A320 B2 Avionic

TRAINING CENTER

A319/320/321 B2 AVIONICS

ONLY FOR TRAINING PURPOSE

Page 6 2

This document must be used for maintenance training purposes only.

Amendment service will not be provided.

MNG Technic Maintenance Training Center

RECORD OF REVISIONS

Rev. No

Rev. Date

Revised Pages

Reason for Changes

Prepared By

Original

02.10.2006

Original

Original

Oktay TÜRKMEN

1

20.10.2009

ALL

General Revision

Oktay TÜRKMEN

2

21.05.2013

ALL

TNA STUDY

Oktay TURKMEN Arif VURUCU

__________________________________________________________________ ONLY FOR TRAINING PURPOSE

TABLE OF CONTENTS AUTO FLIGHT GENERAL......................................................................................................................................................................2-13 AUTOPILOT/FLIGHT DIRECTOR................................................................................................................................14-57 AUTO THRUST............................................................................................................................................................58-69 FLIGHT SUGMENTATION...........................................................................................................................................70-99 FLIGHT MANAGEMENT..........................................................................................................................................100-117 FAULT ISOLATION..................................................................................................................................................118-121 TESTS......................................................................................................................................................................122-133 INDICATING&RECORDING FLIGHT WARNING SYSTEM...................................................................................................................................134-135 PARAVISUAL INDICATING (PVI) ...........................................................................................................................136-139 APPENDIX...............................................................................................................................................................140-150 TROUBLE SHOOTING EXERCISES.......................................................................................................................151-320 COMMUNICATIONS AUDIO MANAGEMENT...........................................................................................................................................321-348 GROUND CREW AND COCKPIT CALL SYSTEM..................................................................................................349-354 RADIO MANAGEMENT SYSTEM............................................................................................................................355-376 VHF SYSTEM...........................................................................................................................................................377-390 HF SYSTEM.............................................................................................................................................................391-404 ACARS.....................................................................................................................................................................405-448 CIDS.........................................................................................................................................................................449-523 PRAM.......................................................................................................................................................................524-531 COCKPIT VOICE RECORDER................................................................................................................................532-540 EQUIPMENT FURNISHING – EMERGENCY LOCATER TRANSMITTER (ELT)...................................................541-552 ELECTRONIC INSTRUMENT SYSTEM (ILS) ........................................................................................................553-554 CENTRAL WARNING SYSTEMS............................................................................................................................555-577 AUDIO WARNINGS.................................................................................................................................................578-618 ELECTRONIC INSTRUMENT SYSTEM (EIS) ........................................................................................................619-632 PRIMARY FLIGHT DISPLAY (PFD) ........................................................................................................................633-634 NAVIGATION DISPLAY ..........................................................................................................................................635-638

TABLE OF CONTENTS CENTRAL WARNING SYSTEMS ...........................................................................................................................639-640 ECAM CONTROL PANEL (ECP) ............................................................................................................................641-642 SYSTEM DATA ACQUISITION CONCENTRATOR (SDAC) ..................................................................................643-648 FLIGHT WARNING COMPUTER (FWC) ................................................................................................................649-659 DISPLAY MANAGEMENT COMPUTER (DMC) – CATHOD RAY TUBE (CRT).....................................................660-676 EIS SWITCHING .....................................................................................................................................................677-688 ELECTRONIC INSTRUMENT SYSTEM .................................................................................................................689-690 EIS-TEST/BITE .......................................................................................................................................................691-702 ELECTRICAL CLOCK .............................................................................................................................................703-775 CENTRALIZED FAULT DISPLAY SYSTEM (CFDS) AND DATA RECORDING SYSTEM.....................................776-803 AIDS ........................................................................................................................................................................804-863 MULTIFUNCTION PRINTER...................................................................................................................................864-869 NAVIGATION STANDBY NAVIGATION SYSTEMS.......................................................................................................................870-875 AIR DATA / INERTIAL REFERENCE SYSTEM.......................................................................................................876-989 SATELLITE NAVIGATION.....................................................................................................................................990-1003 ILS SYSTEM........................................................................................................................................................1004-1021 ILS (MULTI MODE RECEIVER) ..........................................................................................................................1022-1047 VOR/MARKER.....................................................................................................................................................1048-1063 DME......................................................................................................................................................................1064-1079 ADF......................................................................................................................................................................1080-1095 RADIO ALTIMETER.............................................................................................................................................1096-1109 WEATHER RADAR..............................................................................................................................................1110-1123 WXR/PWS............................................................................................................................................................1124-1143 ATC/MODES........................................................................................................................................................1144-1155 TCAS....................................................................................................................................................................1156-1175 GPWS...................................................................................................................................................................1176-1193 ENHANCED GPWS.............................................................................................................................................1194-1216 INFORMATION SYSTEM INFORMATION SYSTEM.....................................................................................................................................1217-1259

MNG TECHNIC

TRAINING CENTER FOR TRAINING PURPOSE ONLY AVIONIC ______________________________________________________________________________________________________________________________________________________________________________________________

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Revision No : 02 Issue Date : 21/05/2013

A319/320/321 / TRAINING MANUAL

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Lufthansa Technical Trainin


22-00

GENERAL

SYSTEM DESIGN PHILOSOPHY This module highlights the new concept of the AutoFlight System and presents the relationship with the Electrical Flight Control Systems and the Full Authority Digital Engine Control ( FADEC ). Introduction The purpose of this module is to explain basic system design aspects included in a modern AutoFlight System. This module is not an introduction of all the functions of the system. General Concept The AutoFlight System calculates orders to automatically control the flight controls and the engines. The system only computes orders. These orders are not executed by actuators ( exept FAC for Rudder Control ) belonging to AFS but by systems which usually control the surfaces and the engines when the AFS is not active i.e. : side sticks and thrust levers. Navigation A fundamental function of AutoFlight System is to calculate the position of the aircraft. When computing A/C position, the system uses several aircraft sensors giving useful information for this purpose.

AFS/Fly by Wire The control wheel steering mode which existed in previous AutoFlight System is now ensured by the manual fly by wire mode of the Electrical Flight Control System. On conventional aircrafts the Control Wheel Steering ( CWS ) mode consists in maintaining the A/C attitude once the control wheel is released. In any case, when the automatic control of surfaces is active, if the pilot moves the stick, it disengages. System Design To meet the necessary reliability, the AutoFlight System is built around four computers. Two Flight Management and Guidance Computer ( FMGC 1 and FMGC 2 ) and two Flight Augmentation Computer ( FAC 1 and FAC 2 ). Each FMGC and each FAC has a command part and a monitor part: it is a fail passive computer. In Approach or Go Around the AFS is automaticly fail operative, if both APs are engaged.

Flight Plan The system has several flight plans in its memory. These are predetermined by the airline. A flight plan describes a complete flight from departure to arrival, it includes vertical information and all intermediate waypoints. It can be displayed on the instruments ( CRTs ). Operation There are several ways to use the Auto Flight System. The normal and recommended way to use the AFS is to use it to follow the flight plan. Knowing the position of the aircraft and the desired flight plan ( chosen by the pilot ), the system is able to compute the orders sent to the surfaces and engines so that the aircraft follows the flight plan. The pilot has an important monitoring role. Note : during AFS operation, side sticks and thrust levers do not move automatically.

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Lufthansa Technical Training


FCU

AUTOFLIGHT SYSTEM CMD

COMMAND MON

SENSORS MONITOR

FLIGHT CONTROL SYSTEM

2 FMGC 1 CMD COMMAND

MON

MONITOR

2

FADEC FAC 1

Figure 1

System Design Philosophy

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CONTROLS AND INDICATIONS IN THE COCKPIT Controls ( 1/18 ) ( 6/11 ) ( 7/10 ) (8) ( 17 )

Flight Control Panel Multipurpose Control and Display Units ( MCDU ) Radio Management Panels ( RMP ) for Navaid selection. Rudder Trim Panel Flight Control Unit ( FCU ).

Indication ( 2/15 ) ( 3/14 ) ( 13 ) (4)

Navigation Display ( ND ) Primary Flight Display ( PFD ) Engine Warning Display ( EWD ) System Display ( SD ).

Miscellaneous ( 5/12 ) Takeover and Priority pushbutton switches ( 9 ) A/THR Instinctive Disconnect pushbutton switches ( 16 ) AUTO LAND warning lights and Paravisual Display Warnings: MASTER WARN and MASTER CAUT lights.

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Figure 2

Controls and Indications in the Cockpit

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LOCATION OF THE FMGC‘S AND FAC‘S

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Figure 3

Location of FMGCs and FACs

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Lufthansa Technical Tra


AUTOFLIGHT SYSTEM PRESENTATION 1. General The auto flight system ( AFS ) installed on the aircraft is made up of two types of computers : - the flight management and guidance computer ( FMGC ) - the flight augmentation computer ( FAC ) and two types of control units : - the flight control unit ( FCU ) - the multipurpose control and display units ( MCDU ).

- acquisition and hold of a flight path - guidance of the aircraft at takeoff by holding runway axis and speed ( available in the FD as long as the aircraft is on ground ) - automatic landing and go around. The autopilot generates the following orders : - position of the control surfaces on the three axes : pitch, roll and yaw - position of the nose wheel during roll out. These orders are taken into account by these computers : FACs, ELACs, SECs and BSCU.

The functions of the FMGC are : - autopilot ( AP ) - flight director ( FD ) - automatic thrust control ( A/THR ) - flight management.

The flight director generates guidance orders used in manual control. These orders are displayed on the PFDs ( primary flight displays ) through the DMCs ( display management computers ).

The functions of the FAC are : - yaw damper - rudder trim - rudder travel limiting - calculation of the characteristic speeds and flight envelope monitoring - acquisition of the yaw AP order. The MCDUs linked to the FMGCs enable : - the introduction and the modification of the flight plan - the display, the selection and the modification of the parameters associated with the flight management function. The FCU is used for : - the engagement of the AP/FD and A/THR systems - the selection of flight parameters ( altitude speed/Mach, vertical speed/flight path angle, heading/track ) - the selection of AP/FD modes. This system description describes the autopilot ( AP ) and the flight director ( FD ) functions, which are : - stabilization of the aircraft around its center of gravity when the AP/FD system holds vertical speed or flight path angle and heading or track

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PFD DMCs ND

EFIS

To ELAC‘s

ND DMCs PFD

EFIS

Figure 4

Layout of AFS Components

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ELECTRICAL POWER The AFS components are supplied by electrical power supply systems as defindet on figure 5. ”CAT 3 DUAL” is indicated, if contactor BTC 1 (11XU1) and BTC 2 (11XU2) and 1PC2 (DC BAT BUS - DC BUS 2) are open. The APU-generator is not accepted for ”CAT 3 DUAL” operation.

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Figure 5

Electrical Power

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FMGC - INPUT / OUTPUT DISCRETES

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AUTO FLIGHT DISCRETES/ANALOG INTERFACES FMGC-INPUT / OUTPUT DISCRETES

22-85-00

SCHEM 04 Page 101 Nov 01 /89

Figure 6

FMGC - Discretes

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22-10

AUTOPILOT / FLIGHTDIRECTOR

AUTOPILOT INTRODUCTION General The autopilot is engaged from the Flight Control Unit by the related pushbuttons. The autopilot guidance modes are selected from the Flight Control Unit or the Flight Management and Guidance Computer. The autopilot function is a loop after a comparison between real and reference parameters, the FMGC computes orders which are sent to the Flight Controls. The loop is closed by real values coming from sensors and given by other systems ( ex : ADIRS ) to the FMGC. When the autopilot is engaged, the load thresholds on the side sticks and pedals are increased. If a side stick is overriden or the Takeover and Priority P/BSw is depressed the autopilot disengages. When AP is engaged : on the side sticks, the pitch and roll load threshold changes. Any force exeeding this tresholds disengages the AP. on the rudder pedals, the load threshold changesalso in the artificial feel and trim unit. Exeeding this threshold results on AP disengagement.

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Figure 7

FMGS - Components

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AUTOPILOT INTRODUCTION ( CONT. ) Modes There are lateral modes and vertical modes. Basically, one of each is chosen by the pilot or by the system. The AP being engaged, a lateral mode and a vertical mode are simultaneously active. According to the flight phases, the lateral mode controls the aileron via Elevator Aileron Computers (ELACs), the spoilers via ELACs and Spoiler Elevator Computers (SECs), the rudder via Flight Augmentation Computers (FACs) and the nose wheel via ELACs and the Braking/Steering Control Unit (BSCU). The vertical mode controls the elevators via ELACs. Autopilot Operation on Ground For maintenance purposes, the autopilot can be engaged on the ground only with both engines shut down. Hydraulic power is not required. When an engine is started, the autopilot disengages. Autopilot Operation at Take-Off The autopilot can be engaged in flight, provided the aircraft has been airborn for at least 5 seconds. Before autopilot engagement, take-off modes can be active for the flight director. Autopilot Operation at Cruise In cruise, only one autopilot can be engaged at a time. Ailerons and Spoilers execute the orders of lateral modes, Elevators execute the orders of vertical modes. Engaging a second AP in cruise disengages the other one. Note : The rudder is not controlled by the AP, but by Flight Augmentation Computer ( FAC ) functions.

Autopilot Operation during Landing If the airfield is equipped with ILS installations, the autopilot can perform a complete landing, roll out included. In addition, the autopilot controls the rudder via the Flight Augmentation Computer. ILS approach : AP is able to perform a complete landing with descent, flare and roll out. A second AP can be engaged (AP 1 active, AP 2 backup ). After landing, the autopilot gives steering orders for the nose wheel. Roll out : Steering order to rudder and nose gear depend on aircraft speed. Ailerons and spoilers AP orders are null. Note : spoilers are directly controlled by SECs as airbrakes. During roll out, at low speed (about 60 kts), the pilot normally disengages the AP function(s) by pressing a take over pushbutton located on the side stick. If the airfield has no glide slope installation, the pilots can select a LOC or a NAV approach, but the autopilot is disengaged at a given altitude. LOC (without glide) or NAVigation approach : same principles as for cruise. Pilots have to disengage AP at a given altitude in order to land manually.

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/ EPR

Figure 8

FMGS - Architecture

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AUTOPILOT INTRODUCTION (CONT.) Nose wheel control Each ELAC receives a nose wheel steering command from the two FMGCs : DELTA ( NOSE WHEEL The ELACs select one off the two commands in function of : AP engagement The selected command is sent to the BSCU. The BSCU uses this command associated with commands from the control wheel and rudder pedals to compute nose wheel control angle. The command from the FMGC and the command from the rudder pedals are limited with respect to the speed. The command from the FMGC is used after landing during rollout. The BSCU generates four discretes ( BSCU HEALTHY ) whose validity is taken into account : For capability computations In the ROLL OUT logic. It also supplies 2 discretes (wheel speed ) for the ROLL OUT logic. Using the ‘Pedal Disconnect P/B‘ on the Handwheel prevents nosewheel movement e. c. during full ruddertravel in Take off ( Crosswind ).

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+ +

to Rudder

+ +

Figure 9

+

Nose Wheel Control

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FD INTRODUCTION Interface with DMCs and Reconfiguration Each DMC receives : - a bus from each FMGC on which are routed : the FD orders, the AP/FD engagements, the AP/FD modes, the landing capabilities, - a wired discrete per FMGC giving the engagement status of the FDs, - a bit on a discrete label of the FCU corresponding to the action on the FD pushbutton switch associated with the PFD. A logic inside the acquisition module selects the FMGC bus(es) required to present the FD orders and the FMA messages. In normal operation : - the DMC 1 transmits data to Capt PFD - the DMC 2 transmits data to F/O PFD. Each DMC is linked to its associated PFD by two connections ( a normal one and an alternate one). The alternate connection is used for different types of changeover. DMC / PFD Reconfiguration In the event of a DMC failure, the DMC 3 in standby can replace the faulty DMC after action on the EIS DMC selector switch. In the event of a PFD failure, the data are transferred automatically from the PFD to the ND ( data on PFD have priority ). This transfer can also be made manually in two ways: - by turning the PFD potentiometer to ” OFF ” - by action on the PFD / ND XFR pushbutton switch. Automatic Selection of FMGC Bus for the FD Orders Each DMC makes a selection depending on the side on which it is installed and on the validity of each FD, in function of: - the engagement wired discretes - the status matrices (SSM) of the labels which the FD orders are routed. So the DMC1 (2) selects the FMGC1 (2) bus if the FD1 (2) is valid. The PFD1 (2) therefore displays: - the FD1 (2) message on the FMA - the FD orders from the FMGC1 (2).

Automatic FD Reconfiguration If an FD1 (2) validity loss is detected by the DMC1 (2) through: - loss of the FD1 (2) ENG condition - non refresh of FMGC1 (2) labels - status matrix of FMGC1 (2) labels coded at F/W status the DMC1 (2) will select the data from the FMGC2 (1) automatically and will display: - FD2 (1) message - the FD orders from the FMGC2 (1). FD Order Removal All the FD orders can be cleared by the DMC in one of the following cases: - action on the corresponding FD pushbutton switch on the FCU - validity loss of both FDs. The DMC clears a given FD order when the associated label is NCD. Selection of FMGC Bus for Display of AP/FD Modes and Landing Capabilities This selection depends on the engagement of the AP/FD systems. FD only engaged : Each DMC utilizes the bus selected for the FD orders as per the logic described in ” Automatic Selection of FMGC Bus for the FD Orders ”. Only one AP engaged : Each DMC utilizes the FMGC bus which corresponds to this AP. Each PFD displays : - AP1 or AP2 message depending on the AP engaged, - the modes corresponding to this AP, - the landing capabilities from the FMGC corresponding to the AP engaged. Both APs engaged : Each DMC is associated with the corresponding FMGC. Therefore the Capt (F/O) PFD displays : - AP1 + 2 message, - the modes corresponding to AP1 (2), - the landing capabilities from the FMGC1 (2). FD Flag ( red ) In case of both FMGC‘s failure or both FD disengaged with FD pushbutton ” ON ” and attitude valid, a red FD - flag is displayed.

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Figure 10

Interface between FMGCs and DMCs

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AUTOPILOT / FLIGHT DIRECTOR - MODES Modes General Engagement Principle : The engagement of the cruise modes on the AP/FD follows the operational utilization principle of Automatic Flight System ( AFS ). When the pilot wants to control a flight parameter manually, he must select the required value on the FCU then pull the associated selector knob. Then, the AP/FD mode of the manual control of this parameter is engaged.

Disengagement Principle : The disengagement of a lateral mode is caused by the engagement of a new lateral mode. During RUNWAY LOC sub-AFS mode ( FD Roll-Takeoff mode ) when a discrepancy between CMD and MONG channels leads to FD disengagement. The disengagement of a longitudinal mode is caused by the engagement of a new longitudinal mode. Each mode ( lateral or longitudinal ) is disengaged at the engine running on ground or at the confirmed loss of AP/FD for more than 0.6 s.

In order to have a flight parameter controlled by the FM part of the FMGC the pilot must push the associated selector knob. The automatic control is then armed or activated. Synchronization of Modes between FMGCs : So as to ensure a consistent operation of the AFS, it is mandatory to have the two FMGCs in operation of the same modes active and armed. The logic for the selection of the FMGC which has priority takes into account the engagement of the AP/FD and A-THR functions ( see Fig. on next page ). In cruise phase there is at least one AP/FD engaged, the FMGC which has priority imposes the cruise modes active and armed to the FMGC which has no priority Engagement on the Ground : In order to facilitate the AFS test, certain cruise modes can be activated on the AP and on the FD, on the ground when the engines are stopped. All these modes are disengaged at engine start-up on the ground and this causes the return to a configuration in conformity with the takeoff phase.

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AP ENGAGEMENT

FD ENGAGEMENT

A-THR ENGAGEMENT FMGC having priority 1 2

1

2

1

2

1

-

-

-

-

-

1

0

1

-

-

-

-

2

0

0

1

-

-

-

1

0

0

0

1

-

-

2

0

0

0

0

1

-

1

0

0

0

0

0

1

2

0

0

0

0

0

0

1 ( if valid )

NOTE:

” - ” means: whatever the state

Figure11 FMGC Priority Logic

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AUTOPILOT / FLIGHT DIRECTOR - MODES ( CONT. ) Mode Selection Principle A mode can be selected through one of the following possibilities : Automaticly, e.g. the altitude acquisition mode is always armed, exept in some cases ( approach ). Action on pushbutton switch located on the FCU. Push or pull action on one of the reference selection knobs ( speed / mach, heading / track, altitude, vertical speed/flight path angle ) on the FCU. CRUISE FLIGHT

LONGITUDINAL

LATERAL

Cancellation of an engaged mode. Position of the throttle control levers ( selection of TO or GARD modes ). AP - A/THR Mode Compability The AFS is such that the AP/FD system or the A/THR function always control the speed. The AP/FD has the priority. To do this, the modes of the A/THR system are function of the AP/FD-longitudinal modes. The table below presents the Cruise Modes.

MODE

AVAILABILITY

PHASES

NOTE

- Vertical speed ( V/S ) ( Acquisition and Hold )

AP / FD

HOLD

Automatic or V/S - FPA select knob

- Flight path angle ( FPA ) ( Acquisition and Hold )

AP / FD

HOLD

- Altitude acquisition ( ALT ACQ )

AP / FD

- Altitude hold ( ALT )

AP / FD

HOLD

- DES - OP DES

( Descent ) ( Open Desct )

AP / FD AP / FD

ARM - HOLD HOLD

- CLB - OP CLB

( Climb ) ( Open Climb )

AP / FD AP / FD

ARM - HOLD HOLD

- Heading ( HDG ) - Track ( Acquisition and Hold )

AP / FD AP / FD

HOLD HOLD

- Navigation ( NAV )

AP / FD

ARM - CAPTURE

ARM - HOLD

Armed automatically

Automatic on selected Altitude Altitude select-knob

Automatic or HDG / TRK selectknob ( pulled ) HDG / TRK selectknob ( pushed )

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COMMON MODES ( Takeoff, Landing, Go Around )

COMMON MODES

TAKEOFF

( TO )

LONGITUDENAL MODES All engines operational : Speed Reference System ( SRS ) : Holding of V2 + 10 kts

LATERALE MODES

AVAILABILITY

Runway ( RWY ): - Holding of LOC centerline up to 30 ft RA,

FD

- Track above 30 ft RA One engine fail :

PHASES

HOLD AP*/ FD ( *AP only 5 sec after lift off )

SRS : Holding of Va if Va > V2 V2 if Va < V2 ( Va : Actual Speed ) GO AROUND ( GA )

SRS : Holding of Va if Va > Vapp or Vapp if Va < Vapp

Track

LOCALIZER ( LOC )

APPROACH ( APP )

Glide capture, track ( GS ), Flare, Rollout or Final desct ( FINAL ) according to the profile determined by the FMGC ( Appr. Page )

AP / FD

HOLD

LOC capture and track

AP / FD

ARM - CAPT - TRACK

LOC capture and track Align and Rollout or R - NAV approach or VOR approach

AP / FD

ARM - CAPT - TRACK

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TRAINING CENTER FOR TRAINING PURPOSE ONLY AVIONIC ______________________________________________________________________________________________________________________________________________________________________________________________ FRA

AUTOPILOT / FLIGHT DIRECTOR - MODES ( CONT. ) The operational use of the AFS is based of the following principle : The short-term pilot orders are entered through the FCU The long.term pilot orders are entered through the MCDU. This principle leads to two distinct operations : Selected and managed controls. Selected Control The aircraft is controlled using reference parameters entered by the pilot on the FCU ( heading / track, vertical speed / flight path angle, speed / mach, altitude ). These parameters are taken into account ( acquisition and then hold ) as follows: Modification of the parameter by means of the corresponding selector knob on the FCU. Pull action on the selector knob. Managed Control The aircraft is controlled using reference parameters computed by the FMGC which takes into account the pilot data selected on the MCDU. A parameter is selected in managed control by pushing the corresponding selector knob. In this case the parameter value is called out by means of a dashed line on the FCU ( exept altitude which is always displayed ) and a white indicator light comes on near the coresponding referens display.

