of a discrete delivered by the frequency control system. The data corresponding to the frequency selected on the RMP is sent to the transceiver in the form of a 32-bit serial word through an ARINC 429 bus. This serial word contains the label, the source/destination identifier, the frequency data, the status and the parity bit. Each serial word is applied to an ARINC 429 receiver made up of differential amplifiers. These amplifiers convert the ARINC data into TTL logic signals for analysis by the microprocessor. After acquisition of these signals, the microprocessor supplies tuning data to the synthesizer in BCD form. The synthesizer processes these data and then generates a frequency to tune the receiver and the power amplifier to the RMP-selected frequency.  In Receive mode: The signal is applied to one of two low-pass filters. Selection of the low band or high band filter is made depending on the received frequency (118 to 127.975 Mhz or 128 to 136.975Mhz). These filters prevent the undesired frequencies from being fed back to the antenna. After filtering, the signal is applied to a first mixer where it is mixed with the signal from the frequency synthesizer. This first Intermediate Frequency (IF) output is then filtered by a crystal filter and amplified (40 dB of gain depending on AGC). The signal is then applied to a second mixer. This mixer receives the modulated 20.025Mhz signal and a 9.325 Mhz frequency signal produced by a crystal oscillator. The output provides a second IF signal of 10.7Mhz. This second IF signal is filtered and amplified twice before detection. Both IF amplifiers provide a gain of 70 dB. The complete amplification channel provides a total gain of 110dB. The modulated signal is then applied to a detection circuit and to a buffer amplifier. This detection circuit also controls the AGC circuit which controls the IF amplifier gain. The buffer amplifier output demodulated signal is applied to the data circuit and audio circuit.

The three systems are each pressurized by one main pump. The Green system pump is connected to the left engine and the Yellow system pump is connected to the right engine. The Green and Yellow pumps supply hydraulic power when their engine operates. The electric pump of the Blue system starts automatically when any one of the engines operates. The three system main pumps are usually set to operate permanently. If necessary (because of a system fault, or for servicing), the pumps can be set to off from the flight compartment.  Most of the components of the systems are in the three hydraulic compartments. The Green system components are in the main landing gear compartment. The Yellow system components are in the hydraulic compartment in the right belly fairing. The Blue system components are in the hydraulic compartment in the left belly fairing. The two hydraulic compartments (Blue and Yellow) are forward of the main landing-gear compartment. There are three ground service panels, one for each main system. The Blue and Green ground service panels are in the left belly fairing. The Yellow ground service panel is in the right belly fairing. All of the ground service panels are aft of the main landing-gear compartment.  All of the ground service panels have self-sealing connectors for hydraulic power supply on the ground. The ground service panel of the Green system has a hand pump, filter and selector valve so that any of the reservoirs can be filled from it. The ground service panel of the Yellow system has a hand pump and selector valve. They give an alternate method to operate the cargo doors. The ground service panel of the Blue system has the ram air turbine control panel.  The RAT of the Blue system is in a compartment in the left belly fairing, forward of the main landing-gear compartment. It extends into the airflow automatically when there is a total failure of AC busbars 1XP and 2XP. The crew can operate it from the flight compartment if necessary. The RAT supplies hydraulic power to the Blue system. If there is an electrical failure, the RAT also supplies the controlled speed motor generator (CSM/G) which gives electrical power to the aircraft.

A/THR Engage and Disengage Logic  The A/THR can only be engaged when all the necessary conditions are present and if there is a request for engagement (pilot action or automatic). (1) Required engagement conditions Two conditions are required to make the engagement possible: (a) AP/FD/A/THR common condition This condition results from the monitoring functions below: ? monitoring of the validity of the ADIRS input parameters (two ADIRS must be valid). This monitoring consists in a check of the SSM and refreshment period. The monitoring of the main parameters is made by vote or comparison. ? monitoring of LGCIUs parameters. This monitoring is made in the FAC which informs the FMGC of the validity of the LGCIU parameters. One LGCIU at least must be healthy. This condition is not required at landing or during go around. ? internal monitoring of the guidance portion healthy. ? internal monitoring of the management portion healthy. This condition is not required at landing or during go around. NOTE: The monitoring functions common to the AP/FD and A/THR are described in AP/FD Engagement.  (b) A/THR specific condition This condition specific to A/THR includes the conditions below: ? A/THR must be either in the manual speed or in the auto speed control mode ? the two ECUs/EECs must be healthy ? the FCU must be healthy ? no discrepancy between the N1/EPR target computed in the FMGC and the N1/EPR feedback from each ECU/EEC when the A/THR is active ? the various parameters used in the flight envelope protection such as VLS, VMAX, etc. must be healthy ? no action on one of the two A/THR instinctive disconnect pushbutton switches lasts more than 15s.  (3) A/THR ACTIVE logic After engagement, the A/THR is active if: ? the two throttle control levers are between IDLE and CL (CL included) ? one throttle control lever is between IDLE and CL (including CL), and the other is between IDLE and MCT (including MCT) with FLEX TO limit mode not selected ? the Alpha floor protection is active whatever the position of the throttle control levers.