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Figure 12

Managed and Selected Control

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AP ENGAGEMENT General The AP is engaged through two pushbutton switches ( AP 1 and AP 2 ) located on the center section of the FCU. In cruise only one AP can be engaged at a time ( priority to the last AP engaged ). Both APs can be engaged when APPR and GO AROUND modes are selected. In these cases, the AP 1 has priority and is active. The AP 2 is in standby and becomes active if the AP 1 is lost. When these modes are released, the AP 2 is disengaged automatically. The AP can be engaged on the ground in any mode with engines stopped. The AP disengages when one engine is started. An AP can be engaged again 5 s after lift-off : - In active FD modes ( if at least one FD is engaged ) - In HDG and V/S modes ( if no FD is engaged ). At AP engagement, the load thresholds on the side stick controllers and on the rudder pedals are increased. AP engagement is indicated by the illumination of the corresponding pushbutton switch ( three green bars ) and by the AP 1 or AP 2 indication in the status column on the PFDs. The pilot can disengage the AP in different ways: - By action on the engagement pushbutton switch, with the green bars on. - By action on one takeover and priority pushbutton switch on the side stick controller.

AP-engage hardware logic Principle: A part of the AP engage logic is accomplished through the hardware. It takes into account the following signals : - AP ENGD boolean generated in the software - FG HEALTHY logic signal - AP SW wired discrete from the FCU. The AP-engage hardware logic utilizes the command and the monitoring channels. Each output discrete takes into account the conditions generated by each channel. During the safety tests ( at power up ) the AP SW signal is inhibited prohibiting engagement through the pushbutton switch. The disengagement takes place in the hardware logic : - Upon loss of one of the AP ENGD and FG HEALTHY signals after confirmation of 200 ms - Through action on one takeover and priority pushbutton switch located on the side stick controllers - Upon detection of Long Power Failure ( LPF ) by the power unit . In the event of short interruption, the engage signal maintains its pre-cutoff state. The final circuits are therefore supplied with back-up current ( VS ). They are isolated from the other signals during the cutoff ( SW RESET signal active ). The AP ENGD wired discretes obtained are used by: -The Elevator Aileron Computers ( ELAC ) ( selection of AUTO mode ) -The FCU ( illumination of the corresponding AP pushbutton switch ) -The opposite FMGC ( disengagement of associated AP if in cruise modes, selection of the FMGC having priority ) -The FMGC OWN ( engagement wrap around ). -The FWCs ( generation of the AP warning ).

Loss of the AP is indicated by an aural and visual warning.

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Confirmation and Trouble Schooting Data

LPF - Long Power Failure

CONFIRMATION AND TROUBLE SHOOTING DATA

Figure 13

AP-Engage Hardware Logic

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AP ENGAGEMENT ( Cont. ) AP-engage software logic. Engagement conditions. This signal is at 1 ( F/F set ) if all the engagement conditions are activated : - Action on the engagement pushbutton switch. - Ground conditions; engagement possible in any mode only if all engines are stopped. - Flight conditions; engagement possible 5 s after lift-off. - Conditions specific to the AP : AP COND. - Conditions common to the AP/FD : AP/FD COND. - Conditions common to the AP/FD and A/THR : AP/FD, A/THR COND. Disengagement conditions. These are : - Action on the engagement pushbutton switch, the associated AP being already engaged - or action on one takeover and priority pushbutton switch - or one engine start on the ground - or loss of one condition: either AP COND, or AP/FD COND or AP/FD/A/THR COND - or in the event of landing in dual-AP operation, disengagement of AP 2 only when the LAND or GO AROUND mode is released - or engagement of the opposite AP if the AP is not in LAND or GO AROUND mode.

AP-specific conditions Disengagement through AP takeover and priority pushbutton switches. Availability and validity of peripherals. These are peripherals which utilize the AP commands. FAC : - Availability of at least one FAC ( CMD and MON FAC HEALTHY wired discretes ). - confirmation of FAC operation in AUTO mode further to AP engagement by the FAC - engagement of the yaw damper function - engagement of the rudder trim function. Loss of one of the above five logic conditions is not taken into account in LAND TRACK, between 100 ft. and the ground. ELAC : Each ELAC generates ELAC AP DISC discretes. The AP disengages only upon a command from the two ELACs. - The pilot takes control by overriding the load thresholds of the side stick controller - or both ELACs not healthy - or servoloops not healthy - or high or low speed protections are reached - or the Alpha floor protection is active - or the roll angle is > 45 - or the EFCS abnormal laws are engaged ( direct or alternate ). . The disconnection command from only one ELAC results in a reduction of landing capability.

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Confirmation and Trouble Shooting Data

LPF - Long Power Failure

CONFIRMATION AND TROUBLE SHOOTING DATA

Figure 14

AP-Engage Hardware Logic

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SIDE STICK CONTROLLERS AND RUDDER PEDALS Increase of load thresholds on side stick controllers and rudder pedals When the AP is engaged, the command and the monitoring channels supply the relays which control the side stick lock solenoids ( the command channel provides the +28 V, the monitoring channel provides the ground ). Each control has its own solenoid. Each AP has its own relays and can therefore lock the controls. Side stick controllers: The loads are increased on both axes. Any load on the side stick controller which exceeds these values, results in AP disconnection ( wired discrete from the ELACs ). Rudder pedals: The load is applied on the rudder artificial feel ( addition of a spring in the artificial feel and trim unit ). Exeeded load results not in an AP disconnection.

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Figure 15

Side Stick Controllers and Rudder Pedals - Locking

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WARNINGS AP Warnings When the AP is disengaged,a warning is provided: If the disengagement is manual through the TAKE-OVER-PB on the SIDE STICK, the visual and associated warnings are temporary. If the disengagement is due to a FAILURE, an ACTION on the FCU PB or FORCE on the SIDE STICK the visual and audio warnings are continuous.

Autoland Warning When ” LAND ” appears in green and at least one AP is engaged, the AUTOLAND red light appears on the glareshield when the aircraft is below 200 ft RA and one of the following events occurs: The APs are lost, or the aircraft gets too far off the beam, or the localizer or glide slope transmitter or receiver fails, or the difference between both radio altimeter indications is greater than 15 ft.

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Figure 16

Location - Warnings

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FLIGHT DIRECTOR ENGAGEMENT The Flight Director ( FD ) generates guidance orders used in manual control and displays these orders on the Primary Flight Display. Engagement The Flight Director is engaged upon energization. Energization on the GROUND : - After the safety tests at power rise : Both FDs engage if no failure is detected by internal monitoring automatically. The white ”1 FD 2” indication appears on each Primary Flight Display ( PFD ), but the FD bars are removed. If a FD does not engage ( FMGC failure detected by internal monitoring ), both PFDs are automatically switched to the valid FD ( FD indication: ”2 FD 2” if FD #1 fails or ”1 FD 1” if FD #2 fails on both PFDs ). Let us see the FD engagement in case of energization in flight.

Display Logic There are three types of FD bars : PITCH BAR, ROLL BAR, YAW BAR. The horizontal PITCH BAR does not appear if there is no active vertical mode or in rollout phase of LAND mode. The vertical ROLL BAR does not appear in ROLLOUT mode or in RUNWAY mode up to 30 feet. The YAW BAR only appears in RUNWAY mode, up to 30 feet, and during LAND mode, align or rollout phases. This bar is said to be centered when just below the central yellow square.

Energization in FLIGHT : The safety test at power rise is not performed. The two FD‘s engage in Vertical Speed ( V/S ) and Heading ( HDG ) modes if no AP is engaged. Flight Director Pushbutton FD pushbuttons, located on the EFIS control panels of the Flight Control Unit, allow the Flight Director symbols to be removed from the Primary Flight Display.

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HDG - V/S SELECTION

TRK - FPA SELECTION

Figure 17

Flight Director Selection and Indications

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FLIGHT DIRECTOR DESCRIPTION OPERATION ( CONT. ) Bar Display The Flight Director bars can be used provided heading/vertical speed is selected on the FCU. HDG / VS displayed at power up. AP/FD modes are correctly followed, when the FD bars are centered on the fixed aircraft symbol of the PFD. The FMGCs can send a command to the DMCs to make the FD bars flash for 10 seconds. The FD pitch and roll bars flash in the following conditions : When one AP or FD is engaged, when no AP/FD were previously engaged If V/S and HDG modes are engaged with approach modes engaged or in the NAV mode associated, with RNAV approach When the glide data is lost above 100 feet with approach modes engaged the PITCH BAR flashes When the LOC data is lost above 15 feet with approach modes engaged the ROLL BAR flashes.

Flight Path Director ( FPD ) Symbol The Flight Path Director can be used provided track/flight path angle ( TRK/ FPA ) is selected on the FCU. The Flight path director symbol shows the pilot how to intercept and fly the vertical and lateral flight path. Flight Path Vector ( FPV ) Symbol The Flight Path Vector symbol represents the track and flight path angle actually being flown. When the pilot superimposes the FPV and the FPD symbols, the aircraft is flying the commanded trajectory.

Yaw Bar Symbol The yaw bar appears in take-off and landing phases and is identical to the FD bar case. It only appears in RUNWAY mode up to 30 ft RA and during align ( at 30 ft ) and rollout phases of LAND mode on ground.

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HDG - V/S SELECTION

TRK - FPA SELECTION

Figure 18

Flight Director Selection and Indications

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FMA DESCRIPTION AP, FD, A/THR ENGAGEMENT STATUS

FLIGHT MODE ANNUNCIATOR ( FMA ) The flight mode annunciator ( FMA ), which is just above the primary flight displays, shows the status of the autothrust, the vertical and Iateral modes of the autopilot and flight director, and the approach capabilities, and the engagement status of the AP/FD and the autothrust. After each mode change, the FMA displays a white box around the new annunciation for ten seconds. In the three Ieft columns: The first Iine shows the engaged modes in green. The second Iine shows the armed modes in blue or magenta. Magenta indicates that the modes are armed or engaged because of a constraint. The third Iine displays special messages: —Messages related to flight controls have first priority : MAN PITCH TRIM ONLY in red, flashing for 9 seconds, then steady USE MAN PITCH TRIM in amber, pulsing for 9 seconds, then steady —Messages related to the FMGS have second priority.

Figure 19

The fourth column: Displays approach capabilities in white. Displays DH or MDA / MDH in blue. The fifth column: Displays the engagement status of AP, FD, and A/THR in white. Displays a box around FD for 10 seconds in case of automatic FMGC switching. Displays A/THR in blue when autothrust is armed but not active. Note: When one AP is engaged, the master FMGC drives both FMAs. If no AP is engaged, each FMA is driven by its onside FMGC. (The onside FD pushbutton must be ON to display AP/FD modes and approach capabilities).

FMA

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Figure 20

FMA

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Figure 21

FMA

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Figure 22

FMA

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Figure 23

FMA

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Figure 24

FMA

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LANDING CAPABILITY CONDITIONS Landing Capabilities Whatever the flight phase, each Flight Management Guidance Computer ( FMGC ) computes its own automatic landing capability according to the availability of the varius sensors and functions. According to this capability, each FMGC computes the landing capacity which takes into account information from both FMGC‘s. When the AP and FD are disengaged for one FMGC, the landing capability corresspons to the category of the only FMGC likely to provide automatic landing. When the AP or FD is engaged for the two FMGC’s, the landing capability corresponds to the lowest category coming from the two FMGC‘s. The master FMGC then sends the category of landing to be displayed on both Primary Flight Displays ( PFD, on FMA ) via the Display Management Computers ( DMC ). The LAND 3 FAIL OPERATIONAL capability is obtained, when both FMGC‘s have the LAND 3 FAIL OPERATIONAL category. In this configuration, the objective is to continue automatic landing in spite of the simple failures which might affect the various systems used during this phase. NOTE : Below 100 ft RA, LAND 3 FAIL PASSIVE and LAND 3 FAIL OPERA TIONAL categories are memorized, until the LAND TRACK mode is disengaged or the 2 AP‘s are disengaged. A failure occurring below 100 ft does not cause any capability down grading. The CAT 1, CAT 2, CAT 3 SINGLE and CAT 3 DUAL messages are displayed on the FMA according to the landing capabilities send by the FMGC‘s.

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FMGC 1

FMGC 2 FMGC 1

FMGC 1

Figure 25

Land Capability - Block Diagram

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LAND 2 Capability One Autopilot engaged involves the validity of at least one Elevator Aileron Computer (ELAC) and one Spoiler Elevator Computer (SEC) including hydraulic condition. The output ”LAND 2 CONDITION” is used in the logic of the land 3 ”fail passive” capability. LAND 3 FAIL PASSIVE Capability Land 3 Fail Passive Capability is frozen below 100 feet as long as one Autopilot remains engaged in LAND TRACK. A failure occuring below 100 feet does not cause any Capability Downgrading. For details and meaning of the ” LAND 2 COND ” input refer to the logic of land 2 capability. The ” RESET CAPABILITY ” output is used in the logic of land 3 ” fail operative ” capability. LAND 3 FAIL OPERATIONAL Capability Each command/monitoring FMGC channel performs Landing Capabilities. Land 3 Fail Operational Capability is frozen below 100 feet as long as one Autopilot remains engaged in LAND TRACK. A failure occuring below 100 feet will thus not cause any Capability Downgrading. Note that the ” EFCS in FAIL OP status ” means that there is a redundancy of hydraulic systems and of surfaces. POWER SUPPLY SPLIT condition means that both power supplies must be dissociated and provided by independent buses. The ” RESET CAPABILITY ” input is detailed in the logic of land 3” fail passive ” capability. Capacity Downgrading Display Land category INOP messages are displayed on ECAM status page, on the right column. Downgraded land category messages are displayed on ECAM STATUS page, on the left column.

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LAND 3 FAIL OPERATIVE CAPABILITY

LAND 3 FAIL PASSIVE CAPABILITY

LAND ARM AND LAND ARM OPP OVER 400FT LAND TRACK AND LAND TRACK OPP

LAND 2 COND

AP AND AP OPP ENG A/THR OR A/THR OPP ENG

S

FWC OWN AND OPP VALID POWER SUPPLY SPLIT ENGINE STOPPED

R

LAND 3 FAIL OPERATIVE CAPABILITY

A/THR ENG A/THR OPP ENG S RA OWN VALID RA OPP VALID

PFD OWN AND OPP VALID

R

LAND 3 FAIL PASSIVE CAPABILITY

LAND 3 FAIL OP CAPABILITY

NO ELAC AP DISC IR OWN,OPP AND 3 AVALID RA OWN AND OPP VALID ILS OWN AND OPP VALID ADR OWN,OPP AND 3 VALID FAC OWN AND OPP HLTY

RA > 100FT RESET CAPABILITY

LAND TRACK LAND TRACK OPP

BSCU VALID

RESET CAPABILITY

NO AP ENG

ADIRS MONITORING BY FAC OWN AND OPP OK YAW DAMPER OWN AND OPP HLTY RUD TRIM OWN AND OPP HLTY

LAND 2 CAPABILITY LAND ARM OVER 400FT OR LAND TRACK AP ENG LAND ARM OPP OVER 400FT OR LAND TRACK OPP AP OPP ENG

FWC OWN VALID FWC OPP VALID PFD OWN VALID PFD OPP VALID ILS OWN OK

LAND 2 CAPAB

ILS OPP OK

LAND 3 FAIL PAS CAPABILITY LAND 3 FAIL OP CAPABILITY

Figure 26

LAND 2 CAPABI LITY

Partial Landing Capabilities

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FLIGHT CONTROL UNIT Purpose In general, the Flight Control Unit (FCU) provides the short term interface between the crew and the Flight Management and Guidance System. The FCU comprises three panels : - one center panel (auto flight control section) which features the controls and the displays associated with the AFS. - two symmetrical panels (EFIS control sections) located on the left side and on the right side of the center panel. These panels include the controls and the displays associated respectively with the Captain and the First Officer EFIS display units. The FCU it is located on the glareshield. The FCU is the main interface to engage functions and guidance modes and to select parameters. The FCU allows: Engagement of Autopilots, Flight Directors and Autothrust. Selection of Guidance modes; e.g. Heading,Vertical Speed or Track Flight Path Angle. Selection of Flight parameters; e.g. Speed, Altitude, Mach. FCU-Reconfiguration The FCU consists of two identical computers ( FCU #1 and #2 ) totally independent. The computers ( SIDE 1 and SIDE 2 ) have separate power supplies. Each side is associated with the controls on the front panel of the unit. The display is common to both sides, whereas the signals are routed via separate paths. Only one is active at the time, the other is in standby for AFS but controls his Baro-correction. When both channels fail, all FCU controls are inoperative. AUTOTHRUST, AP/FD 1 AND AP/FD 2 are not available.

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Figure 27

Flight Control Unit (FCU)

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FLIGHT CONTROL UNIT (CONT.) Changeover In order to ensure segregation of barometric selections and displays, the CAPT and F/O BARO parameters are controlled, in normal operation, independently by the two different FCU - processors. If both FCUs are healthy, the FCU 1 is active and controls Capt BARO selection, AFS display, AFS and EFIS pushbutton switches as well as ARINC 1 bus. The FCU 2 controls only F/O BARO selection and ARINC 2 bus. When FCU 1 is failed, there is a changeover on FCU 2 which becomes full active. It then controls the whole FCU. When FCU 2 is failed, FCU 1 remains active and also controls F/O BARO selection and ARINC 2 bus.

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FCU 2 NOT VALID

FCU 1 NOT VALID

Bus #1

FCU DISPLAYS

Bus #2

FCU 1 NOT VALID FCU 2 NOT VALID

FCU 1 NOT VALID

FCU

FCU

Figure 28

Changeover Block Diagram

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FCU PHILOSOPHIE The guidance targets and their associated modes are of two types: managed by the FMGS selected by the crew Those managed by the FMGS are indicated by dashed windows with the associated dot illuminated white. Those selected by the crew are displayed on the windows with the associated dot extinguished. Note: ALTITUDE WINDOW IS NEVER DASHED AND ALWAYS DISPLAYS PILOT SELECTED ALTITUDE.

FCU PANEL DESCRIPTION

There are four selector knobs: SPD-MACH HDG-TRK ALT V/S-FP A Selector knobs can be turned and pulled or pushed ( exept on A 320 V/S-FPA which cannot be pushed ).

APRR mode engagement P/B Sw Arms, disarms, engages or disengages approach modes.

AP 1 - AP 2 Engage P/B Sws Engages or disengages autopilot functions. Illuminated green when the AP is engaged. A/THR - Engage P/B Sw Arms, activates or disconnects the autothrust functions. Illuminated green if the A/THR is armed or active. Meter selection P/B Sw Used to display the FCU selected altitude target and QNH in meters on ECAM.

LOC mode engagement P/B Sw Arms, disarms, engages or disengages the LOC mode.

In order to arm/engage managed guidance the pilot must push the associated selector knob e.g. HDG selection knob pushed = NAV mode engaged/ armed. In order to engage a selected guidance mode, the pilot has to turn (to set the required value) then pull the selector knob to engage the mode on the selected target. In managed guidance (window dashed), turning the selector knob (without pulling it) dispays the set value for 45 seconds ( A 320 for 10 seconds) in the HDG-TRK and V/S-FPA windows and 10 seconds in SPD/MACH window

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SPEED/MACH control area

Lateral control area

AP-A/THR control area

Vertical control area

SPEED/MACH WINDOW

HDG/TRK WINDOW

AP1-AP2 pb

ALTITUDE WINDOW

-shows selected SPD or MACH in ”selected guidance” -shows after PWR UP: SPD 100

-shows selected HDG or TRK in ”selected guidance” -shows after PWR UP: ---”---” in ”managed guidance” -Display range: between 0 -359 deg.

-engages or disengages autopilot function

-always displays a target value selected by the crew. The window is never dashed

HDG TRK selector knob

A/THR pb

-”---” in ”managed guidance” -Display range: between 100 and 399 KT for speed, between 0.10 and 0.99 for MACH number SPD/MACH selector knob

-Knob pushed: armes/engages NAV for ”managed guidance”

-Knob pushed: engaged SPD/MACH for ”managed guidance”

-knob pulled: engages HDG or TRK in ”selected guidance”

-Knob pulled: engaged SPD/MACH for ”selected guidance” SPD/MACH pb

-illuminated green when the AP is engaged

-arms, activates or disconnects the autothrust -illuminated green if the A/THR is armed or active

Altitude selector knob (outer and inner) -outer knob has 2 selectable positions: 100ft or 1000ft -inner knob sets the altitude in the FCU window - knob pushed: CLB / DESCT. - knob pulled: OPEN CLB /OPEN DESCT. ECAM meter selector push button -is used to display the FCU altitude target in meters on the ECAM

LOC mode engagement pb Arms, engages or disengages the LOC mode

V/S-FP A window -shows selected V/S or FPA in ”selected guidance” -shows ”---” in manged guidance

-Depressing this pb changes SPD target to corresponding MACH target and vice versa (automatic on FL 305)

V/S-FP A selector knob -turning sets V/S or FPA value to be displayed in the V/S/FPA window. V/S range: -6000 +6000 ft/min FPA range: -9.9 +9.9 deg. -knob pushed: engages an immediate Level Off (V/S or FPA=0) -knob pulled: engages V/S or FPA mode APPR push button -arms, disarms, engages or disengages approach modes

MANAGED DOT

HDG-V/S/TRK-FP A PB SPD/MACH PB

ECAM METER SELECTOR PB

Figure 29

Panel Description

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ECAM - FCU Warnings.

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FCU 1 AND 2 FAULT

FCU 1 (2) FAULT

CAT 2 only

Figure 30

CAT 3

FCU Warnings

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22-30

AUTOTHRUST

Purpose The autothrust ( A/THR ) system is part of the auto flight system .The autothrust system ensures the functions below through the control of the thrust : speed hold ( selected by the pilot : ”manual control” or computed by the FMS: ”auto control” ) Mach hold ( selected by the pilot : ”manual control” or computed by the FMS: ”auto control” ) thrust hold thrust reduction during flare-out ( RETARD ) protection against excessive angle of attack ( ALPHA FLOOR protection ) The A/THR is integrated in the Flight Management and Guidance System. The Engine Interface Units ( EIUs ) and the Electronic Control Units ( ECUs ) / Electronic Engine Control ( EECs ), ensure the link between this system and the engines. The use of digital engine control units permitted to simplify the autothrustsystem through : the deletion of the autothrottle actuator ( use of a digital link between the FMGC and the ECUs/EECs ) the deletion of the limit thrust computation ( already performed by the ECUs/EECs ) the deletion of the limit thrust panel ( the ECUs/EECs make this selection automatically depending on the position of the throttle levers ). the deletion of the TO/GA levers ( the engagement of these modes is made through push action on the throttle control levers ).

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-----------------------------------------------------------------------------! ENGAGEMENT OF AP ! ENGAGEMENT OF FD ! A/THR ACTIVE ! !---------------!------------!-------------!--------------! ! ! 1 ! 2 ! 1 ! 2 ! ! !---------------!------------!-------------!--------------!------------------! ! ON ! * ! * ! * ! A/THR 1 ! !---------------!------------!-------------!--------------!------------------! ! OFF ! ON ! * ! * ! A/THR 2 ! !---------------!------------!-------------!--------------!------------------! ! OFF ! OFF ! ON ! * ! A/THR 1 ! !---------------!------------!-------------!--------------!------------------! ! OFF ! OFF ! OFF ! ON ! A/THR 2 ! !---------------!------------!-------------!--------------!------------------! ! OFF ! OFF ! OFF ! OFF ! A/THR 1 ! ! ! ! ! ! (OR A/THR2 IF ! ! ! ! ! ! A/THR 1 FAIL) ! !----------------------------------------------------------------------------! * either ON or OFF

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DESCRIPTION AND OPERATION Engagement The engagement of the Autothrust function can be MANUAL or AUTOMATIC. The Autothrust ( A/THR ) is engaged MANUALLY by pressing the A/THR pushbutton on the Flight Control Unit ( FCU ). This is inhibited below 100 feet RA, with engines running. The A/THR is engaged AUTOMATICALLY : - when the Autopilot/Flight Director ( AP/FD ) is engaged in TAKE-OFF or -GO AROUND modes, - or - in flight, when the Alphafloor is detected ; this is inhibited below 100 feet RA except during the 15 seconds following the lift-off. Note: To effectively have A / THR on engines, the engagement of the A / THR is confirmed by a logic of activation in the Engine Control Unit ( ECU ). A/THR Loop Principle To perform the A/THR function, the Flight Management and Guidance Computer ( FMGC ) communicates, on the one hand, with the FCU and, on the other hand, with the ECU via the FCU and the Engine Interface Units ( EIUs ).

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ACTIVE RANGE

1

1

FLX TO

EXCEPT FLX TO

>IDLE

2

Figure 31

A/THR Engagement

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Description and Operation (Cont.) Thrust Levers The thrust levers are manually operated and electrically connected to the Engine Control Units. Each lever has 4 positions, defined by DETENTS or STOPS, and 3 operating segments. Note that the thrust levers never move automatically. The Engine Control Units compute the thrust limit which depends on the position of the thrust levers. The thrust levers can be moved on a sector which includes specific positions : - ”0” : corresponds to an IDLE thrust, - ”CL” : corresponds to a CLIMB thrust, - ”FLX/MCT” : corresponds to a FLEXIBLE TAKE-OFF thrust or a MAXIMUM CONTINUOUS thrust, - ”TO/GA” : corresponds to a MAXIMUM TAKE-OFF/GO AROUND thrust. The thrust reverser levers only allow REVERSE thrust to be performed. If a thrust lever is in a detent, the thrust limit corresponds to this detent. If a thrust lever is not in a detent, the thrust limit corresponds to the next higher detent. The FMGCs select the higher of the ECU1,and ECU2 thrust limits. A/THR Function Logic The A/THR function can be ENGAGED or DISENGAGED. When it is engaged, it can be ACTIVE or NOT ACTIVE. DISENGAGEMENT case : - the thrust levers control the engines, - on the FCU, the A/THR pushbutton light is OFF, - the Flight Mode Annunciator ( FMA ) displays neither the A/THR engagement status nor the A/THR modes.

When the A/THR engage logic conditions are present, the A/THR can be engaged. It is active or not active depending on the thrust lever position. A/THR is ACTIVE if : - setting thrust levers between CL and IDLE detents ( with two engines running ) - or between MCT and IDLE detents if one engine inoperative - selecting the FCU A/THR pb on while the thrust levers are in the A/THR active range - activation of ALPHA FLOOR regardless of A/THR initial status and thrust levers position. Note : While A/THR is active: - If at least one thrust lever is set out of the CL detent anywhere within the A/THR active range, A/THR remains active. ASYM amber message is displayed on FMA. When the A/THR function is ENGAGED and ACTIVE : - the A/THR system controls the engines, - on the FCU, the A/THR pushbutton light is ON, - the FMA displays the A/THR engagement status ( in white ) and the A/THR mode. A/THR is NOT ACTIVE if ; - as soon as one thrust lever is placed outside the active range, the two engines are controlled by the position of the thrust levers. This lasts as long as the ALPHA FLOOR protection is not activated. When the A/THR function is ENGAGED and NOT ACTIVE - the thrust levers control the engines, - the A/THR pushbutton light is ON, - the FMA displays the A/THR engagement status ( in cyan ) and the A/ THR mode. Note that in case of one engine failure, the A/THR activation zone becomes between ”MCT” and ” > 0” stop.

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ACTIVE RANGE

1

1

EXCEPT FLX TO

>IDLE

2

Figure 32

FLX TO

A/THR Engagement

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Description and Operation (Cont.) Modes The A/THR function works according to modes and their related reference parameters. The reference parameter can be : - a SPEED or a MACH NUMBER; in this case, the source is either the FCU ( value chosen by the pilots ) or the FMGC itself, - a THRUST; in that case, the sources are either the ECUs ( which compute the thrust limit ) when the thrust limit is needed, or the FMGC itself. The possible Autothrust modes are SPEED, MACH, THRUST, RETARD and ALPHAFLOOR - PROTECTION. The choice of the mode is made by the FMGCs : - SPEED or MACH mode, the reference which are selected on FCU or managed by the FMGC, - THRUST mode, where the reference corresponds to a thrust limit computed by the ECUs ( according to the thrust lever position ), idle thrust in descent or optimum thrust computed by the FMGC, - RETARD mode : a thrust reduced to and maintained at idle during flare, - ALPHAFLOOR - PROTECTION : a TO/GA thrust setting to protect the aircraft against excessive angle-of-attack and windshear.

Alphafloor The A/THR function protects against an excessive angle of attack. The Alphafloor signal is detected by the FACs or ELACs. In case of excessive angle-of-attack, the FACs send an order to the FMGCs which activate the Alphafloor protection. The Alphafloor detection automatically engages and activates the A/THR function, whatever the thrust lever position and the A/THR engagement status : the engine thrust becomes equal to Take-Of f / Go Around thrust. When the A/THR is active with the Alphafloor protection active, the amber message ” A. FLOOR ” is displayed on the Flight Mode Annunciator. When the A/THR is active with the Alphafloor protection active but, with the Alphafloor detection no longer present in the FACs, the amber message ” TOGA LK ” ( LK for LOCK ) is displayed on the FMA. The Alphafloor protection can only be cancelled through the disengagement of the A/THR function, via the A/THR pushbutton or the A/THR instinctive disconnect switches.

The A/THR modes depend on the active vertical mode of the Autopilot or Flight Director. When no vertical mode is engaged, the A/THR operates in SPEED / MACH modes except : - when THRUST mode engages automatically in case of Alphafloor, - when, A/THR being in RETARD, APs and FDs disengage, the A/THR function remains in RETARD mode.

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ACTIVE RANGE

1

1

EXCEPT FLX TO

>IDLE

2

Figure 33

FLX TO

A/THR Engagement

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Description and Operation ( Cont. ) A/THR Operation in Flight The Aircraft is on GROUND and ready for take-off. Neither AP nor A/THR are engaged. The engines are controlled by the thrust levers.

Disconnection Besides the normal A/THR operation, the A/THR function is disengaged either by pilot action or in case of a system failure.

To TAKE-OFF , the pilot sets the thrust levers to the TO/GA stop or to the FLEX/MCT detent if a flexible temperature is selected on the MCDU. This engages the A/THR function ( but it is not active ). At THRUST REDUCTION ALTITUDE, a message on the Flight Mode Annunciators indicates to the pilots that they have to set the thrust levers in the CL detent. As soon as the thrust levers are in ” CL ” detent, the A/THR is active. Then, the thrust levers remain in this position until the approach phase. If only one thrust lever is set into ” CL - MCT ” area, a message on the FMAs warns the pilot to set the thrust lever to ” CL ” detent ( LVR ASYM ). The A/THR remains active. During AUTOMATIC LANDING, before touch down, an auto call out, ” RETARD ”, indicates to the pilot that he has to set the thrust levers to the ” IDLE ” stop. When he does it, the A/THR disengages.

The A/THR function can be disengaged either by pressing at least one of the two red instinctive disconnect switches on the side of thrust levers 1 and 2 or by pressing the A/THR pushbutton on the FCU. A/THR disengagement can also be due to an external system failure. When the A/THR function is active, the actual engine thrust does not necessarily correspond to the thrust lever position. Consequently, it is important to know what happens after Autothrust disconnection. - As long as a thrust lever remains in its detent, the thrust on the corresponding engine is frozen at its last value just before the disconnection. - As soon as a thrust lever is moved from the detent, or if it was not in a detent, the thrust on the corresponding engine is smoothly adapted to the thrust lever position.

This allows the automatic activation of the ground spoilers if they are in armed condition. Then, on GROUND, the pilot sets the thrust reverser levers to the REVERSE position.

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ACTIVE RANGE

1

1

EXCEPT FLX TO

>IDLE

2

Figure 34

FLX TO

A/THR Engagement

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Description and Operation ( Cont. ) Isolation of the Engines from the A/THR System A separation of the ECUs from the FMGCs after a disengagement, is done through the wired discrete that the ECUs receives directly. This disconnection can be done in two different ways: Standart disconnection; - depress the INST DISC P/B on thrust levers, or - set all thrust levers to IDLE detent. Non standart disconnection; - depress the FCU A/THR P/B while A/THR is active ( no effect in LAND TRACK ), or - loss of the arming conditions ( e.g. failure condition ). Action on one of the two INST DISC P/B forces the relays of FMGC and ECU to the separat the systems. Action of one of the two INST DISC P/B for more than 15 sec. inhibits any engagement of the A/THR function, what ever the reason ( FCU A/THR P/B switch, Alpha floor protection etc. ) Recovering is only possible at next computer power up.

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15 sec

Figure 35

Isolation of the Engines

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22-60

FLIGHT AUGMENTATION

Functions The FAC performs the functions given below : - yaw damper - rudder trim ( manual or 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. Operating principles The FAC is a dual-dual type system for yaw damper, rudder trim and rudder travel functions ( fail operational ). FAC 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 per function ( yaw damper, rudder trim, RTL ) is possible. The following functions are achieved upon energization independently of 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. Yaw Damper The yaw damper 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 enga ged in cruise only. - engine failure recovery when the autopilot is engaged ( the ELACs provide this function in manual flight ). Rudder Trim The rudder trim 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. Position of the trim is indicated on the center pedestal. Rudder Travel Limitation This function provides the limitation of the rudder travel by displacement of a stop as a function of the speed. Monitoring of flight envelope and computation of maneuvering speed This function provides the primary flight display ( PFD ) with different data displayed on the speed scale. The FAC also computes the conditions of activation of the alpha floor mode of the A/THR functions (angle of attack protection in case of windshear). BITE function of the system The FAC 1 performs BITE function of the whole AFS / FMS. Each computer includes its own BITE function and is linked to the FAC 1. The MCDU displays the content of the maintenance data via the CFDIU.

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ARTIFICIAL FEEL

YAW DAMPER ACT (YELLOW) YAW DAMPER ACT (GREEN) RUDDER POSITION TRANSDUCER UNIT

Figure 36

Rudder Components

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Figure 37

FAC Components

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Figure 38

FAC Peripheral

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26 V AC POWER SUPPLY The rotary variable differential transformers ( RVDT ) and the linear variable differential transformers ( LVDT ) associated with the FAC1 are supplied with 26V/400 Hz : - from the 115VAC ESS BUS 4XP through the 26VAC ESS BUS 431XP.A via 3A circuit breaker 14CC1. The components associated with the FAC2 are supplied with 26V / 400 Hz : - from the 115VAC BUS 2 2XP through the 26VAC BUS 2 231XP-A via 3A circuitbreaker 14CC2.

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HEATER

This CB supplied also the heater of the RTL unit

Figure 39

Power Supply - Block Diagram

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ENGAGEMENT - DESCRIPTION AND OPERATION General Each flight augmentation computer ( FAC ) includes two independent computation channels with digital processors. The engagement and monitoring principles ensure : - safe operation through the failure detectors - maximum availability through the automatic reconfigurations further to failures. System Description Each FAC comprises the following devices for function monitoring : - an engagement device per FLT CTL/FAC pushbutton switch common to the yaw damper, rudder trim and rudder travel limiting functions. - global internal monitoring of the computer in software ( real-time monitor ) and hard-wired circuitry ( FAC HEALTHY, watchdog ). - Monitoring of reconfiguration of certain peripherals - monitoring of sensors. The result of the failure is memorized in flight only. The logic circuits common to all the functions lead to the total loss of the FAC with illumination of the FAULT legend. The result of the failure is memorized. The reset will be possible only upon manual action by the pilot on the FLT CTL / FAC pushbutton switch. The Flightcontrol function can be disengaged through action on the pushbuttonswitches, but not the flight envelope protection.

Connection with FLT CTL/FAC Pushbutton Switches Each FAC is associated with an engagement pushbutton switch located on the FLT CTL panel, on the overhead panel. This pushbutton switch serves for : - The engagement or the disengagement of all the flight control func tions, engagement status : no indication on the pushbutton switch, disengagement status : the OFF legend is on. - The indication of FAC failures with the FAULT legend. This authorizes a pilot action ( FAULT/OF F ) to reset the digital section of the FAC. If the action is operative, the FAULT legend goes off and the system can be re-engaged. Therefore in normal operation the legends are off. In abnormal operation these indications are given : - computer not energized or not installed : FAULT legend on ; ECAM warning. - FAC failures specific to one function : FAULT legend off ; ECAM warning. - Common FAC failures which can be reset : FAULT legend on with possible reset by the pilot ; ECAM warning. - Power-supply transient failures : FAULT legend on with possible reset by the pilot. - FAC failures on the ground with engines shut down : FAULT legend with automatic reset at failure suppression.

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Figure 40

FAC Engagement

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Warnings - FAC Faults

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FAC 1 (2) Fault

FAC 1 and 2 Fault

CAT 3 DUAL

Figure 41

Warnings - FAC Faults

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CONFIGURATION AND OPERATIONAL SPEED COMPUTATION - DESCRIPTION AND OPERATION General The flight augmentation computer ( FAC ) fulfills several functions independently of the engagement status of the FLT CTL/FAC pushbutton switches. These functions are necessary for : - the control of the speed scale on the primary flight displays ( PFDs ). - the adaptation of gains of the flight management and guidance computer ( FMGC ) and elevator aileron computer ( ELAC ). - the distribution of signals for the FMGC control laws - the protection of the flight envelope in automatic flight ( speed limits for the FMGC, alpha-floor for the autothrust ) - the display of the rudderposition input. The FAC therefore computes : - the weight and the center of gravity - the characteristic speed data - the aerodynamic flight-path angle ( Gamma actual ) and the potential flight-path angle ( Gamma command ) - the alpha-floor protection - the position of the rudder trim for the ECAM system. - the windshear detection. - the low energy warning.

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Figure 42

Interconnection between FAC and Users

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The signification of the different speed data is given below : - VSW : stall warning speed - VALPHA PROT : speed corresponding to angle-of-attack reached when ELAC Alpha Protection is triggered. - VALPHA LIM : minimum speed which can be reached in ELAC Alpha Protection - VLS : lower selectable speed for a given configuration - VMAN (Green dot) : maneuvering speed : This speed represents the drift down speed which corresponds to the optimum speed (max. lift-to-drag ratio) in the event of engine failure. - V3 and V4 : minimum flap and slat retraction speed V3(F) = minimum flap RETRACTION speed V4(S) = minimum slat RETRACTION speed - VMAX : maximum allowable speed It determines a maximum value not to be exceeded. It represents, depending on the configuration, the smallest value of the following : VFE = maximum flap and slat extended speed VLE = maximum landing gear extended speed in clean configuration VM0/MM0 = maximum operating limit speed - VMAXOP : maximum selectable speed - VC TREND : airspeed tendency. It corresponds to the speed increment in 10s with the actual acceleration of the aircraft - VFEN : in landing phase, it corresponds to the VFE at next flap/slat position.

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B

SPD LIM FLAG

A

Appears when both FACs are inoperative. In this case, the following PFD information is lost : - VSW, VLS, S, F, Green Dot, Speed Trend, VMAX, VFE, VFE NEXT.

B

SPD LIM

FLT / CTR IN PITCH ALTN OR DIRECT LAW

FLT / CTR IN PITCH NORMAL LAW

Figure 43

PFD - SPD Scale

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DETECTION OF ALPHA FLOOR Computation of Alpha Floor Protection The alpha floor protection is calculated in the FAC. This function enables : - To protect the aircraft against excessive angle-of-attack. To do this, a comparison is made between the aircraft angle-of-attack and predetermined thresholds function of configuration. Beyond the thresholds the FAC transmits a command signal to the autothrust which will apply full thrust. - To protect the aircraft against windshear in approach by determining a wind acceleration ( deduced from the difference between ground acceleration and air acceleration ).

The ELAC direct computation of the alpha floor protection is taken into account directly as soon as the first detection is made either by the FAC or by the ELAC

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FAC 1/2 ELAC 1/2

Figure 44

Detection of Alpha Floor Condition

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WINDSHEAR DETECTION

ENERGY AWARENESS

General The windshear is a sudden change in wind direction and / or speed over a relatively short distance in the atmosphere. This can have an effect on aircraft performance during takeoff and landing phases. In windshear conditions, the principle is to reduce the detection threshold according to the detected windshear in order to get the possibility of performing a go around maneuver sooner.

General It is the generation of an aural warning from FWCs in the cockpit, telling the crew that with the current thrust, it won‘t be possible to recover flight path angle through pitch control. The only warning is. ‘ SPEED SPEED SPEED ’ ! Whenn this low energy warning appears, thrust must be increased until warning disappears, or alpha floor may be triggered. This warning is available only with flaps and slats in configuration 3 or Full and with radio altitude between 100 ft and 2000 ft. The alpha floor function inhibits the low energy warning. The low energy warning is triggert if the A/C angle of attack is greater than the computed ‘ low energy ‘angle of attack. The latter depends on A/C configuration, deceleration rate and flight path angle.

Warning The windshear warning is only active at takeoff and at landing below 1000 ft. During the landing phase the warning is inhibited at 50 ft. The crew is informed of a windshear by activation of the following warnings: - Red ” WINDSHEAR ” legend displayed on the PFDs at least 15 s - ” WINDSHEAR ” three times announcement generated by the FWC.

The computation is performed in the FACs, and relies only on ADIRU avaibility.

In case of warning inhibition in both FACs, the ” WINDSHEAR DET FAULT ” message appears on the upper ECAM display unit during TO or landing as soon as the slats are extended.

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WIND SHEAR

Figure 45

Windshear Indication on PFD and ECAM

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RUDDER TRIM - DESCRIPTION AND OPERATION Components There are two rudder trims. All components are duplicated exept the RUDDER TRIM selector and RESET button. The rudder trim function is achieved by ; an electromechanical actuator which comprises 2 asynchronous motors connected to a reduction gear by rigid linkage and 4 Rotary Variable Differential Transduceres ( RVDT ). a RUD TRIM selector for manual trim control. a RESET pushbutton. a rudder trim indicator located to the left of the RUD TRIM selector. two Flight Augmentation Computers ( FAC 1 and FAC 2 ). General The rudder trim function has two modes: Manual Mode, when the autopilot is not engaged and Automatic Mode when the autopilot is engaged. The autotrim order is computed by the laws, whereas the manual trim order transits through them. The order is then sent to the actuator. This order is reproduced at the rudder pedals. Priority is given to the rudder trim of FAC 1; a changeover logic enables to the switch to FAC 2 in case of failure. If both rudder trims fail, the last deflection is maintained The rudder position is displayed on the RUD TRIM indicator and on the ECAM display unit.

Power Loop During the autotest triggered by the FAC power up the internal actuator monitoring checks the actuator servo-loop and monitor circuit validity, and the enabling signal reception. Then the changeover logic enables the trim motor to be supplied and the rudder trim laws control it. The laws compute the trim order and sent it to the actuator’s motor via the Electronic Control Circuit. The feedback in the power loop is provided by two Rotary Variable Differential Transducers ( RVDT ) for each side. Monitoring The computation and the power loop are monitored by comparators. The input parameters are also monitored. The computation is monitored by the comparators between the FAC Command and Monitor parts. The FMGC and ADIRS peripheral inputs are always monitored. The power loop is monitored by the the comparators between the rudder trim order and the position feedback signal.

Manual Mode When the autopilot is not engaged, the rudder trim order is given by the RUDDER TRIM selector. Note : The RESET pb. enables to return the rudder to the neutral position. Automatic Mode With the autopilot engaged, the Flight Augmentation Computer calculates the trim order using Flight Management and Guidance Computer and Air Data Reference System data. Note : At touch-down, the AUTO RESET function moves the rudder to the neutral position.

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Figure 46

Rudder Trim Schematic

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Warnings - Rudder Trim faults

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RUD TRIM 1 AND 2 FAULT

RUD TRIM 1 (2) FAULT

AP 1+2

Figure 47

Warnings - Rudder Trim Faults

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YAW DAMPER - DESCRIPTION AND OPERATION Components There are two yaw dampers. In normal operation both are engaged, but only one is active. Yaw damper 1 has priority. The yaw damper function is achieved by : 2 electrohydraulic actuators with an external centring element. Each actuator comprise 1 jack, 1 Linear Variable Differential Transducer ( LVDT ), 2 Electro Valves ( EV ), 2 Bypass Valves, 1 Pressure Switch ( PS ), 1 Servo Valve ( SV ). 2 Flight Augmentation Computers ( FAC 1 and FAC 2 ) 2 Rotary Variable Differential Transducer ( RVDT ). General Yaw Damper one and two operate with the changeover logics.The yaw damper actuators does not move the rudder pedals. The Yaw Damper function operates as follows : Order is computed by the laws and sent to the rudder via the related yaw damper actuator. YD actuator 1 is powered by the green hydraulic system. YD actuator 2 is powered by the yellow hydraulic system. Manual Mode In manual mode, the autopilot is not engaged and the Elevator Aileron Computer sends the turn coordination, and the dutch roll damping yaw orders to the FAC.

Land Mode When the land mode is engaged, the yaw order is computed directly by the FMGC. Power Loop The yaw damper laws control the servovalve and the changeover logic enable to pressurize the jack. The feedback in the power loop is provided by a Linear Variable Differential Transducer ( LVDT ) for the Command side and a Rotary Variable Differential Transducer ( RVDT ) for the Monitor side. In case of dual monitor loss, a centring spring rod moves the rudder to the neutral position. Monitoring At power up, the yaw damper function safety tests are initiated. The continuity between the standby yaw damper and its servo valve is tested. The computation is monitored by the comparators between Command and Monitor part. The ELAC, FMGC and ADIRS peripheral inputs are always monitored. The power loop is monitored by a comparator between the yaw order and the rudder position feedback. In flight, the hydraulic pressures are monitored by the FAC. The LVDT‘s and the RVDT‘s are always monitored.

Manual Alternate After a dual Elevator Aileron Computer failure, turn coordination is lost and a simplified alternate law of dutch roll damping is computed by the FAC. Auto Mode In auto mode, the FAC computes the dutch roll damping in clean configuration, the engine failure recovery in take-off, go-around and runway modes. The turn coordination law is computed by using roll orders from the FMGC.

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Turn Coordin. , Yaw damp. Yaw damping Align, Rollout

Figure 48

Yaw Damper Schematic

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YAW DAMPER DESCRIPTION AND OPERATION (CONT.) Pressure Switch Function The table shows the reaction of the Pressure Switch depending of the engagement-state of the electro valve #1 and #2. You see also the actuator modes.

EV 1

EV 2

0

0

0

1

1

0

1

1

ACT

PR.SW

0 BY PASS

1 Y/D FAIL

1 ACTIVE

0

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YAW DAMPER 1 (2) FAULT

YAW DAMPER 1 AND 2 FAULT

Figure 49

Warnings - Yaw Damper Faults

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RUDDER TRAVEL LIMITING DESCRIPTION AND OPERATION Components The Rudder Travel Limiting function is dual. All the components are duplicated. The Rudder Travel Limiting function ( RTL ) is achieved by ; - an electromechanical rudder travel limitation unit with two motors. - two Rotary Variable Differential Transducer ( RVDT ) integrated in the unit. - two Flight Augmentation Computers ( FAC 1 and FAC 2 ). General The Rudder Travel Limiting function acts through a control law, which is a function of the corrected airspeed, and returns to the low speed limitation in case of failure. Normal operation :

The return to low speed logic connects the motor for 30 sec directly to 26 VAC in order to recover full rudder deflection. Monitoring The computation and the power loops are monitored by comparators. The computation is monitored by the comparator between the FAC Command and Monitor channels. The ADIRS parameters ( Vcas ) are monitored by a two-by-two comparison and then one of them is selected. The power loop is monitored by the comparators between the Rudder Travel Limiting order and the RTL unit position feedback.

Rudder Travel Limation Unit

The RTL law in the command channel of the FAC 1 ( active side ) controls the limitation unit stops through a motor. Return to low speed : If both Rudder Travel Limiting function fail when the slats are extended, the full rudder deflection is obtained. Priority is given to RTL of FAC 1; a changeover logic enables to switch to FAC 2 in case of failure. Laws The Rudder Travel Limiting control law generates a rudder deflection order in relation to the corrected airspeed. Power Loop The Rudder Travel Limiting law controls the unit’s motor, and the changeover logic enables the motor to be supplied. The return to low speed function has an independent power supply. The law computes the RTL order and sends it to its motor via an electronic control. The feedback in the power loop is provided by one Rotary Variable Differential Transducer ( RVDT ) for slaving and monitoring.

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Figure 50

Rudder Travel Limiting Schematic

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Warnings - RTL Faults

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Rudder Travel Limit 1 and 2 Fault

Rudder Travel Limit 1 (2) Fault

CAT 3 Single only

Figure 51

CAT 3 DUAL

Warnings - Rudder Travel Limitation Faults

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22-70

FLIGHT MANAGEMENT

Purpose The Flight Management part has several functions linked to the flight plan such as lateral and vertical guidance, or displays.The FLIGHT MANAGEMENT function of each FMGC, in association with the FCU and two MCDUs, performs ; - aircraft position computation, - radio nav tuning, - flight planning, - lateral navigation and flight phase management, - speed management, - performance data, - displays of flight management data ( on MCDU, ND and PFD ). Flight Plan The flight plan is defined by various elements which indicate the routes the aircraft must follow with the limitations along these routes. The elements are mainly taken from the data bases or directly entered by the pilot. The limitations are mainly speed, altitude or time constraints originated by the Air Traffic Control (ATC). The function that integrates these elements and limitations to construct a flight plan is called FLIGHT PLANNING. In addition to this, the FM part provides the aircraft position and the follow-up of the flight plan, this is called NAVIGATION. Everything can be prepared prior to the take-off but can also be modified quickly and easily during the flight operation. In case of a FM problem, the remaining valid FMGC can be used as sole source to command both MCDUs and NDs ( single mode ).

This data base is updated every 28 days. Besides this, some room is kept to allow manual entry of 20 navaids, 20 waypoints, 3 routes and 10 runways. The data base cannot be erased, however, the manually entered data can be erased. Two cycle data bases can be inserted, the selection is made automatically using data from the aircraft clock. Navigation The navigation process provides the system with current aircraft state information consisting of present position, altitude, winds, true airspeed and ground speed. This is achieved using inputs from the inertial reference system, air data sensors and navigation radios ( Global Positioning System (GPS) can also be used if it is installed ).

Navigation Data Base The navigation data base provides all necessary information for flight plan construction and follow-up. The pilot will either select an already assembled flight plan ( company route CO ROUTE- ), or will build his own flight plan, using the existing data base contents.

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FMGC 1

Figure 52

FM - Schematic

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FM - DESCRIPTION AND OPERATION (CONT.) Lateral Flight Plan The lateral flight plan provides the sequential track changes at each waypoint within 3 main sections. Departure : initial FIX (origin airport), SID... En Route : waypoints, navigation aids... Arrival : STAR, approach, missed approach, go around... The lateral steering order can be followed by the pilot or the autopilot through the NAV mode selected on the Flight Control Unit (FCU).

Vertical Flight Plan The vertical flight plan provides an accurate flight path prediction which requires a precise knowledge of current and forecast wind, temperature and the lateral flight path to be flown. The vertical flight plan is divided into several flight phases : PREFLIGHT : fuel / weight / V2 insertions. TAKE-OFF : speed management, thrust reduction altitude, acceleration altitude. CLIMB : speed limit, speed management. CRUISE : top of climb ( T / C ), cruise altitude, top of descent ( T / D ). APPROACH / MISSED APPROACH / GO AROUND : thrust / acceleration altitudes. The vertical steering order can be followed by the pilot or the autopilot. Any level change in the vertical profile is initiated after a push action on a level change selector.

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Performance The performance data base contains optimal speed schedules for the expected range of operating conditions.

Display According to the pilot selection on the FCU, the flight plan is shown in relation to the aircraft position on the ROSE-NAV or ARC modes.

Several performance modes are available to the operator with the primary one being the ECONOMY mode. The ECON mode can be tailored to meet specific airline requirements using a selectable COST INDEX ( CI ). A Cost Index is defined as the ratio of cost of time ( $ / h ) to the cost of fuel ( cts / pds ). The speed and the thrust values associated with a given Cost Index are used to determine the climb and descent profiles. FUEL and TIME are the main ”actors” in this particular part of the FM function and direct the airline choice.

The aircraft model is fixed and the chart moves. The difference between the two modes is that the half range is available when the Navigation Display ( ND ) is set to ROSE mode as there is only frontal view when it is set to ARC mode. In PLAN mode, the flight plan is shown, with NORTH at the top of the screen, centered on the TO waypoint. Depending on the selected range, the aircraft may or may not be visualized on this display. The PLAN display can be decentered by scrolling the flight plan on the MCDU. The Primary Flight Display ( PFD ) shows the FM guidance following engagement of the AP / FD lateral and longitudinal modes.

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FLIGHT MANAGEMENT PRIORITY LOGIC FM OPERATING MODES There are three operating modes : DUAL, INDEPENDENT, SINGLE. At FM initialization, that means at power up, both FM parts exchange information. Initial cross-comparison is made on the following parameters : Nav data base Perf. data base FM operational program software part numbers A/C and engine type Pin program. If the Flight Management ( FM ) parts agree, DUAL mode is active. When keys are pressed, they are immediately processed by both FMs, regardless of the MCDU from which they originate. If the FM parts disagree, INDEPENDENT mode is automaticly active. Each FM part manages its own Multipurpose Control and Display Unit. If one FM part has failed, SINGLE mode is active. Both MCDUs are driven by the remaining FM part. An independent configuration results in the MCDU - messages ” INDEPENDENT OPERATION ” . MODE OPERATION In DUAL mode, the FM part receives the master / slave activation from the Flight Guidance part. The Master computer imposes the following parameters upon the Slave computer : Flight phase Flight plan sequencing Active performance mode and speeds Clearence and maximum altitudes ILS frequencies and courses, if any. After a flight plan change, there is a comparison on the active leg and, every second, on the active performance mode and active guidance mode.

If it is different, the slave computer will synchronize itself to the master one by copying the master values. In DUAL mode operation three parameters are computed independently by each FMGC. This parameters, aircraft position, gross weight and target speeds from master and slave computers are compared every second. If the difference is greater than 5 Nm, 2 tons or 2 Kts respectively, an appropriate message is displayed on the MCDUs : - FMS1 / FMS2 POS DIFF - FMS1 / FMS2 GW DIFF - FMS1 / FMS2 SPD TGT DIFF Note :In some dynamic conditions, vertical and lateral computations may temporarily disagree and may be evident on the ND. In this case, the flight director and autoflight system use the master FMGC for tracking. In INDEPENDENT mode, there is no interaction from one system to the other one. The FMGCs only send their status information to each other (e.g. in this case, the INDEPENDENT mode). In SINGLE mode, both MCDUs are driven by the same FM part, but they can still display different pages. Messages linked to the navigation process are displayed on both MCDUs.

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Figure 53

FMGC System Architecture

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FLIGHT MANAGEMENT PRIORITY LOGIC (CONT.) MCDU As already presented in the FM OPERATING MODES topic, the MCDUs work differently. In Normal mode, the MCDUs can be used simultaneously on different pages. Any modification or entry on one MCDU is transmitted to the other MCDU via the FMGC crosstalk. In INDEPENDENT mode, both MCDU‘s operate separatly. The message ” INDEPENDENT OPERATION ” in the scratchpat indicates this operation. In SINGLE mode, both MCDUs basically work as in normal mode, but with the only valid FMGC. This mode is indicated with the message ” OPP FMGC in PROCESS ” on the corresponding MCDU. Displays Flight Management information is displayed on Navigation Displays and on Primary Flight Displays. For FM information, in DUAL or INDEPENDENT modes, FMGC1 supplies PFD1 and ND1, FMGC2 supplies PFD2 and ND2. In SINGLE mode, the remaining FMGC supplies all the displays.

Radio Navigation The schematic shows the architecture of the radio navigation receivers controlled by the FMGCs in DUAL or INDEPENDENT modes. For the selection of radio navigation frequencies and courses, in DUAL or INDEPENDENT modes, each FMGC controls its own side receivers through a Radio Management Panel ( RMP ). Only the actual frequencies and courses from the receivers are displayed on the PFDs and the NDs. In case of a FMGC failure, the valid FMGC controls its own side receivers as usual, through a Radio Management Panel, but also the other side receivers, directly without going through a RMP. If both FMGCs fail, the crew must use the Radio Management Panels to select the frequencies and courses.

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Figure 54

FMGC System Architecture

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MULTIPURPOSE CONTROL AND DISPLAY UNIT Purpose The Multipurpose Control and Display Units ( MCDU ) provide access to the following : -

FMGC ( Flight Management function ) DATA LINK ( ACARS )-optional CFDS ( Centralized Fault and Display System ) AIDS-optional.

They are composed by a keyboard and a screen for entry / display between the pilot or the line maintenance. The 2 MCDUs are located on the center pedestal.

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MCDU 1

FMGC 1

MCDU 2

AIDS

ACARS

Figure 55

CFDS

FMGC 2

MCDU Architecture

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Front Panel Annunciators There are three annunciator lights on the MCDU front panel. The ” FAIL” annunciator comes on amber when the MCDU has failed. The ” FMGC” annunciator comes on white when the FM is not the active system and it has sent an important message to display. In this case, any page key can be pressed to return to the Flight Manage ment related display. Important messages are those displayed in amber. The ” MCDU MENU” annunciator comes on white when a system, linked to the MCDU, other than the FM, requests the display.

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Figure 56

MCDU Annunciator

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EFIS FM - DISPLAY FM Display on PFD The Primary Flight Display ( PFD ), as main guidance instrument, displays the data computed or inserted on the Multipurpose Control and Display Unit. This data can be ECON speed targets and target altitudes in managed guidance modes, V1 and V2, Decision Height ( DH ) or Minimum Descent Altitude ( MDA ) in approach. At the top of the Primary Flight Display, the Flight Mode Annunciator ( FMA ) provides the pilot with the DH or the MDA. The speed scale displays the Flight Management data such as the speed target and V1. The altitude scale displays the altitude constraint from the Flight Management ( FM ) part and the linear vertical deviation with respect to the FM theoretical vertical flight plan ( F-PLN ). Landing elevation is also indicated by a blue horizontal bar on the altitude scale.

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Figure 57

FM Display on PFD

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EFIS DISPLAY (CONT.) FM Display on ND The Navigation Display works in six different modes selected on each Electronic Flight Instrument System ( EFIS ) control panel of the Flight Control Unit ( FCU ). In ROSE-NAV, ARC and PLAN modes, the Navigation Display ( ND ) displays the flight plan computed in the FM part at a scale defined by the range selected on the FCU. The ND represents basically : the aircraft position, the flight plan data, the range selected on the FCU, autotuned navaids. Note : The aircraft position is fixed in all display modes except in PLAN mode, where it moves along the flight plan. There is correspondance between the flight plan displayed on the ND and the MCDU FLIGHT PLAN page if no scrolling has been done on this page.

Note : Wind speed and direction, ground speed and track are computed by the FM part and transmitted to the Display Management Computers ( DMC ) which also receive the same data from the Air Data and Inertial Reference Units (ADIRU). In accordance with the pin programming, the DMC selects the ADIRU data to be displayed on the ND. Radio navaids are displayed in cyan when they are autotuned by the FM part. Specific symbols can appear, along the flight plan, corresponding to some maneuvers such as Start of Climb ( S/C ) in white, Top of Climb ( T/C ) in cyan, Top of Descent ( T/D ) in white, holding pattern and turn procedure.

The TO waypoint characteristics are displayed in the top right hand corner of the Navigation Display ( ND ) : ident ( in white ) and bearing ( in green ), distance to go ( in green ), Estimated Time of Arrival ( ETA ), ( in green ). The rest of the flight plan line and waypoints is displayed in green. A crosstrack deviation, if any, is also provided, in green, on the left or right hand side in nautical miles.

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ILS

Figure 58

FM Display on ND

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EFIS DISPLAY (CONT.) Data Base Display P/B By pressing these five interlocked pushbuttons : WPT, VOR.D, NDB, ARPT and CSTR, different information from the Navigation data base is available in magenta. Note that these options are exclusive and the priority is given to the last which has been selected. When the WPT pushbutton is pressed in, all waypoint locations in the related range, are transmitted to the ND to be displayed. When the VOR.D pushbutton is pressed in, all VOR and / or DME stations locations in the related range, are displayed on the ND. When the NDB pushbutton is pressed in, all Non Directional Beacon station locations in the related range, are transmitted to the ND to be displayed. When the ARPT pushbutton is pressed in, all airport locations in the related range, are transmitted to the ND to be displayed. When the CSTR pushbutton is pressed in, all speed, altitude and time constraints ( if any ) on one or several waypoints, are transmitted to the ND to be displayed. For example, the constraints on BGN waypoint are : - the flight level is constrained to below ( - ) FL180, and the - speed is constrained to below ( - ) 250 kts.

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ILS

Figure 59

CSTR P/B pressed in

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22-90

FAULT ISOLATION

BITE AND FIDS DESCRIPTION AND OPERATION General The Auto Flight System is a type 1 system, able to maintain a two-way communication with the Centralized Fault Display System. It comprises a system BITE located in FAC 1 called Fault Isolation Detection System ( FIDS ). Basically, the faults detected by the computer BITEs are concentrated in the system BITE called Fault Isolation Detection System ( FIDS ), and can be accessed through the MCDU and the CFDS. Like for other systems, the CFDIU works in NORMAL MODE and MENU MODE ( see ATA 31 - CFDS ). A FIDS card is fitted in each FAC. Both FACs are interchangeable, but only the FAC 1 FIDS is active due to side 1 signal. Note : When the FIDS has failed, BITE‘s continue to work, the results can be read in the shop after FAC 1 change.

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FMGC 1 FAC 1

CFDIU

FMGC 2 FIDS FAC 1

FAC 2

FCU 2 MCDU 1

MCDU 2 MCDU 1

AFS Computer BITEs

Figure 60

AFS BITE Architecture

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BITE AND FIDS DESCRIPTION AND OPERATION (CONT.) MCDU BITE The MCDU performs tests on its processor, memory and display unit; If a failure is found by the MCDU BITE. - the ” FAIL” annunciator comes on in amber and the display is blank - no snapshot is taken - the MCDU FAIL output discrete is set and sent to FG 1 and FG 2 parts. FCU BITE Each FCU BITE computes the maintenance status of its related part and permanently sends this maintenance data to the FG part.

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D+I+M MON

D+I+M FM

D: DETECTION I: ISOLATION M: MEMORIZATION

D+I+M CMD

T O

FMGC 1 F I D S

FMGC 2 F

D+I+M CMD

A I L

D+I+M MON

D+I+M FM

Figure 61

MCDU BITE

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22-91

TESTS

Operation Normal Mode: The use of the system in normal mode is described in ATA REF 31-32-00. Menu Mode: Access to the main menu of the FIDS : chaining of the operations enabling display of this menu is described in the next figure.

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Figure 62

Menu Chaining

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Menu Mode The menu mode is relevant to a specific operation enabled only on the ground. It is based on an interactive dialogue betweenthe FIDS and the MCDU. The functions of the system in menu mode are described in the next figure:

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Digit 4

Figure 63

AFS Main Menu

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GROUND SCAN This function is accessible from the MCDU when the system is in menu mode The three following functions can be accessed when the AFS / GROUND SCAN page is displayed: - GROUND REPORT - PRESENT FAILURES SCAN - PROGRAM GROUND REPORT function This function enables the failures recorded in the ground area of the FIDS memory to be displayed. The content of the ground area is erased during computer power up and engine start . The failures memorized and visible in the GROUND REPORT are the ones which occurred after the last ground area initialization. Two types of content can be displayed:

PRESENT FAILURES SCAN function (GROUND SCANNING) This function is used to isolate failures present when the function is selected. Therefore an inhibited failure will not be announced. Once the function is activated ( push action on the line key adjacent to the PRESENT FAILURES SCAN indication ), a wait message is displayed for 40s while the system isolates the present failures. After this time, the messages are displayed on the GROUND REPORT page. NOTE : As soon as the PRESENT FAILURES SCAN function is selected, the ground contexts previously recorded are erased and thus definitely lost.

Only the internal failures that occured on ground are normally displayed by the GROUND REPORT function. After selection of the PRESENT FAILURES SCAN function ( Ref. para. PRESENT FAILURE SCAN ) all internal and external failures ( considering a limit of three contexts ) found during this operation are seen in this report. As selection of the PRESENT FAILURES SCAN function erases the content of the ground area, it is highly recommended, prior to this selection, to display this content using the GROUND REPORT function. Failures are presented with the following data: - the flight counter ( Leg - 00 ) which indicates that the failure occurred on the ground. - the ATA reference ( AMM PGBK. 400s ) and the associated Failure message. - the computer which identified the failure. Additional information can be obtained by selecting the TROUBLE SHOOTING DATA item.

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Digit 4

Figure 64

Ground Scan

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AFS TEST The purpose of the AFS TEST is to check the integrity of the AFS after replacement of an LRU ( line replaceable unit ). The AFS TEST completes the AFS computer monitoring and safety tests. This test, which is performed in the FACs and the FMGCs ( FM and FG sections ) consists in: using the computer safety test results ( FAC, FG, FM, FCU and MCDUs ) the test of symmetrical discrete inputs : FAC COM and FAC MON, FG COM and MON the test of the symmetrical ARINC inputs the plausibility test of the information delivered by: - the RUD TRIM/RESET pushbutton switch - the rudder trim control switch - Capt A/THR instinctive disconnect pushbutton switch - F/O A/THR instinctive disconnect pushbutton switch - Capt takeover and priority pushbutton switch - F/O takeover and priority pushbutton switch - FAC engagement pushbutton switch.

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Figure 65

AFS Test

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LAND CAT 3 CAPABILYTY TEST The purpose of the test is to verify the capability of the involved systems to perform a CAT 3 fail - operational automatic landing. It also verifies the takeover and priority pushbutton switches, the A/THR instinctive disconnect pushbutton switches and the warnings associated to the automatic landing. The LAND TEST function is mainly performed in the FIDS and utilizes FG failure detection ( snapshot, analysis and reporting ). Consequently, the LAND TEST efficiency is identical to the FG BITE efficiency. Test Principle This test consists in checking the correct operationof the systems inside and outside the AFS and involved in CAT 3 automatic landing ( correct operation of BITE‘s, system reception, self - test results, interconnections validity ).

Test Running If a failure occurs prior to the acceptation phase, the test is refused. If a failure occurs after the acceptation phase, the FMGCs remain in LAND TEST condition. From AFS / LAND TEST-4 page, the operator must answer questions by YES or NO via the MCDU. NOTE:Please answer by YES if agree with sentence, NO if disagree. If the answer is YES, the test continues until the last page is displayed ( AFS / LAND TEST-9 ) with TEST OK final message. If the answer is NO, an analysis is made at the level of the AFS BITEs in order to detect and isolate the failure. A failure message is displayed on the AFS / LAND TEST REPORT page requesting to check the system concerned by the analysis. NOTE:Each AFS / LAND TEST page displays an END OF TEST indication. Pressing the line key adjacent to this indication results in the transmis sion of an END OF TEST FIDS command to the four FG BITEs. Reception of this command causes loss of the LAND TEST ACCEPTA TION condition for each BITE. Chaining of the various pages of the Land Test are described in the next figure.

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1

Figure 66

Land Test ( Accepted )

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Figure 67

Land Test ( Refused )

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ATA 31

INDICATING AND RECORDING

31-51

FLIGHT WARNING SYSTEM

ALTITUDE ALERT Operation An altitude warning ( ” C ” chord sound and altitude window of PFD pulsing yellow or flashing amber ) is generated by FWC when A/C approaches a preselected altitude or flight level. This warning is based on comparison of altitude ( ADIRS ) with preselected altitude displayed on FCU. Continuous ” C ” chord is cancelled by a new altitude selection or the EMER CANC pushbutton of the ECAM control panel or the MASTER WARN pushbutton. The altitude box frame flashing is extinguished by a new altitude selection. The altitude alert is inhibited when the slat are out with L/G selected down or in approach after capture of glide slope or when L/G is downlocked.

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Figure 68

Altitude Alert

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ATA 34

NAVIGATION

34-34

PARAVISUAL INDICATING ( PVI )

PVI DESCRIPTION AND OPERATION Principle The PVI is a roll out piloting aid system when on runway, during take-off and landing phases below 30 feet, in reduced visibility conditions. The PVI is installed on the Captain side of the glareshield and generates a head up paravisual image for the Captain. The electronic and display modules are integrated in a single unit.The autoland indicator is also integrated in this unit. The Liquid Cristal Display consists of two fixed stripes and moving box. The Captain can correct the aircraft trajectory following the PVI indication. Components The PVI is provided for the Captain only. The electronic and display modules are integrated in a single unit. The autoland indicator is also integrated in this unit. ON/OFF Switch An ON/OFF switch controls the PVI. When it is set to OFF the PVI display is black. PVI in Standby When the PVI is ON and guidance command presentation conditions are not fullfilled, fixed marks are displayed. The conditions of display are : PVI ON no PVI internal failure correct reception of Yaw Flight Director control AP/FD modes and engagement such as RUNWAY LOC mode or ROLL OUT mode.

PVI in Guidance Command Presentation The PVI displays a moving symbol and a fixed mark. When the moving symbol is on the right, the Captain has to correct on the right. Opposite for the left. When the moving symbol is centered, no action is required. External Failure When a non valid signal input is detected, the PVI display is white. Internal Failure When the PVI integrated test detects a failure, a black display appears. Test Condition The test can be performed when the PVI is not in guidance command presentation. The test can be performed when : PVI is ON not in RUNWAY LOC mode not in ROLL OUT mode. Test without Failure To perform the test the PVI has to be switched on. The test P/B has to be pressed. The AUTOLAND and the PVI display are tested. The autoland light comes on red. The moving symbol moves from the center to the right, then to the left, and so on until the test pushbutton is released. Test with Failure To perform the test the PVI has to be switched on. When a failure is detected by the power up test, the PVI display is black.

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Figure 69

PVI -Schematic

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Test A self test controlled by a pushbutton is integrated in the PVI. When the test pushbutton is pressed in, a discrete signal is transmitted to each Flight Warning Computer. The autoland lights comes on red. Note: If the ILS switches on the EFIS control panel are ON, the LOC / GS scales and indexes are flashing too.

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Figure 70

Autoland Light

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ATA 22

AUTOFLIGHT

APPENDIX TAKE OFF SEQUENCE Take Off Mode This mode provides lateral guidance function, at takeoff, on the runway centerline by means of the LOC beam and by following an optimum longitudinal flight path after rotation. The mode is engaged when the pilot selects the takeoff thrust by positioning the throttle control levers beyond the MCT / FLX TO gate. Engagement of the mode is shown by the green SRS ( Speed Reference System ) and RWY indications in the Flight Mode Annunciator ( FMA ) columns corresponding to the longitudinal modes. The pitch guidance law enables holding of V2 + 10 kts in normal engine configuration. Prior to mode engagement, the pilot must select speed V2 on the TAKE OFF page of the MCDU or on the FCU if the FM is faulty.

NOTE : In engine fail detection, the law enables to hold: The aircraft speed ( Va ) if it is greater than V2 when the engine failure occurs, or V2 if the aircraft speed ( Va ) is lower than V2 when the engine failure occurs. In addition, the guidance law includes: An attitude protection to reduce the A / C nose-up attitude during this phase. A flight path angle protection to ensure a minimum climbing rate.

With V2 selected, the managed speed control is activated and the TO longitudinal mode ( SRS ) can be engaged. Without V2 selection on the MCDU, the mode is not engaged on this axis. The guidance law on the lateral axis provides guidance of the aircraft on the runway centerline by means of the LOC beam. For this, the FM or the pilot selects the ILS frequency associated with the takeoff runway. This selection can be made: Implicitly by selecting the takeoff runway or departure procedure on the MCDU. Expressly by selecting the frequency on the RMP or the MCDU. The laterale TO ( RWY ) mode can be engaged when the aircraft is at the end of the runway and receives the LOC deviation signals. If the ILS is not available or if the ILS frequency is not selected, theTO mode is not engaged on this axis.

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E

Figure 71

Take Off Sequence - System Preparation

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NAV

1FD2

1FD2

CLB

CLB

NAV 1FD2

NAV

NAV

1FD2

1FD2

1FD2

RWY

NAV

1FD2

1FD2

Figure 72

Take Off Sequence - with NAV Armed

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APPROACH AND LANDING LOC Mode Localizer capture shall be achieved with only one overshoot followed by a constand convergent heading ( if needed ) under the following conditions: A track angle error of between 20 and 60. Capture initiated at a distance of at least 10 nautical miles from the runway threshold. Aircraft ground speed of 200 kts. LOC beam sensitivity of 75 mA per degree. Still in air the LOC beam shall be tracked to within 7.5 mA. GLIDE The overshoot on glideslope beam capture shall not exeed 75 mA. If the capture is initiated when the aircraft is on, or above the beam center line, the overshoot shall not exeed 150 mA providing that the capture altitude is above 1500 ft. Still in air the glideslope beam shall be tracked to within 20 mA.

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Figure 73

Appr. and Ldg. Sequence - System Preparation

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NAV

NAV

Figure 74

Approach and Landing Sequence - ILS Approach

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GO AROUND On the laterale axis, the engaged mode enables to hold the track followed by the aircraft. On the longitudenal axcis, it ensures managed speed control.The speed reference of the guidance law is the aircraft speed when the mode was engaged ( the lower limit of speed is the approach speed ). This mode is available on the AP and the FD. It is engaged when the pilot selects the maximum thrust by positioning the throttle control levers in the TO / GA gate. Engagement is indicated by the green ”SRS” and ”GA TRK” indications displayed in the FMA sections corresponding to the longitudenal and lateral modes.

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Figure 75

GO Around Sequence

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FOR TRAINING PURPOSE ONLY AVIONIC TRAINING CENTER ______________________________________________________________________________________________________________________________________

Training Manual A 319/320/321 ATA 22 Autoflight TROUBLE SHOOTING EXERCISES

Lufthansa Issue: MARCH 2000 Technical Training GmbH For Training Purposes Only Lufthansa Base Lufthansa 1995 Book No: ______________________________________________________________________________________________________________________________________



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ATA 22

T/S EXERCISES

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Revision No : 02 FRA US/T hz APR 98 Issue Date : 21/05/2013


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Familiarize yourself with the different documentations and references when trouble shooting. Use the following exercises refered on a TLB-report and PFR and answer the question. You can check your answer given on page 4.

Exercise 1 TLB: - A/P disengagement in Cruise Flight - ECAM: AFS: AP OFF PFR: ECAM WARNING MESSAGE: - 22-oo AUTO FLT AP OFF, FAILURE MESSAGE: - 22-83-34 AFS : FMGC 1 / C-M ARINC LINK SOURCE: AFS IDENTIFIERS: ECAM 1,ECAM 2, EIS 1, EIS 2, EIS 3 Question 1: Which TSM Task do you have to use ? Question 2: Which test do you have to perform for Fault Confirmation ? Question 3: After removal and installation of the FMGC the fault is still present. You have to check the Wiring. What is the FIN of the ” first terminal block ”? Question 4: Which test do you have to perform at the end of your trouble shooting ?

Exercise 2 TLB: - MCDU 1 FAIL- Light comes on. PFR: ECAM WARNING MESSAGE: - none FAILURE MESSAGE: - 22-82-12 AFS : MCDU 1 SOURCE: AFS IDENTIFIERS: CFDS Question 1: Which Fault Symptom Index do you have to use to find this malfunction ? Question 2: Which TSM Task do you have to use ? Question 3: Which test do you have to perform for Fault Confirmation ? Question 4: After removal and installation of the MCDU which test do you perform to check the proper installation ?

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Exercise 3 TLB: - AF / FCU 1 FAULT PFR: ECAM WARNING MESSAGE: - 22-00 AUTO FLT FCU 1 FAULT FAILURE MESSAGE: - 22-81-12 AFS : FCU SOURCE: AFS IDENT : ADR 1, ADR 3, ECAM 1, EIS 1, EIS 3, RADAR 1 Question 1: Which TSM Task do you have to use ? Question 2: The Aircraft is CAT III certyfied. You have to change the FCU. After removal and installation of the FCU which test do you perform to check the proper installation ? Question 3: Do you have to perform a LAND 3 TEST ? Question 4: Which task do you have to perform for the LAND 3 TEST ? Question 5: Which test do you have to perform at the end of your trouble shooting ?

Exercise 4 TLB: - During CRUISE-FLIGHT A/P 2 disengaged, reengage. possible PFR: ECAM WARNING MESSAGE: - 22-00 AUTO FLT YAW DAMPER 1 FAILURE MESSAGE: - 29-32-12 AFS : HYD G 1151GN SOURCE: AFS

Question 1: Which test do you have to perform for Fault Confirmation ? Question 2: TS-DATA Word 3 is: 39AC. Do you have to remove and install a new FAC ? Question 3: Which test do you have to perform at the end of your trouble shooting ?

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ANSWERS : Exercise 1: 1/1: 221100 T 810 821 1/2: TASK 22-96-00-710-001 Operational Test of the AFS 1/3: 1851 VT 1/4: TASK 22-96-00-710-001 Operational Test of the AFS

Exercise 2: 2/1: 22-LOCAL 2/2: 228200 T 810 802 2/3: TASK 22-91-00-710-001 Operational test of the ground scanning 2/4: TASK 22-70-00-710-001 MCDU Operational Test

Exercise 3: 3/1: 228100 T 810 805 3/2: TASK 22-96-00-710-001 Operational Test of the AFS 3/3: YES 3/4: TASK 22-97-00-710-001 Land 3 Capability Test. 3/5: TASK 22-91-00-710-001 Operational test of the ground scanning

Exercise 4: 4/1: TASK 22-91-00-710-001 Operational test of the ground scanning 4/2: NO 4/3: TASK 22-91-00-710-001 Operational test of the ground scanning

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Training Manual A 319/320/321 ATA 23 Communications ATA Spec. 104 Level 3

Lufthansa Issue: January 1999 Technical Training GmbH For Training Purposes Only Book No: A320 23 L3 Lufthansa Base Lufthansa 1995 ______________________________________________________________________________________________________________________________________________________________________________________________



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23-51

AUDIO MANAGEMENT

AUDIO INTEGRATING SYSTEM PRESENTATION The Audio Management Unit (AMU) is the heart of the Audio Integrating System. The AMU acts as an interface between the users and the various radio communication and navigation systems. The AMU provides the following functions : radio transmission, radio and navigation reception visual and aural warnings of the ground crew and the Cabin Attendant calls, flight interphone, interface with the Cockpit Voice Recorder (CVR) SELCAL calls, emergency function for the Captain and the First Officer.

CALLS Ground crew and cabin Attendants calls are visualized on the Audio Control Panels (ACPs).

TRANSMISSION For transmission, the AMU collects the microphone inputs from the various acoustic equipment and directs them to the radio communication transceivers selected on the Audio Control Panels (ACPs). RECEPTION For reception, the AMU collects the audio outputs from the various communication and navigation systems and directs them to the various crew stations and acoustic equipment, whatever the election made on the ACPs. FLIGHT INTERPHONE The flight interphone allows telephone links between the various crew stations in the cockpit and between the cockpit and the ground mechanic through the External Power Control Panel. SELCAL (SELective CALling) The SELCAL system provides the crew with visual and aural warnings from ground stations equipped with a coding device.

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1

1 OXY STWG BOX

OXY STWG BOX

OXY STWG BOX

SDAC

FWC CFDS

(aural Call Indication) 1

aural Warnings (FWC,GPWS,TCAS)

4TH OCCUPANT (parallel to 3rd Occupant)

Figure 1

(

)

AMU Schematic

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AUDIO CONTROL PANEL PRESENTATION GENERAL Three basic Audio Control Panels are provided in the cockpit for the Captain, First Officer and 3rd occupant. Two other optional ACPs can be installed, one in the cockpit for the 4th occupant and one in the avionics bay for ground service. Each Audio Control Panel (ACP) allows : the use of various radio communication and radio navigation facilities installed in the aircraft for transmission and reception of the audio signals, the display of various calls (SELCAL, ground crew calls and calls from the Cabin Attendants), the use of flight, cabin and service interphone systems. The Audio Control Panels (ACPs) are connected to the Audio Management Unit (AMU) via an ARINC 429 bus. TRANSMISSION KEY The front face features : seven rectangular pushbutton keys for transmission. Transmission channel selection : when a transmission key is pressed (CALL, MECH or ATT), three green bars come on. The selection is accepted (e.g : VHF1): the selected system is ready for transmission. only one radio system can be selected at a time for transmission. When a new transmission key is pressed, the green bars come on and the previously selected key is disabled. When a SELCAL/CALL, MECHanic or ATTendant call is received, the associated system key flashes amber and a buzzer sound is heard. CALL : For a SELCAL/CALL (HF/VHF). MECH : For a ground mechanic call. ATT : For a call from Attendant station.

PASSENGER ADDRESS (PA) KEY The PA key is used for Passenger Address announcements. When the Passenger Address (PA) key is pressed, three green bars come on (not LH-version). Boomsets, oxygen masks or hand-microphones can be used for Passenger Address announcements. (The PA key must be pressed and held) RECEPTION KNOB The fifteen Reception knobs, with associated potentiometers, are used for the selection of reception channels and adjustement of the received audio signals. The 15 reception knobs are also pushbutton switches of the pushpush type : Pressed in : The reception is inhibited Released out : Reception Knob comes on white and the reception is active. ON VOICE The ON VOICE key is used for attenuating morse code identification signals from ADF and VOR/DME navigation systems, in order not to hinder voice reception information. When the VOICE pushbutton key is pressed, the ON legend comes on green. RESET The RESET key cancels any amber lighted calls and buzzer sounds. INT/RAD SWITCH The INTERPHONE/RADIO selector switch is used for selecting radio or interphone mode. It is a three-position switch. Neutral position : The transceiver is in reception mode. RAD position (moment position): The radio system selected on the ACP changes from reception mode to transmission mode. For transmission, the switch must be held in the RAD position. INT position (fix position): The flight interphone operates regardless of the transmission key selection.When the PTT is activated, the interphone is cut : Radio transmission has priority over INT selection on the ACP.

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ACPs

Figure 2

AMU Audio Control Panel

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AUDIO SWITCHING General The AUDIO SWITCHING selector is used in case of communication failure on captain or first officer channels. Norm Position This positon corresponds to the normal allocation of the ACPs F/O 3 Position In this postion, the first officer is switched on the 3rd occupant part of the AMU controlled by the 3rd occupant ACP. The first officer now uses the 3rd occupant ACP. The 3rd occupant Audio equipment can not be used. CAPT 3 Position In this postion, the captain is switched on the 3rd occupant part of the AMU controlled by the 3rd occupant ACP. The captain now uses the 3rd occupant ACP. The 3rd occupant Audio equipment can not be used. Note: If the switch is in the CAPT 3 or F/O 3 position, the message ”AUDIO 3 XFRD” is displayed in green on the ECAM MEMO display.

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AUDIO SWITCHING

ÎÎÎÎ ÎÎÎÎ ÎÎÎÎ ÎÎÎÎ

CAPT 3

NORM

F/O 3 48 VU

SDAC (AUDIO XFR on ECAM) F/O ACP F/O AUDIO EQUIPMENT

F/O 3rd OCCUPANT ACP

3rd OCCUPANT AUDIO EQUIPMENT

3rd OCCUPANT CAPT ACP

CAPT AUDIO EQUIPMENT

CAPT

AMU

Figure 3

AMU Audio Switching Schematic

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POWER SUPPLY The system components are supplied with 28VDC from DC BUS1 and DC ESS BUS via 2 sub-busbars 101PP and 401PP respectively. Busbar 101PP Supply of the 3rd Occupant ACP and its associated electronic circuit located in the AMU via 3A circuit breaker: COM NAV/ACP/THIRD/OCCPNT (121VU) Supply of the calls card in the AMU via 3A circuit breaker: COM NAV/SELCAL (121VU) Busbar 401PP Supply of the Captain ACP and its associated electronic circuit located in the AMU via 3A circuit breaker: COM/AUDIO/ACP/CAPT (49VU) Supply of the 1st Officer ACP and its associated electronic circuit located in the AMU via 3A circuit breaker: COM/AUDIO/ACP/F/O (49VU) Supply of the Flight-Interphone Electronic Card located in the AMU via 3A circuit breaker: COM/AUDIO/FLT/INTPH (49VU)

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CALLS CARD BITE

Figure 4

AMU Power Supply Schematic

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DESCRIPTION The audio management unit (AMU) ensures the interface between the user (jack panel and ACP) and the various radio communication and radio navigation systems. The AMU ensures the following functions : Transmission Reception SELCAL and display of ground crew and Cabin Attendant calls Flight interphone Emergency function for the Captain and First Officer stations It also serves to record communications (FAA recording) and is equipped with a TEST circuit (BITE). This TEST circuit enables the AMU to be connected to the CFDIU. The AMU comprises 3 independent channels associated with the 3 ACPs. Each channel comprises : its reception function its transmission function its logic processing function its power supply The SELCAL, BITE and Flight Interphone sections are connected to the different channels.

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FWC

AMU

Figure 5

AMU Detailed Schematic

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EMERGENCY FUNCTION General The emergency function is used in case of loss of communications on the Captain or the First Officer channels. This function switches the Captain or First Officer communications to the 3rd Occupant station. In this case, the Captain (or the First Officer) uses the ACP located on the overhead panel to make his microphone or audio selections. Operation The AUDIO SWITCHING selector-switch, located on the overhead panel is used to switch to emergency configuration.Turning this switch, sends a ground to the Captain (or First Officer) and 3rd Occupant switching relays. The various microphone inputs, commands and audio outputs are connected to the microphone inputs, commands and audio outputs of the 3rd Occupant. This switchover is indicated on the upper ECAM display unit. Message: ”AUDIO 3 XFRD” NOTE : When the emergency function is activated, the various audio inputs and outputs at the 3rd Occupant station are no longer connected to their circuit. Therefore, the 3rd Occupant cannot use his audio integrating circuits.

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AMU

AMU

log 1 = GND

Figure 6

AMU Emergency Switching Schematic

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TRANSMIT CIRCUIT Transmission with boomset The analog signals of the boomset microphone are connected to the OdB gain amplifier then sent to the output transformer. At the transformer output, the switching module switches these signals to the transmitter selected by the ACP in accordance with information received from the logic processing function. Transmission with oxygen mask microphone In normal flight configuration, the oxygen mask microphone is not connected to the microphone circuit. Operation is as follows in flight configuration with use of oxygen mask. This system sets a control switch contained in the stowage box of the oxygen mask to the ground. This activates the relay which sets the oxygen mask into service. The pressurization of the oxygen circuit when masks fall out automatically activates this control switch.

Transmission on passenger address channel Transmissions can be made on the passenger address channels in 2 ways: In normal configuration, use the handset installed aft of the pedestal to make the PA announcements. This handset is part of the cabin intercommunication data system (Ref. ATA 23-73-00, Circuit RH). In RADIO configuration, use the rectangular PA pushbutton switch located on each ACP to make the passenger address announcements. This pushbutton switch is unstable, i.e. hold it pressed to make the announcements : this avoids unwanted transmissions. The electronic processing of this channel is identical to that of the other transmission channels. The operation of this pushbutton switch can be made identical to that of the other transmission channels (stable operation) : to achieve this, modify the AMU pin-program.

Transmission with hand microphone The hand microphone can be used in two ways : Radio transmission The logic processing card associated with relay K1 delivers a command. This command supplies relay K1 (AND function between the PTT switch of the hand microphone and the selected radio transmission, except for INT). Relay K1 directly connects the hand microphone to the transmission selection circuit. The station selected in transmission mode then supplies the hand microphone. Flight Interphone transmission When INT transmission is selected, relay K1 is not supplied ; the logic processing card associated with relay K2 delivers a command. This command supplies relay K2 (AND function between the PTT switch of the hand microphone and the INT transmission selection). Relay K2 connects the power supply of the boomset microphone to the hand microphone. This system applies the analog signals of the hand microphone to the OdB amplifier, then to the INT channel via the transmission line. This removes the microphone power supply from the interphone amplifier.

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K1 = Handmike PTT and NOT(INT) K2 = Handmike PTT and INT

Figure 7

AMU Power Supply MIC

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MUTING CIRCUIT The feedback produced by the loud speaker - microphone acoustic coupling when the microphones are used (acoustic feedback) is eliminated by a muting circuit. To achieve this, the muting circuit reduces the gain and/or the frequency range of the loud speakers. This attenuating circuit is controlled by the PTT switch of any of the radio communication microphones. The attenuating circuit is an integral part of the loud speakers. The logic processing channel receives PTT switch type information. From this information it activates the muting module. A ground is sent to the loud speaker units which set the direct muting function into service .

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WARNINGS

Figure 8

AMU Muting Circuit Schematic

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FLIGHT INTERPHONE General The flight interphone enables : telephone conversations between the various stations in the cockpit telephone conversations between the cockpit and the ground crew via the external power panel. The input signal from the various microphones used in the aircraft (hand microphone, boomset, mask microphone) is applied to inputs 1 to 7. A specific power supply circuit is provided for the microphones of inputs 6 and 7 (they have no transmission card to supply them). A current detection circuit on channel 6 and a cut-off relay on channel 7 cuts off the channels when they are not used. The L/G relay controls this cut-off relay. The amplified LF output signal is then available on the 3 windings of the secondary of the output transformer : 600 ohm output for ground crew 600 ohm output for audio output No. 6 2.2 Kohm output for the various AMU audio cards.

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(n.u.)

FLIGHT

GROUND

LGCIU INPUT 7 EXT POWER PANEL

Figure 9

AMU Flight Interphone Schematic

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VOR / ILS / DME SWITCHING Purpose In normal configuration, the DME reception is coupled with the VOR reception. However, in certain ILS or MLS approach conditions, the DME used must be aurally identified. The DME reception must therefore be coupled with the ILS or MLS reception. Operation The ND (Navigation Display) mode selector switch or the ILS pushbutton switch is used for switching control (see ATA 31 - DMC circuit ). Action on one of these commands sends a ground to the switching relays which connect the DME receptions to the ILS or MLS receptions.

VOICE ON/OFF FUNCTION

A compensation amplifier is provided to compensate for the insertion losses of this filter. Action on the VOICE/ON switch located on each ACP switches the attenuation filter into or out of service. Released position, VOICE/ON off The filter is not used, the operator simultaneously receives the marker identification and the voice transmission. Pressed in position, VOICE/ON on A command from the CPU sets the filter into service. The 1000 -1020 - 1350 Hz frequencies are greatly attenuated. Only the voice transmissions are audible. NOTE : The audio outputs of the communication channel and the ILS, MLS, MKR navigation do not transit via the filtering module.

Purpose The VOR, ADF navigation ground stations transmit a morse code which is used to identify them. However, certain stations, in addition to their code, transmit recorded voice information. This information informs the crew of subjects such as : latest weather information, state or special information concerning terrains etc. (e.g. : ATIS station). In order not to hinder the reception of this information, the VOICE/ON function greatly reduces the morse code reception. It is attenuated until it becomes practically inaudible while this information is being transmitted. Operation The transmission modulation frequency for ground station codes is 1020 Hz. However, certain onboard equipment receive a 1020 Hz frequency-modulated signal and at same time transmit this signal at 1000 Hz to the audio system. The 1000 Hz signal is generated by their synthetizer (the aeronautical standards specify that the ADF ground stations must be modulated at a frequency of 1020 Hz plus or minor 50Hz). Furthermore, the DME reception is coupled to the VOR reception (in normal operation). Thus the DME marker identification-code is transmitted with a frequency modulation of 1350 Hz. The filtering circuit of the navigation channels therefore comprises an attenuater filter for the reception bands of the ADF and VOR systems. This filter attenuates the 1000, 1020 and 1350 Hz frequencies by more than 32 dB.

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ILS/DME

VOR/DME ADF

Figure 10

AMU DME Switching and Voice ON/OFF Function

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CALLS Selective Call The SELCAL-CALL system of the audio management system gives a visual and aural indication of the calls from the ground stations equipped with a coding device which can be used by the aircraft installation (SELCAL system Selective Calling). The calls are sent on the radio frequencies which link the aircraft to the ground. The communication channels used are : VHF1 - VH2 and VHF3, HF1, HF2 if installed The aircraft receivers detect and capture the call signals transmitted by the ground stations (VHF or HF). Once detected, the signals are sent to the AMU SELCAL card. This SELCAL card is equipped with 5 inputs. These inputs correspond to the various communication facilities on the aircraft (VHF1 - VHF2 - VHF3 - HF1 HF2 in accordance with aircraft definition). The SELCAL decoder permanently scans the 5 inputs on which the calls may be present. It analyzes the received signals to check if they comprise the frequencies relevant to aircraft code. The operator programs this code on the SELCAL code panel. If the frequencies and aircraft code correspond, the warning system transmits an aural signal. The CALL legend on each ACP associated to the system which received the call (VHF1 - VHF2 - VHF3 - HF1 - HF2) comes on. Press the RESET pushbutton switch located on each ACP to reset the aural and lighted call.

Ground Crew Call This circuit displays the call from the ground crew in the cockpit. NOTE : Chapter ATA 23-42 (Cockpit-to-Ground Crew Call System-Circuit WC) gives the operation of the ground crew call circuit. When Capt Call pushbutton switch (located on external power panel 108VU) is pressed, it sends ground information to the Call card. The information is processed and a message is sent to the various audio cards and then to the ACPs. This causes the MECH legend to flash (coupled with INT transmission pushbutton switch) for 60 seconds. After 60 seconds, or when the RESET pushbutton switch is pressed, like the SELCAL system, the circuit is re-initialized. Cabin Attendant Call This circuit displays calls made from the cabin by the Cabin Attendants in the cockpit. NOTE : Chapter ATA 23-73 (Cabin Intercommunication Data System Circuit RH) gives the operation of the call circuit. When a call is made from the Cabin Attendant station, the CIDS generates ground information. This information is sent to the Call card. The information is processed then sent to the various audio cards and then to the ACPs. On the ACPs, this causes ATT legend to flash (coupled with CAB pushbutton switch) for 60 seconds. After 60 seconds or when the RESET pushbutton switch is pressed, like the SELCAL system, the calculating unit re-initializes the circuit. It also sends information back to the CIDS for re-initialization . NOTE : It is possible to inhibit the automatic function which causes the MECH and ATT flashing call legends to stop.

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COCKPIT CALL

Figure 11

ACP Call Indications

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L

FAULT ISOLATION AND BITE The audio system BITE (Built-In Test Equipment) serves as an aid for line maintenance in workshop and Service Department. It is used when faulty units are to be detected, replaced or repaired. It limits the number of unwanted removals of the system components. The BITE : Constantly transmits the actual status of the system (availability-unavailability). Memorizes any failures which occurred during the 63 previous flight segments or up to memory capacity. Monitors the data exchanges between the system components. Centralizes the triggered tests or self-test results. Dialogs with the CFDIU by means of menus An additional function is the transmission of the pin-program and of a message which serve to identify the system. General Operation The BITE may operate in two modes : the normal mode the menu mode. Normal Mode This mode cyclically interrogates the AMU cards in order to know their status and the status of the associated ACP. It transmits this information to the CFDIU and if a failure is detected, records this information in the failure memory. It interrogates the cards one after each other every 13 ms. The processing card generates information with respect to the self-test of this card and the data from the associated ACP. This information is sent to the BITE. This information is sent to CFDIU cyclically. Menu Mode This mode is used only on the ground. It enables dialog between the AMU and an operator via the multipurpose control and display unit (MCDU). An airground discrete gives the ground-flight information. The LGCIU (Landing Gear Control and Interface Unit Circuit GA) delivers the air-ground discrete.

The different menu selections are: LAST LEG REPORT PREVIOUS LEGS REPORT LRU IDENTIFICATION TEST AMU ACP AUDIO TEST AMU SELCAL CALL Transmission of PIN-PROGRAM The BITE circuit generates a message giving the installation status of certain equipment (VHF3 - HF1 - HF2 - ADF2). This message is generated from the information sent by the SELCAL card which receives the pin-program. It is sent to the CFDIU. The CFDIU system requires this information in order to transmit this information to the relevant circuits such as the RMPs (Radio Management Panels) and the SDAC (System Data Acquisition Concentrator). CFDS Messages Faults detected by the system and transfered to the CFDIU causes the following messages displayed on MCDU screen: FAULT ACP X BITE detected a faulty ACP X. AUDIO NO DATA FROM ACP X There is no communication between AMU and ACP X. AUDIO NO DATA FROM CFDIU No connection to the CFDS FAULT SELCAL The SELCAL part of the AMU is faulty. FAULT CALL The CALL part of the AMU (Att call, ground crew call) is faulty.

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ÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂ ÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂ ÂÂÂÂ ÂÂÂÂ ÂÂÂÂ

CFDS monitored

Figure 12

AMU CFDS monitored LRUs

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Figure 13

AMU MCDU BITE Menu

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Figure 14

AMS Location Cockpit

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Figure 15

AMS Location Cockpit and 80 VU

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23-42

GROUND CREW AND COCKPIT CALL SYSTEM

CALL SYSTEM PRESENTATION General The Ground Crew Call system enables the member to ground mechanic or ground mechanic to crew member calls. Ground mechanic to cockpit call When pressing the COCKPIT CALL pushbutton, the MECH light flashes amber on all ACPs and a buzzer is heard. An action on the RESET key on any ACP will make all MECH lights go off. Note: MECH lights go off automaticly after 60 sec if the call is not cancelled by the RESET key. Cockpit to ground mechanic call The horn sounds as long as the CALL/MECH pushbutton is pressed in and the cockpit CALL blue light on the panel 108 VU stays on. The RESET pushbutton makes the COCKPIT CALL light go off. Additional Horn Warnings The HORN can also be activated by following warnings: 26-13 APU FIRE on ground 21-26 BLOWERS LO FLOW on ground with engines shut down 34-14 ADIRS ON BAT on ground with engines shut down 25-65 ELT operation on ground 24-38 BATT discharge on ground

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CALLS BUZZER

MECH

FWD

CALLS

EMER

AFT

FWD

MID

EMER

EXIT

CALL

CALL

ON

ON

A319/320

MECH

CALL

CALL

CALL

CALL

CALL

MECH

ATT

VHF1

VHF2

VHF3

HF1

HF2

INT

CAB

ON VOICE

RESET

ALL

AFT

A321

INT PA

RAD VOR1

VOR2

MKR

ILS

MLS

ADF1

ADF2

FLT INT EXT PWR 108VU

HORN NOT IN USE

AVAIL

LIGHT TEST

COCKPIT CALL

ADIRU & AVNCS VENT

COCKPIT CALL

RESET

APU

FIRE

APU SHUT OFF

Figure 16

Call System Panels

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DESCRIPTION General The ground crew call system enables crew member-to-ground mechanic or ground mechanic-to-crew member calls. System Description The ground crew call system consists of : A CALLS/MECH pushbutton switch 1WC located on the overhead panel 21VU in the cockpit. It is associated with the RESET pushbutton switch 12WC located on the panel 108VU of the ground power receptacle. A mechanic call horn 15WC located in the nose gear well.The horn sounds to warn the mechanic of a call. A COCKPIT CALL indicator light 14 WC located on the panel 108VU.This indicator light comes on to warn the mechanic of a call. A COCKPIT CALL pushbutton switch 10 WC located on the panel 108VU. This pushbutton switch enables the ground mechanic to call the crew members via the circuit WW for the audio function and circuit RN for the visual indication. The system operates on the ground only, with the left and right main landing gear shock absorbers compressed. However, in flight, if the LGCIU is not energized, the ground crew call is activated following pilot’s action.

Ground Mechanic-to-Crew Member Call When pressing the COCKPIT CALL pushbutton switch 10WC, a ground signal is applied to the FWCs (31-52) triggering the buzzer circuit which feeds the aural warning signal to the loud speakers. This ground signal is applied to the circuit RN for the illumination of the MECH legend on the ACPs. Crew Member-to-Ground Mechanic Call During all the time the pilot presses the CALLS/MECH pushbutton switch 1WC located on the overhead panel, the mechanic call horn sounds. The blue COCKPIT CALL indicator light comes on. When the pilot releases the CALLS/MECH pushbutton switch, the mechanic call horn stops but the indicator light remains on. This indicator light goes off when pressing the RESET pushbutton switch 12WC located on the panel 108VU. In addition this system provides warnings for the following circuits : 26-13 APU FIRE on ground 21-26 BLOWERS LO FLOW on ground with engines shut down 34-14 ADIRS ON BAT on ground with engines shut down 25-65 ELT operation on ground 24-38 BATT discharge on ground

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GROUND MECHANIC TO CREW MEMBER CALL

ELT OPERATION BATT DISCHARGE

CREW MEMBER TO GROUND MECHANIC CALL AND ADD. WARNINGS

Figure 17

Call System Detailed Schematic

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Figure 18

Call System Location

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Figure 19

Call System Location Cockpit

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23-13

RADIO MANAGEMENT SYSTEM

RMP SYSTEM PRESENTATION The RMPs are used for the selection of radio communication frequencies. They are also used for the selection of radio navigation frequencies as back-up of the Flight Management Guidance Computers (FMGCs). There are 3 RMPs for frequency selection : Each RMP can control any VHF or HF system. RMP1 and RMP2 can control the radio navigation systems in back-up mode. RMP3 cannot control the radio navigation systems. The 3 RMPs permanently dialog so that each RMP is informed of the last selection made on any of the other RMPs If two RMPs fail, the remaining RMP controls all the VHF and HF transceivers. The transmission of data to the communication and navigation systems and the dialog between the RMPs are performed through data buses.

1 WINDOWS There are 2 display windows : The ACTIVE window displays the operational frequency. The STandBY/CouRSE window displays the standby frequency or the course in back-up navigation mode. The windows are liquid crystal displays with a high contrast.

RMP2 allocated with VHF2 RMP3 allocated with VHF3, HF1/2. If VHF2 is selected on RMP1, the SEL light comes on WHITE on RMP1 and RMP2.

4 DUAL SELECTOR KNOB The DUAL SELECTOR KNOB is used for the selection of the frequency/course displayed in the STandby/Course window.

5 ON/OFF SWITCH The latching ON/OFF switch allows the crew to set the RMP on or off.

6 TRANSFER P/B When the TRANSFER key is pressed, the operational frequency becomes the STandBY frequency and the STandBY frequency becomes the operational frequency.

7 AMPLITUDE MODULATION KEY The Amplitude Modulation (AM) key is associated with the HF system for communication with stations using amplitude modulation transceivers.

8 NAVIGATION KEYS

2 COMMUNICATION KEYS There are 5 pushbutton keys for the radio communication systems. When a key is pressed, the ACTIVE and the STandBY frequencies are automatically displayed in the dedicated windows.

3 SEL INDICATOR

The NAVigation guarded pushbutton key allows the radio navigation systems to be selected, in back-up mode only, when the Flight Management Guidance Computers (FMGCs) are failed. In radio navigation back up mode, navigation frequency/course selection is performed using the dual selector knob.

The SEL indicator light comes on WHITE, when a non dedicated Radio Management Panel takes control of the system frequency selection. The normal configuration is : RMP1 allocated with VHF1

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A

B

FMGC 1

1

NAV RECEIVERS SYS 1

1

6

VHF 1 (HF 1) TRANSCEIVER VHF 3 TRANSCEIVER

RMP 1

4

RMP 3

ACARS

VHF 2 (HF 2) TRANSCEIVER

RMP 2 2

3

7

8

5 NAV RECEIVERS SYS 2

FMGC 2 B Figure 20

A

RMP Schematic

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POWER SUPPLY RMP 1 The RMP1 is supplied with 28VDC from the 28VDC ESS BUS 4PP (sub-busbar 401PP) through 3A circuit breaker 2RG1 on the overhead panel 49VU (in the cockpit). The RMP1 is supplied by the emergency system. RMP 2 The RMP2 is supplied with 28VDC from the 28VDC BUS 2PP (sub-busbar 204PP) through 3A circuit breaker 2RG2 on the rear C/B panel 121VU (in the cockpit). RMP 3 The RMP3 is supplied with 28VDC from the 28VDC BUS 1PP (sub-busbar 103PP) through 3A circuit breaker 2RG3 on the rear C/B panel 121VU (in the cockpit).

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Figure 21

RMP Power Supply Schematic

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RMP DESCRIPTION Operation The RMPs have two modes of operation : the normal mode the radio-navigation back up mode. Normal Mode In normal mode the RMPs control the frequencies of the VHF1, VHF2 and HF1 transceivers. For frequency control on the VHF3 system, refer to the ACARS. The operating frequencies of all the transceivers can be displayed and modified on one RMP. The RMPs exchange the various frequencies selected for the transceivers through dialogue buses. Any new selection made on one RMP is taken into account by the two others. Each RMP has two output buses connected to the radio communication equipment : The RMP1(2) COM BUS 1 delivers the VHF1 and HF1 frequencies. The RMP2 COM BUS 1 delivers the VHF3 frequencies. The RMP2(1) COM BUS 2 delivers the VHF2 frequencies. Each transceiver receives the appropriate output bus from the RMP1 and RMP2. The transceiver only takes into account one of the two signals (depending on the status of a discrete received from the RMP1 or 2). In addition, the RMP1 or the RMP2 (set to OFF) can be made transparent for the RMP3 (its output buses are linked to the RMP1 and RMP2 only). In the event of failures of one or two RMPs, the reconfigurations are possible to control the radio communication equipment.

Radio-Navigation Back Up Mode This mode is selected in the event of failure of both FMGCs, on the RMP1 and the RMP2 only. In addition to normal mode functions it also enables the frequency control of the radio navigation equipment : on Captain side (VOR1, DME1, ILS1, ADF1) for the RMP1 on First Officer side (VOR2, DME2, ILS2, ADF2) for the RMP2. The RMP1 and the RMP2 transmit on a dedicated output bus the frequencies to the radio navigation equipment. In addition, the RMP1 (RMP2) receives the FMGC1 (FMGC2) management bus. In normal mode, these input and output are directly interconnected by means of internal relays. The RMP is thus transparent to the onside FMGC. In radionavigation back up mode, the output bus transmits frequencies generated by the RMP. Each radio-navigation system receives the output bus from the onside RMP and the management bus from the offside FMGC. Only one input is taken into account according to the status of a discrete received from the RMP. This enables reconfigurations in case of failure of one or two FMGCs. The RMP1 and the RMP2 exchange, through the dialogue buses, the frequency and the course for the ILS : the selected values are identical for the ILS1 and the ILS2 at selection of the back up mode on the RMP1 and the RMP2. The ILS course and frequency are the only radio navigation data exchanged through the dialogue buses.

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Figure 22

RMP Detailed Schematic

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Figure 23

RMP COM Tuning Architecture

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Figure 24

RMP COM Tuning Architecture

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Figure 25

RMP NAV Tuning Architecture

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1

1 Figure 26

for ILS frequency transfer only

RMP NAV Tuning Architecture

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RMP COMMUNICATION TUNING The radio management panels (RMP) are used for radio communication equipment frequency selection. They are also used for radio navigation equipment frequency selection in back up mode. When the ON/OFF switch is set to on, the RMP displays the frequency previouly selected. By means of the dual selector knob the desired frequency can be selected in the stand by window. The transfer pushbutton must be pressed to render it active and the displayed values are changed over. The RMP modifies its output data accordingly. Note : only the stand-by frequency can be modified by means of the dual selector knob. The new active frequency is transmitted to all RMPs through the dialog buses. When the VHF2 tranceiver is selected on RMP 1 the SEL indicator lights on RMP 1 and RMP 2 come ON. The AM pushbutton controls the selection of the amplitude modulation (AM) mode for the HF transceivers. By default, the single side board (SSB) mode is selected on the corresponding HF system.This selection is memorized when another system is selected. The other RMPs take into account this selection through their dialog buses.

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Figure 27

RMP COM Tuning

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RMP NAV BACK UP TUNING We are going to study the actions to be performed for a back up tuning of an ADF frequency and a VOR/ILS frequency and course. The fist thing to do is to open the guard on the NAV key. When the NAV key is pressed in, the on side VOR/ILS and ADF receivers are controlled by the RMP and no longer by the FMGC. The green LED comes on indicating that you are in STANDBY tuning mode. When the STBY NAV key is pressed, (i.e VOR), its green LED comes on and the previously memorized frequency is displayed in both windows. The knob is turned to select a new frequency. First, the selected frequency is displayed in the STBY / CRS window. When the transfer key is pressed, the STANDBY frequency becomes ACTIVE and the active course is displayed in the right hand side window. The outer knob is turned to select a new course. to select another frequency, the transfer key must be pressed again to get the active frequency displayed in both windows. NOTE: The operation of course and frequency tuning is the same for VOR and ILS ADF tuning is performed as for ILS or VOR ecept that when the transfer key is pressed, the standby and active frequencies are interchanged.

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Figure 28

RMP NAV Back up Tuning

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RADIO NAV TUNING FROM RMP AND MCDU General The navaid selection includes tuning of the following sensors. VOR (frequency/course) and DME for display ILS (frequency/LOC course) ADF (frequency/BFO) There are three ways of selection which are : RMP selection (radio management panel) manual selection through the MCDU automatic selection (in FMGC software).

displayed. If frequency is entered, the ident field is filled if found in data base. If not, brackets are displayed.

RMP Selection The RMP selection in the radio nav architecture has to be considered as a back-up of selection. It is activated/deactivated upon selection of the nav mode for the RMP (NAV pushbutton switch). Since then, the pilot may select VOR, ILS or ADF. If selection of any RMP is active, neither the pilot nor the FMGCs can tune the radio frequencies on both sides. For display, the selected VOR ADF are shown on the navigation display with a character R near the ident or frequency to indicate that the navaid selection mode is RMP. On the MCDU, the RMP select navaids are displayed on the RADIO NAV page or PROG page in green small fonts. Manual selection through the MCDU Selection through the MCDU is possible through two pages : RADIO NAV page VOR tuning On RADIO NAV page, the pilot may select for display a VOR by identor frequency in line 1L, 1R. He may also optionally enter a course in line 2L, 2R. Upon modification of the selected VOR, the course is automatically cleared. Manually selected navaids are displayed in cyan large fonts on the MCDU and on the navigation display there is a character M near the navaid ident or frequency. - Selection mechanization If ident entry is made, the nav data base is searched and if there is a match, the FMGC outputs the frequency. If not, NEW NAVAID page is

If the VOR field is cleared, the display reverts to autotuned navaid with associated course (if any). ILS tuning On RADIO NAV page only, the pilot may select an ILS by frequency or ident in field 3L. The entry mechanization is the same as forVOR.However upon entry of an ILS by frequency, this frequency is compared : - In preflight and takeoff phases to the ILS frequency at origin - else to the ILS frequency at destination. In both cases, if a match is found, the ident and frequency are displayed (cyan small fonts for the ident, cyan large fonts for the frequency). If not, only the frequency is displayed (in cyan large fonts) and a message appears in scratchpad RWY/ILS MISMATCH. In field 4L, the pilot may select the LOC course. This will be used for LOC capture and ILS guidance in approach. This LOC course may only be entered through the MCDU on the RADIO NAV page. It is cleared if the pilot changes the selected ILS. ADF tuning With the same mechanization as for VOR, the pilot may select an ADF by ident or frequency in line 5L, 5R. Since the second ADF is an aircraft option (program pin on FMGC), the second ADF is available only when this option is valid. When an ADF is selected, the ADF BFO prompt appears in line 6. Selection of the BFO operation by pressing the LS key displays the prompt ADF BFO and activates the BFO function for the current ADF frequency selection. The BFO operation is deactivated by clearing the associated field. The display reverts to ADF BFO. It is also deactivated by entering a new ADF frequency or ident. Automatic selection Automatic selection is performed in the FMGC software. From a display point of view, autotuned VOR, ILS or ADF are displayed on RADIO NAV page or PROG page in cyan small fonts. On navigation display, there is no indicator M or R near the VOR or ADF for display showing that the navaid is autotuned.

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2

1 1 ) If selected Station is valid, the Station Identifier is displayd instead of the frequency. 2 ) Tuning Mode: R Tuned via the RMP M Tuned via the MCDU Nothing when auto tuned by the FMGC

Figure 29

MCDU and RMP NAV Tuning

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FAULT ISOLATION AND BITE The BITE facilitates maintenance on in-service aircraft. The BITE detects and determines a failure related to the RMP. The BITE of the RMP is connected to the Centralized Fault Display Interface Unit (CFDIU). The BITE : transmits permanently RMP status and an identification message to the CFDIU. memorizes the failures occured during the last 63 flight legs. monitors data input from the various peripherals (VHF, HF and CFDIU). transmits to the CFDIU the result of the tests performed and self-tests. can communicate with the CFDIU by the menus.

CFDS Messages Faults detected by the system and transfered to the CFDIU causes the following messages displayed on MCDU screen: RMP X NO DATA FROM RMP Y There is no communication between RMP X and RMP Y. RMP X NO DATA FROM FMGEC 1 (2) There is no communication between RMP X and FMGEC 1 (2). NO DATA FROM CFDIU No conection to the CFDS

General Operation The BITE may operate in two modes : the normal mode the menu mode. Normal Mode During the normal mode the BITE monitors cyclically the momentaneous status of the RMP. It transmits these information signals to the CFDIU during the flight concerned. In case of fault detection the BITE stores the information signals in the fault memories. Menu Mode The menu mode can only be activated on the ground. This mode enables communication between the CFDIU and the RMP BITE by means of the MCDU (Multipurpose Control Display Unit). The RMP menu mode is composed of : LAST LEG REPORT PREVIOUS LEGS REPORT LRU IDENTIFICATION TROUBLE SHOOTING DATA TEST. Note: Only RMP 1 (or RMP 3, if RMP 1 is switched off) is connected to the CFDIU. The other RMPs are tested via RMP 1 (or RMP 3)!

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NAV RECEIVERS SYS 1

FMGC 1

CFDS

ÂÂÂÂÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂÂÂÂ

VHF 1/3 (HF 1) TRANSCEIVER

RMP 1

ÂÂÂÂ ÂÂÂÂ

RMP 3

CFDS monitored

RMP 2

VHF 2 (HF 2) TRANSCEIVER

FMGC 2

Figure 30

NAV RECEIVERS SYS 2

RMP CFDS monitored LRUs

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Figure 31

RMP MCDU BITE Menu

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LOCATION

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Figure 32

RMP Location

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23-12

VHF SYSTEM

VHF SYSTEM PRESENTATION General The VHF is used for short range voice communications. The VHF system allows short distance voice communications between different aircrafts (in flight or on ground) or between the aircraft and a ground station. Principle Let’s see the main components of the VHF system. For voice communications, the crew use acoustic equipment. 2 side-stick radio selectors. 2 loudspeakers. 3 oxygen-masks. Facilities for boomsets, headsets and hand-microphones. The Audio Management Unit (AMU) acts as an interface between the crew and the VHF system. The Audio Control Panels (ACPs) allow selection of the VHF1,2 or 3 transceiver in transmission or reception mode and for the control of the received audio signal. The Radio Management Panels (RMPs) serve to select the VHF frequencies. The VHF transceiver, tuned on the frequency selected by one of the 3 Radio Management Panels (RMPs), transforms the audio signals into VHF signals (in transmission mode) or VHF signals audio signals (in reception mode). Note : The VHF3 is dedicated to ACARS system, but can be used for radio voice communications.

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Figure 33

VHF Schematic

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POWER SUPPLY VHF1 System The VHF1 system is supplied with 28VDC : from the 28VDC ESS BUS 4PP (sub-busbar 401PP) through circuit breaker 2RC1 located on the overhead panel 49VU, in the cockpit. The VHF1 system is supplied by the emergency system. VHF2 System The VHF2 system is supplied with 28VDC : from the 28VDC BUS 2 2PP (sub-busbar 204PP) through circuit breaker 2RC2 located on the rear panel 121VU, in the cockpit. VHF3 System The VHF3 system is supplied with 28VDC : from the 28VDC BUS1 1PP (sub-busbar 101PP) through circuit breaker 2RC3 located on the rear panel 121VU, in the cockpit.

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Figure 34

VHF Power Supply Schematic

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DESCRIPTION Transmit Function The audio signals from the microphones are transmitted to the VHF transceiver through the AMU. The VHF transceiver tuned on the frequency selected on one RMP, transforms the audio signals into VHF modulated signals. The VHF signals are fed to the antenna by a coaxial cable. They are then transmitted to the various stations. A connection between the VHF transceiver and the SDAC enables to record the use of the VHF system in transmit mode. The connection is obtained through the PTT switch. Receive Function The antenna picks up the VHF radio-communication signals from the stations. These signals are transmitted to the transceiver by a coaxial cable. The transceiver, tuned on the frequency selected on one RMP demodulates the VHF received signals into audio signals. The AF signals are transmitted via the AMU, to the audio equipment or SELCAL system. Tuning 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 (port select) delivered by the frequency control system. The data corresponding to the frequency selected on the RMP is sent to the transceiver through an ARINC 429 bus. This serial word contains the label, the source / destination identifier, the frequency data, the status and the parity bit. The ACARS MU applies a command signal to the VHF3 to take into account its frequency inputs through the port select discrete. when this discrete is a ground signal, the VHF3 takes into account input A and operates on the frequency transmitted by the ACARS MU. when this discrete is a open circuitl, the VHF3 takes into account input B and operates on the frequency transmitted by the RMPs. Signals The LGCIU indicates the aircraft status (flight or ground) for flight leg switching.

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A B

A B

B A

Port Control

Figure 35

VHF Detailed Schematic

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FAULT ISOLATION AND BITE The BITE facilitates maintenance on in-service aircraft. The BITE detects and determines a failure related to the VHF system. The BITE of the VHF transceiver is connected to the Centralized Fault Display Interface Unit (CFDIU). The BITE : transmits permanently VHF system status and an identification message to the CFDIU. memorizes the failures occured during the last 63 flight legs. monitors data input from the various peripherals (RMP and CFDIU). transmits to the CFDIU the result of the tests performed and self-tests. can communicate with the CFDIU by the menus. General Operation The BITE may operate in two modes : the normal mode the menu mode. Normal Mode During the normal mode the BITE monitors cyclically the momentaneous status of the VHF system. It transmits these information signals to the CFDIU during the flight concerned. In case of fault detection the BITE stores the information signals in the fault memories.

ECAM Message A connection between the VHF tranceiver and the SDAC enables to record the use of the VHF System in transmit mode (PTT). If the system is in transmit mode longer than 60s the following message appears on the ECAM: COM : VHF-X CONT EMITTING CFDS Messages Faults detected by the system and transfered to the CFDIU causes the following messages displayed on MCDU screen: VHF-X TRANSCEIVER A transceiver fault has been detected VHF-X: NO DATA FROM CONTROL SOURCE No data from RMPs VHF-X: NO DATA FROM CFDIU No conection to the CFDS CHECK VHF-X ANTENNA CIRCUIT A antenna fault or a antenna coaxial cable fault has been detected.

Menu Mode The menu mode can only be activated on the ground. This mode enables communication between the CFDIU and the VHF transceiver BITE. This is by means of the MCDU (Multipurpose Control Display Unit) of the maintenance system. The VHF transceiver menu mode is composed of : LAST LEG REPORT PREVIOUS LEGS REPORT LRU IDENTIFICATION TEST.

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RMPs

ÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂ ÂÂÂÂ ÂÂÂÂ ÂÂÂÂ VHF TRANSCEIVERS

AMU

SDAC

CFDS

Figure 36

CFDS monitored

VHF CFDS monitored LRUs

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Figure 37

VHF MCDU BITE Menu

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FRONTPANEL TEST After installation, correct operation of the VHF transceiver can be checked by using the following controls located on the transceiver face : SQL/LAMP TEST pushbutton switch When pressing the SQL/LAMP TEST pushbutton switch : - the squelch is disabled and allows background noise to be heard - the green LRU PASS and red CONTROL INPUT FAIL annunciator lights come on (lamp test). TEST pushbutton and CONTROL INPUT FAIL and LRU PASS annunciator lights When pressing the TEST pushbutton switch : - the green LRU PASS indicator light comes on for 1s approximately to indicate correct operation - the red CONTROL INPUT FAIL warning light is off The red warning light comes on to indicate control data failure (control unit or bus line). - the stationary wave ratio appears in the front display. RFL/OFF/FWD selector switch When placing the RFL/OFF/FWD selector switch in FWD and RFL positions, the forward and reflected powers appear respectively in the front display.

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Figure 38

VHF Front Panel Test

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Figure 39

VHF Location 80 VU

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Figure 40

VHF Location Cockpit

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23-11 HF-SYSTEM GENERAL The high frequency (HF) system serves for all long-distance voice communications between different aircraft (in flight or on the ground), or between the aircraft and one or several ground stations. The HF system operates within the frequency range defined by ARINC 719, (i.e. 2.8 to 23.999 MHz, with 1 KHz spacing between channels). The aircraft is provided with a single HF system. The HF system is composed of : one transceiver one antenna coupler one shunt-type antenna The HF system is associated with : the Radio Management Panels (RMP) which are centralized systems enabling the frequency display of the HF system and the mode switching the Audio Management Unit (AMU) for connection to the audio integrating and SELCAL systems the Centralized Fault Display Interface Unit (CFDIU) (by the MCDU) which is a centralized maintenance system the Landing Gear Control Interface Unit (LGCIU) which indicates the aircraft status (flight or ground) the System Data Acquisition Concentrator (SDAC) which collects transmission information from the HF system (COM: HF1 EMITTING if PTT longer than 60s) The HF1 system is supplied with three-phase 115VAC through 5A circuit breaker (1RE1) in cockpit panel 121VU, from sub-busbar 101XP. The HF1 transceiver (3RE1) provides the HF1 antenna coupler (4RE1) with 28VDC and monophase 115VAC.

interruption of the signal after 15 s approximately. triggering of the signal at each attempt to transmit. Operation The HF transceiver complies with the standards defined in ARINC 719. The transmission and reception of coded messages between the various control units (CFDIU, RMP) comply with ARINC 429. The RMP controls the various operations which are transmitted to the transceiver by a numeric message in compliance with ARINC 429. This message can be received by the port A or the port B of the transceiver. The RMP performs the selection by a discrete. A microprocessor performs the decoding of the frequency and mode (AM or USB). The microprocessor checks the message from the RMP and controls the system operation. In case of failure it controls the illumination of the lights located on the face and/or acts on the transmitter.

Indication of Transmission Out of Frequency Range The HF system is designed to operate within the frequency range from 2.8 to 23.999 MHz. However, an operational facility enables frequency display in the 2 to 29.999 MHz range on the RMP. If the out-of-range values of the HF transceiver are displayed on the RMP, the operating anomaly is indicated as follows : at first activation of the PTT switch : a 1000 Hz audio signal is triggered.

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Figure 41

HF System Schematic

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HF- ANTENNA COUPLER General The antenna coupler enables matching of the aircraft HF shunt-type antenna with the output circuit (50 ohms) of the HF transceiver. The coupler is a pressurized sealed box. The face features : a connector J1 for connection with the transmitter a coaxial connector J2 to connect the coaxial cable from the transmitter a connector J3 for test equipment connection a pressurizing valve a fault warning light a handle an identification plate Operation The coupler is tuned in six sequences : start, reception/standby, tune A, tune B, tune C and operational position. A sequence counter controls the six sequences.The counter starts the next sequence only when all the conditions related to the previous one are met. If a failure is detected during the tuning phase, tuning is stopped. The tuning phase is initiated at HF system energization or when a new frequency is selected. The tune control line is then grounded. Servomotors controlled by servo-amplifiers place the tuning elements in start position. Start sequence In this sequence, the capacitors and inductors are positioned so that they present minimum impedance to signals. When all these conditions are met, a pulse is applied to the sequence counter. The system is forced to the reception/standby phase. Reception/standby sequence In this position, the coupler is in reception condition and ready for a tuning cycle. PTT control grounding causes interlocking of couplers (case of dual system). A pulse is applied to the sequence counter and the system is forced to the next tuning sequence : tune A.

Tune A sequence The purpose of tune A is to adjust the antenna circuits so that HFsignal current and voltage are in phase. To this end after detection a discriminator delivers an errorsignal proportional to the phase difference during 50 ms. The polarity of this signal determines the elements required to achieve tuning. Tune B sequence The purpose of tune B is to match the antenna load with the transmitter output circuits. To this end, a load discriminator compares the HF current and voltage. This comparison gives an error voltage proportional to the difference between the HF circuit impedance and an impedance of 50 ohms. Tune C sequence The purpose of the tune C is to complete previous adjustments and obtain a VSWR (voltage standing-wave ratio) lower than 1.3. When a VSWR lower than 1.3 is obtained, the sequence counter controls start of the next sequence, i.e. operational position. Operational position In this sequence, the tuning control line is disconnected from ground. The antenna coupler can operate. If a new frequency is selected, the antenna coupler goes back to the start sequence and the tuning cycle starts again. Fault Indication Fault information of the coupler can be transmitted by discretes to the HF transceiver. In this case, the HF transceiver will take these items of information into account and will transmit them to the CFDIU.

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Figure 42

HF- System Antenna Coupler

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FAULT ISOLATION AND BITE The BITE facilitates maintenance on in-service aircraft. The BITE detects and determines a failure related to the HF system. The BITE of the HF transceiver is connected to the Centralized Fault Display Interface Unit (CFDIU). The BITE : transmits permanently HF system status and an identification message to the CFDIU. memorizes the failures occured during the last 63 flight legs. monitors data input from the various peripherals (RMP and CFDIU). can communicate with the CFDIU by the menus. General Operation The BITE may operate in two modes : the normal mode the menu mode. Normal Mode During the normal mode the BITE monitors cyclically the momentaneous status of the HF system. It transmits these information signals to the CFDIU during the flight concerned. In case of fault detection the BITE stores the information signals in the fault memories.

ECAM Message A connection between the HF tranceiver and the SDAC enables to record the use of the HF System in transmit mode (PTT). If the system is in transmit mode longer than 60s the following message appears on the ECAM: COM : HF-1 EMITTING CFDS Messages Faults detected by the system and transfered to the CFDIU causes the following messages displayed on MCDU screen: HF-X TRANSCEIVER A transceiver fault has been detected HF-X: NO DATA FROM CONTROL SOURCE No data from RMPs HF-X: NO DATA FROM CFDIU No conection to the CFDS HF-X ANTENNA CIRCUIT A antenna fault, a antenna coaxial cable fault or a coupler fault has been detected.

Menu Mode The menu mode can only be activated on the ground. This mode enables communication between the CFDIU and the HF transceiver BITE. This is by means of the MCDU (Multipurpose Control Display Unit) of the maintenance system. The HF transceiver menu mode is composed of : LAST LEG REPORT PREVIOUS LEGS REPORT LRU IDENTIFICATION CURRENT STATUS.

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ÂÂÂÂ ÂÂÂÂ

CFDS monitored

ÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂÂ Figure 43

HF System CFDS monitored LRUs

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Figure 44

HF- System BITE Menu Sheet 1

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Figure 45

HF- System BITE Menu Sheet 2

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Figure 46

HF System BITE Menu Sheet 3

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HF-TRANSCEIVER FRONTPANEL TEST On transceiver face are located: two jacks (PHONE and MIC) a SQL/LAMP TEST pushbutton switch three red warning lights : - LRU FAIL - KEY INTERLOCK - CONTROL INPUT FAIL a transportation handle an identification plate

Test Correct operation of the transceiver can be checked by means of the various lights on its face. LRU FAIL red light (LED) LRU FAIL red light comes on in the event of a transceiver warning such as : - output power drop (detected only if PTT is active) - microprocessor or synthesizer failure - power failure KEY INTERLOCK red light (LED) KEY INTERLOCK red light comes on when a failure is detected in antenna circuit (if PTT is active), such as : - coupler failure - excessive tuning time - excessive antenna reactance CONTROL INPUT FAIL red light (LED) CONTROL INPUT FAIL red light comes on when there are serial message faults such as : - abscence of label - insufficient refresh rate - message not valid. SQL/LAMP TEST pushbutton switch When pressing the SQL/LAMP TEST pushbutton switch, all the lights come on, the squelch is disabled and causes background noise to be heard in the headset.

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Figure 47

HF- System Transceiver & Fault Annunciator

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Figure 48

HF- System Location

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Figure 49

HF- System Antenna & Coupler Location

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23-24

ACARS

ACARS PRESENTATION Purpose The ACARS Data Link System is an air/ground communication network that enables aircraft to function as a mobile terminal associated with modern airline command, control and management systems. The ACARS is used to transmit or receive automatically or manually generated reports or messages to or from a ground station. The ACARS is dedicated to Maintenance, Operation and Commercial purposes. The choice of ACARS applications and the definition of the operational programs are under Airline responsibility because of high customization of the system. The ACARS is a Buyer Furnished Equipment (BFE). Principle The ACARS can manage both transmission or reception of data. Ground-to-air (uplink UL) and air-to-ground (downlink DL) digital messages are transmitted or received via the VHF3 transceiver. VHF3 is mainly dedicated to the ACARS Data Link System, but can be used as a backup for voice communications. The transmitted information is relayed via the ground stations to a central computer where data is converted into airline messages. A ground network (SITA for EUROPE, ASIA, AFRICA and SOUTH AMERICA, ARINC for the USA and CANADA and AVICOM for JAPAN), transmits the data from the ground receiver to the airline main base. SITA network is exclusively dedicated to the airline community, transmitting technical, commercial, flight operation and safety information. Any of the ACARS functions can be modified by the airline, through the ACARS MU programming. The unit needs a Operating Software and a Customer Database. Both are loadable via a portable Dataloader direct on the frontface of the MU (not via the installed Airborne Data Loader).

Components The ACARS Management Unit is connected to various computers : Flight Management function of the Flight Management Guidance and Envelope Computers (FMGECs). Central Fault Display and Interface Unit (CFDIU). Data Management Unit (DMU). Flight Warning Computers (FWCs) and the System Data Acquisition Concentrator (SDAC1) Air Data and Inertial Reference Unit (ADIRU3). Fuel Quantity Indication System (FQIS). Various units are used to control the ACARS MU : 2 Multipurpose Control Display Units (MCDUs). 1 Printer and 3 Radio Management Panels (RMPs), located in the cockpit. The Unit receives various discrete informations for several functions (e.g. A/C type).

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MCDU

VHF # 3

RMP 1 /2 /3

VOICE FREQUENCY

B

Control

DATA FREQUENCY

A

Control Control

ACARS MU 103 XP 115V AC BUS 1

PRINTER PORTABLE DATA LOADER

ADAPTER CABLE

FMGC 1/2 VARIOUS DISCRETES

A/C TYPE PIN PROGRAMS

AIDS DMU

SDAC 1 CFDIU ADIRS # 3

ACARS MU

FWC 1/2

FQIS

Figure 50

ACARS Schematic

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DESCRIPTION ACARS MU The ACARS Management Unit (MU) manages all tasks related to the ACARS. It controls both emission and reception of data through the VHF3 transceiver. The ACARS MU transmits data to the various aircraft systems through its two general output buses. It receives data from the avionics systems through their general input buses. The ACARS MU is supplied with 115 VAC. MCDU The ACARS MU is interfaced with the Multipurpose control and Display Units (MCDUs). The dialog between one MCDU and the ACARS MU is initiated when ACARS is selected on the MCDU menu. The MCDU enables the following functions : display of data generated by the MU, display of data transmitted by the ground or by peripheral computers, selection of the various ACARS MU functions, test and entry of data by the crew. MCDU 1 and 2 are connected to the ACARS MU. Only one can communicate with the system at a time. FMGEC The FMGECs send a pre-flight and a post-flight report via ACARS MU by manual action through the MCDUs. It also sends report on ground request via ACARS MU. The FMGECs also automatically send the in-flight report to the ACARS MU after take off. The airline can initialize and update the flight plan in the FMGECs through ACARS. Note: The ACARS - FMC Interface is not full active in the moment. ACMS DMU The ACARS MU transmits data to and receives data from the Aircaft Condition Monitoring System (ACMS). Each report generated by the ACMS can be programmed individually for transmission to the ACARS MU either automatically or manually. CFDIU The ACARS MU receives data from the CFDIU. The CFDIU transmits automatically or manually the following messages to the ACARS MU : post flight report on ground or current flight report in flight,

real time failure and real time warning in flight, BITE data messages and class 3 report on ground. The ACARS MU transmits its own maintenance information to the CFDIU (not active). The ACARS MU is provided from the CFDIU with the following : aircraft identification (tail number), flight number and flight phase, identification of departure and destination airports, date and time, installed optional systems. PRINTER The ACARS MU is connected to the multi-purpose cockpit printer. The ACARS MU can buffer data printing, when the printer is buzy with another system. ADIRU 3 The ADIRU 3 sends LAT/LONG information form frequency tuning and groundspeed information. FQIS The Fuel Quantity Indication System sends FOB, preselected Fuel e.g. information FWC/SDAC The ACARS MU receives parameters sent by the System Data Acquisition Concentrator (SDAC) 1 and the Flight Warning Computers (FWCs) 1 and 2. The parameters sent by SDAC 1 and FWCs allow the ACARS MU to establish the EVENT TIME OOOI (pax door closed, gear compressed...). The ACARS MU sends a status parameter to the FWCs. FWC 1 and FWC 2 display on the Engine Warning Display (EWD), one of four ACARS configurations provided by the ACARS MU. The four possible configurations are: ACARS MSG : an ACARS message has been received by the aircraft, ACARS STBY : loss of communication between aircraft and ground, VHF3 VOICE : VHF3 operates in VOICE mode, ACARS CALL : a message requesting a voice conversation has been received from the ground. A amber COM- ACARS FAULT message appers, when the FWCs do not receive normal information from the ACARS MU.

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SDAC1

ADIRS#3

FQIS

Figure 51

ACARS Detailed Schematic

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ACARS ACTIVATION VHF3 FREQUENCY SELECTION VHF3 can be tuned either via the ACARS MU automatically or manually using the MCDUs, or via the Radio Management Panels (RMPs) depending on the PORT SELECT discrete status. The PORT SELECT discrete status is manually controlled by the selection made on the RMPs, or automatically by the ACARS MU. When the PORT SELECT DISCRETE is grounded, the ACARS MU tunes the VHF3 through its input A. When the PORT SELECT DISCRETE is in open circuit, the RMPs tune the VHF3 through its input B. The frequency controlled from the MCDUs is used to force the MU to work with another frequency. In normal case, the frequency is tuned automaticly from the MU by using present position information from the ADIRUs. In case of missing this information or other failure a manuell selection of the 5 ACARS frequencies is possible via the MCDU. VHF3 AUDIO SELECTION VHF3 will handle audio information from ACARS MU or from the Audio Management Unit (AMU) depending on the VOICE DATA SELECT discrete status. When the VOICE DATA SELECT discrete is grounded, VHF3 handles audio information from the ACARS MU. When the VOICE DATA SELECT discrete is in open circuit, VHF3 handles audio information from the AMU. These selections can be defined by means of a pin program .

RMP Each RMP receives the PORT SELECT discrete. When this discrete is grounded, each RMP displays the same kind of information in VHF3 mode : - ACARS in the ACTIVE window, - a frequency in the stand-by window. Note : VOICE DATA SELECT can be grounded or open. When this discrete is in open circuit, each RMP displays the same kind of information in VHF3 mode : - the same frequency in the ACTIVE display, - ACARS in the stand-by display. Note : VOICE DATA SELECT is in open circuit. The PORT SELECT discrete can be changed automatically or manually by the ACARS MU or manually by one RMP. Each time the TRANSFER KEY in one RMP is selected, the REMOTE VOICE/ DATA SELECT discrete status will change momentarily forcing the ACARS MU to change the PORT SELECT and VOICE SELECT discrete status. As a consequence, VHF3 changes from VOICE to DATA or DATA to VOICE mode and RMPs will switch the display between ACTIVE and STAND-BY windows. Note: If the ACARS is active on VHF 3 and the MU fails, the active window displays dashes. When a transfer is made, ACARS is displayed in the standby window.

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Activation of ACARS on RMP

Deactivation of ACARS on RMP

temporarily displayed if continuously --> ACARS Fault

Figure 52

ACARS Activation on RMP

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DATA LOADING General For correct operation the ACARS MU needs the correct: Operating Software (SW P/N) Customer Database (DB P/N) Aircraft registration (A/C REG) The aircraft registration is received from the CFDIU in case of a ”Cold Start”. A Cold Start is activated in case of: MU removal. reset via MCDU (see maintenance pages). new software load. activation of test switch on the MU frontface. During a COLD Start the ACARS MU is completly new initialized (with A/C registration initialisation) and a self test is activated. Software Loading The software have do be loaded via a portable Dataloader direct on the frontface of the MU (not via the installed Airborne Data Loader). After operating software or database loading the correct SW P/N or DB P/N have to be checked on the MCDU (see maintenance pages - part numbers).

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TEST PORTABLE

ADAPTER CABLE (for ACARS loading only)

DATA LOADER

ACARS MU

Figure 53

ACARS Data Loading

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Label 270 Chart

MESSAGES ECAM Messages (Memo) The Label 270 messages are shown on the upper ECAM display (E/W).

Message Text ECAM

Trigger Condition

Reset Condition

ACARS CALL

An ACARS CALL uplink has been received (Voice request from ground)

Activate the ACARS MCDU Operation and select ” ACARS REC MSG ” Pg.

ACARS MSG

An ACARS UPLINK has benn received (e.g. Telex, ATIS, Loadsheet, Weather Data)

Activate the ACARS MCDU Operation and select ” ACARS REC MSG ” Pg.

VHF3 VOICE

VHF 3 set in VOICE mode

Use TFR switch on RMP to select ACARS mode

ACARS STBY

ACARS link not possible. Out of ground station range

Check VHF 3, Frequency, Service Provider avalibility

Indications in the cockpit Situation

Indication on MCDU

Indication on ECAM

Indication on RMP

No link to ground station

NO COMM, MSG NOT GEN

ACARS STBY

ACARS in the ”ACTIVE” WINDOW

ACARS not controling VHF 3

VOICE MODE, MSG NOT GEN

VHF3 VOICE

ACARS in the ”STANDBY” WINDOW

ACARS MU fault

ACARS not visable on MENU page

COM - ACARS FAULT

----- in the ”ACTIVE” WINDOW

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MCDU Messages The MCDU Scratch Pad (SP) shows messages in the following priority: 4. user entries 5. ERROR / ADVISORY messages

The ACARS ERROR / ADVISORY messages (priority 2) are shown, when the ACARS system is selected on the MCDU. Most of them can be cleared by pressing the CLR-key on the MCDU ( see chart).

ACARS Error / Advisory Messages Chart Message Text

Trigger Condition

Reset Condition

NO COMM, MSG NOT GEN ( white )

LSK that initiates a downlink is pressed while MU is in a NO COMM condition.

5 seconds or ” CLR ” key pressed or data entry

VOICE MODE, MSG NOT GEN ( white )

LSK that initiates a downlink is pressed while MU is in VOICE Mode.


PRINTER FAIL ( white )

LSK that initiates a print is pressed, and the printer cannot accept a message.


Invalid data entry

CLR key pressed or valid data entry

MU has not received the A / C Registration Number from the CFDIU.

A / C Registration Number from CFDIU ( Cold Start )

LSK that initiates a downlink is pressed, and the downlink buffer is full.


AUTO / MAN FREQ MISMATCH ( white )

Manually selected VHF data frequency differs from frequency indicated by automatic frequency select logic.

Select correct frequency or ” CLR ” key or data entry

NO LAT / LON, USE MAN FREQ ( amber )

MU is not receiving latitude and longitude data from aircraft

VHF data frequency manually selected or aircraft starts broadcast data or ” CLR ” key pressed or data entry

Hardware part number is invalid (invalid format)

Hardware part number

AIRCRAFT TYPE MISMATCH ( amber )

Aircraft type pins are not set for A320/321 aircraft

Reconfigure pins and reset MU

ACRFT REGNUM DBASE FAIL ( amber )

Aircraft registration number initially received from broadcast does not match database

Install MU on proper aircraft or modify database

INVALID ENTRY ( white ) NO A / C REG, MU IN STBY ( amber ) BUFFER FULL, MSG NOT GEN ( white )

BAD H / W PART NUM ( amber )

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FRA U

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REPORTS / REQUESTS Typ

Name 1.Movement Reports(OOOI)

O P E R

Out DL

: DL of flight number, out time (off blocks) and depature airport

Trigger Automatic

RET IN : DL of flight number, return in time and station

Automatic

OFF

: DL of flight number, off time ( takeoff time ), destination and ETA

Automatic

ON

: DL of flight number, on time ( touch down time ), destination, destination runway Automatic

IN

: DL of flight number, touch down time, in time ( on block time ) and destination

Back- Up Movement Message

Automatic

DL

DL of estimated time of arrival, destination and runway

Automatic

Voice

3a.Voice Cont. Req.

DL

Voice contact requests to various addresses

Manual

Voice

3b.ACARS CALL

UL

Call request from ground

Manual

Voice

DL

Free text ( telex ) to various addresses

Manual

Voice

UL

Free text ( telex ) to pilots from ground

Manual

Voice

DL

Crew ready for ACARS messages ( e.g. release for WX, ATIS, PDC, Loadsheet uplink )

Manual

Voice

DL

Request for forecast and actual weather

Manual

Voice

UL

UL of forecast and actual weather

after req.

Voice

DL

Request for ATIS

Manual

Voice

UL

UL of ATIS

after req.

Voice

DL

Request of Pre Departure Clearance

Manual

Voice

UL

UL of Pre departure Clearance

after req.

Voice

3h.Loadsheet

UL

Ul of loadsheet

after req.

Manual Ramp

4.Refueling Report

DL

DL of supplied fuel, remaining fuel and APU fuel

Automatic

Fuel Message

DL

Request for A/C crew rotation

Manual

Voice

UL

Info about arrival position and next leg for PIC and A/C

Auto or req Voice

DL

Request for PIL

Manual

UL

UL of PIL

Auto or req Manual Ramp

DL

Request for connecting gates

Manual

UL

Departure infos for connecting flights ( Gate and time within the next 30 minutes )

Auto or req Voice

DL

Engine condition monitoring reports

Automatic

3c Free Te 3c.Free Textt 3d.Initial Req.

N

3f ATIS 3f.ATIS

3e Airport Weather 3e.Airport

3g PDC 3g.PDC

5 A/C Crew 5.A/C Cre Rotation S 6 Pa Info List 6.Pax

R V

Function

2.Progress Report

A T I O

E

Dir

7 Connecting Gates 7.Connecting 8.ACMS/AIDS

Voice Voice Print out

______________________________________________________________________________________________________________________________________________________________________________________________

Revision No : 02US/E MK 27.2.96 FRA Issue Date : 21/05/2013


Page: 96

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MANUAL action required Profile Independent DL Reports:

Profile Independent UL Reports:

Conn Gate Reqest A/C Crew Rotation Request Telex Report Voice Request ATIS Request Weather Request PDC Request

Telex Report ACARS CALL

OUT

OFF

ON Trigger

AUTOMATIC

IN

Time

First A/C movement OUT : and all doors closed

now

OFF : Ldg gear decrompressed

now

ON

: Ldg gear compressed

now

IN

: Min one door open

Last time park brake set

Profile Independent DL Reports: ETA Change Report Destination Airport Change Report Destination Runway Change Report Frequency Change Report

Figure 54

ACARS Flight Profile

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7

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MANUAL action required Profile Independent DL Reports:

Profile Independent UL Reports:

Conn Gate Reqest A/C Crew Rotation Request Telex Report Voice Request ATIS Request Weather Request PDC Request

Telex Report ACARS CALL

OUT

OFF

ON Trigger

AUTOMATIC

IN

Time

First A/C movement OUT : and all doors closed

now

OFF : Ldg gear decrompressed

now

ON

: Ldg gear compressed

now

IN

: Min one door open

Last time park brake set

Profile Independent DL Reports: ETA Change Report Destination Airport Change Report Destination Runway Change Report Frequency Change Report

Figure 54 ACARS Flight Profile ______________________________________________________________________________________________________________________________________________________________________________________________ Revision No : 02 Issue Date : 21/05/2013


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SYSTEM REPORT/TEST COM

1L

ACARS 1L

3L

4L 5L 6L

1R

TEST>

2R

1L

2L

2R

2L

3L

3R

4L 5L

1/2

CIDS 2>

1R

HF 1>

2R

3L

HF 2>

3R

4R

4L

VHF 1>

4R

5R

5L

VHF 2>

5R

6R

6L

VHF 3>

6R

3R

TROUBLE SHOOTING DATA

COM

CLASS 3

FAULTS>

PREVIOUS LEGS

2L

SYSTEM REPORT/TEST 1R

6L

LAST LEG

2/2

4R

GROUND

REPORT>

5R 6R

ACARS

ACARS 1R

1L

ATA

2R

2L

231313 3 FQIC (3QT) / ACARS MU (1RB)

3R

3L

4L

4R

4L

5L

5R

5L

6R

6L

1L 2L 3L

TEST IN PROGRESS

PRINT*

6L

SEND :->

Figure 55

CLASS

1R 2R 3R

235534 3 SDAC1 (1WV1) / ACARS MU (1RB)

4R 5R

PRINT*

SEND :->

6R

ACARS TEST VIA CMS

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OPERATION MENU DESCRIPTION Initial Menu Pages When the ACARS MU is given control of an MCDU the MU displays the ACARS MENU page. The ACARS MENU page is the root page through which all other ACARS pages may be accessed. Each page of the ACARS Main Menu can display different data bases. These pages are created upon energization of aircraft electrical network and are recreated when the flight phase changes from preflight to inflight, from inflight to postflight, and from postflight to preflight. When the flight phase changes or the aircraft electrical network is energized, then the MU updates the text and functions displayed on the ACARS Main Menu. -ACARS Requests DL of a request. As long as no data UL is received the request is displayed without carret. -ACARS Reports Data entry for DL reports. The data will be send automaticlly. -T elex Send a TELEX to a predefined or self entered (Free Telex) address. Enter text and press send pushbutton. -V oice Contact Send a voice contact request to a predefined address -Received Messages Shows all messages received by the ACARS MU -MISC (Miscellaneous Menu) Entry to various function of the ACARS System (e.g. Maintenance)

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ACARS INFLT MENU

2/2->

CREW ROT

1L

1R

ACARS INFLT MENU

CONN GATE

2L

2R

REQUEST>

2L

3R

REQUEST>

PIL

3L

4L

4L

CONTACT>

5L

MESSAGES>
6R

ATIS

1L

3L 2L 4L 5L 6L

3L

4L 5L 6L

CONTACT>

5R

RECEIVED

MESSAGES>

6R

Air to Ground

ACARS PREFLT MENU 2/2-> PIL

2L

4R

VOICE

ON to IN Status

Ground to Air

2R 3R

5R

RECEIVED

6L

1L

1R

VOICE

6L

WEATHER

4R

3L

5L

1/2-> ATIS

1L

ACARS POSTFLT MENU 1R

1L

ACARS PREFLT MENU 1/2-> INITIAL

MESSAGES>
2R

ATIS

REQUEST>

1R

3R WEATHER

REQUEST> 4R

2R

3L

3R

4L

5R 4R 6R VOICE

CONTACT> RECEIVED

MESSAGES>

2L

5R

5L 6L

6R

Figure 56

2/2->

CREW ROT

1R

ACARS POSTFLT MENU 2R

1L
REQUEST> WEATHER 3R

2L

REQUEST> 4R

3L

VOICE

CONTACT>

4L

RECEIVED

5L 6L

1/2-> ATIS

MESSAGES>
2R 3R

5R 6R

1R

4R

VOICE

CONTACT> RECEIVED

MESSAGES>

5R 6R

ACARS Menus

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ACARS Telex Address Screen The ACARS TELEX ADDRESS is selected by pressing LSK 5L from the ACARS main menu. The hardcoded data entry screen creates a multi-page screen that provides address selection. The telex downlinks can contain either a 4 or 7 character ground address. Line select keys 1 through 5 left and right are used to select the character address to be used with the telex downlink. The action of pressing a line select key on the ACARS Telex Address screen causes the MU to select an address and generate the ACARS Telex screen, 4 character address. Line select key 6R is used to select the ACARS Telex screen, 7 character address (SITA Address). Line select key 1L on the ACARS Telex screen is used to manually enter an address. The validation criteria is exactly 4 alpha characters for the 4 character address and 7 alphanumeric characters for the 7 character address. The crew can reset an address to its default value by performing the following actions: Press the ”CLR” key so that string ”CLR” appears in the scratchpad. Press the LSK adjacent to the address to be reset to the default value. The entry of free text is performed on the ACARS telex screen. Four lines of text, of 24 characters each, may be entered on the first page and four more lines of TEXT may be entered on the second page. Text may be entered two ways. One way is to key the data into scratch pad then press a line select key in order to transfer the data into that line. If the data contains less than 24 characters, then the remainder of the line is filled with spaces. The other method is to key in the data without stopping. When the scratch pad is full, the data is automatically transferred, when the 25th character is entered, to the uppermost free data line. Character 25 is then displayed in column 1 of the scratch pad. When the data is automatically entered into line 4 of the first page, the MU will switch to the second page. The user may switch between the two methods at any time, except for the last line, which requires using line select key 5L or 5R to enter the data. If the upper-most free line of text is on page 1 and page 2 is displayed on the MCDU, and the 25th character is keyed into the scratchpad, then the data is automatically entered into the upper-most free line, the MU switches to page 1 of the ACARS Telex screen and character 25 is displayed in column 1 of the scratchpad. When the telex message is ready to be transmitted then press line select key 1R, ”SEND*”, on either page of the ACARS Telex screen. The MU creates a downlink containing the entered data, reinitializes the text parameters to their

default values of dashes, and switches to the ACARS Telex Address Select screen. ACARS REFUELING REPORT Page The ACARS REFUELING REPORT page is accessed by pushing the line key adjacent to the REFUELING REPORT indication on the ACARS MENU page. It is used to enter data for the calculation of fuel data for billing purposes only and displaying the results. QTY BEFORE: The MU will display the remaining fuel quantity, in metric tons, adjacent to line key 1R. The crew can enter a remaining fuel quantity value. Note: This value is erased when a MU Reset is performed.

______________________________________________________________________________________________________________________________________________________________________________________________



Page: 102

Page 421

MNG TECHNIC



ACARS REFUELING REPORT 1L 2L 3L 4L 5L 6L

ACARS TELEX 1L

TROUBLE SHOOTING

EDDF

SUPPLY VOL

(

5L

RETURN TO

PRINT*

(

)

3R

(

)

PRINT*

ACARS TELEX 1L

1/2-> SEND*

5R 6R

1R

2L

(

)

2R

3L

(

)

3R

4L

(

)

4R

5L

(

)

5R

6L

RETURN TO

PRINT*

REQUEST>

LOAD DATA

4R

2R 3R

1R

WEATHER

REFUELING

6R 6L

ATIS

REQUEST>

VOICE

CONTACT> RECEIVED

MESSAGES>

5R 6R

ACARS TELEX ADDRESS

4R

OPS

1L

2L

3L

STATION

4L

INITIAL

5R

JETA1

1R

RETURN TO

4R

----.-

FUEL TYP

3L

3L

FUEL DIFF

( )

2R

6L

3R

6.1

SUPPLIER

1/2-> SEND*

)

2L

FOB

0.889

)

(

2R

----.-

DENSITY

(

5L

1L

SUPPLIED

LT

ACARS PREFLT MENU 1/2->

1R

---- / 6.1

UNITS

2L

4L

TYP/QTY BEFORE

)

CREW TROUBLE

4L

5L

6L

RETURN TO

STATION

DEST> STATION

ORIGIN> MAINT

DEST> MAINT

ORIGIN>

1R 2R 3R 4R 5R

SITA ADDR>

6R

6R

Figure 57

ACARS Telex and Refuling Pages

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MNG TECHNIC



Received Messages Page The ACARS REC MSGS page is accessed by pushing the LSK 6R on the ACARS Main Menu. This page is used to view REPORT uplink titles, and to select REPORT uplinks. Each REPORT uplink has its Title, or Default Title, displayed on the Prompt Lines of LSKs 1L through 5L. Each REPORT uplink is queued for display on the ACARS REC MSGS page in the order in which it was received. A REPORT will be identified by its associated title. The most recent REPORT Title appears at the top of the ACARS REC MSGS page. Up to five REPORT uplinks may be displayed on a single page of this screen. A maximum of ten pages are allowed for a total display capability of fifty REPORT uplinks. The Message Display pages (ACARS MSG DISPLAY prompt) are selected by pushing the LSK on the Received Messages Screen associated with a specific Report uplink. Each REPORT Title is initially displayed in large font upon reception. After the crew selects a report or after the report is automatically printed, then the associated Title changes to small font. The transition to small font does not occur until the pilot returns to the Received Messages page. Upon subsequent returns to the Message Display page the Title is displayed in small font. A report may be deleted from the Message Display page by initiating the following sequence: push the CLR function key, the string CLR appears in the scratpad push the LSK adjacent to the RETURN TO REC MSGS indication. The results are the same as if the report was deleted from the ”ACARS REC MSGS” screen. For all the Message Display prompts (i.e., ACCEPT, REJECT, PRINT) if the star is absent then LSK presses for that prompt will be ignored.

Voice Contact The VOICE CONT REQUEST page is accessed by pushing the line key adjacent to VOICE CONTACT indication on the ACARS MENU page. The hardcoded data entry screen creates a multi-page screen that provides address selection. The voice request downlinks contain a 4 character ground address. Line select keys 1 through 5 left and right are used to select the character address to be used with the voice request downlink. The action of pressing a line select key on the VOICE CONTACT screen causes the MU to select an address and generate the ACARS VOICE CONTACT REQ screen. ON VHF: Voice Contact Request is displayed adjacent to the line key 1L. The default value is cyan brackets. The crew can enter the six-digit frequency prior to initiating a downlink. On HF: HF frequency is displayed adjacent to the line key 2L. The default value is cyan brackets. The crew can enter the four or five-digit frequency prior to initiating a downlink. SEND: This key will attempt to downlink the voice report via the downlink media defined by the data base.

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ACARS PREFLT MENU 1/2->
2L

3L

4L 5L 6L

ATIS

INITIAL

1L

REQUEST> WEATHER

FPL DATA

REQUEST>

REFUELING

4R

CONTACT> RECEIVED

2R 3R

LOAD DATA

1R

MESSAGES>

5R 6R

1L 2L

ACARS VOICE CONTACT

ACARS REC MSGS <-PRINTER UPLINK

1R

1L

2R

<-TELEX FRA MAINTENANCE

3L

2L

3R

3L

OPS

STATION

CREW

STATION

<-CONTROL <-SCHEDUL TROUBLE

<-SHOOTING

DEST-> ORIGIN->

1R 2R

MAINT

DEST->

3R

MAINT 4L 5L 6L

-RETURN - TO

--------------------

TELX FRA MAINTENANCE 1L 2L

HALLO ANDY DEIN TELEX KAM AN MFG KUBENS FRA MS

1L

2R

2L

3R

3L

4L

4R

4L

5L

5R

5L

6R

6L

RETURN TO

PRINT*

Figure 58

4L

<-ST ATION

5R

5L

<-DISP ATCH

5R

6R

6L

RETURN TO

6R

4R

ACARS VOICE CONTACT REQ 1R

3L

6L

ORIGIN->

4R

TROUBLE SHOOTING

1R

EDDF

2R

ON VHF ( . ) ON HF ( . )

3R 4R 5R

RETURN TO

SEND*

6R

ACARS Voice Contact and REC MSG Pages

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MNG TECHNIC



ACARS MISC Page OUT, OFF, ON, IN STATUS Pages The ACARS OOOI STATUS 1/4 page is accessed by pushing the line key 3L of the ACARS MISC menu screen. These are four ACARS OOOI status pages. On page 1/4, OOOI states not yet encountered will have time values of white dashes. When the aircraft transitions to one of these states, the time will be inserted as hhmm. The current state will be indicated with an arrow in column 12 pointing to the appropriate state name. Absence of the arrow indicates the current state is INIT. FLT NO: This field displays the flight number. RETURN: This field displays time when the aircraft returns just after an OUT event OUT, OFF, ON, IN: These fields display the OUT, OFF, ON, IN times. DEPT/DEST: This field displays the departure and destination stations. UTC: This field displays the universal coordinated time. BLOCK: This field displays the aircraft time from the OUT event to the IN event. The block time is computed by subtracting the OUT time from the IN time. Whenever a new OUT or IN time is posted, this value is updated. FLIGHT: This field displays the time that aircraft has been airborne during one flight. The flight time is computed by subtracting the OFF time from the ON time. Whenever a new OFF or ON time is posted this value is updated. Pages 2/4, 3/4 and 4/4 display the values of the inputs used to determine the OOOI state and the time of the last change in value. Pages 3/4 and 4/4 display the individual door discretes.The door status on page 2/4 represents the output of the door logic. Door status is displayed in a data field. When one of the doors is open, then the OPEN indication is displayed, and when all the doors are closed then the CLOSED indication is displayed. Slide status is displayed in a data field. When the slide is armed, then ARMED is displayed, and when the slide is not armed, then UNARMD is displayed. Parking brake status is displayed in a data field. When the parking brake is set, then the SET indication is displayed, and when the parking brake is released, then the REL indication is displayed. Aircraft movement status is displayed in a data field. When the aircraft movement is detected, then the MOVE indication is displayed, and when no aircraft movement is detected, then the STABLE indication is displayed.

The flight phase is displayed in a data field. The current OOOI state is displayed in a data field with OUT, OFF, ON, IN, RET IN, INIT, and HOLD. Each entry has a corresponding time tag. Actuation of line key 6R (PRINT prompt) on any page will attempt to print only that page of the ACARS OOOI STATUS screen on the cockpit printer. Actuation of line key 6L (RETURN TO ACARS MENU prompt) will return the user to the ACARS MENU page.

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Page 425

MNG TECHNIC



1L

ACARS MISC DATA
1R

VHF

2L

OOOI 3L

STATISTICS>

2R

SATELLITE

3R

STATISTICS>

4L

4R

PARAMETER 5L

DISPLAY>

5R

MAINT>

6R

RETURN TO 6L

ACARS OOOI STATUS 1/3 1L 2L 3L 4L 5L 6L

FMC FLIGHT NO DLH437 RETURN IN ----Z OUT 0159Z
UTC

DEPT/DEST KDFW/EDDF DATE 10 OCT 95 OFF 0215Z ON 1156Z FLIGHT 0941

PRINT*

ACARS OOOI STATUS 2/3 1R

1L

2R

2L

3R

3L

4R

4L

5R

5L

6R

6L

DOORS SLIDES PARK BRAKE A/C MOVE FLIGHT PH 1 STRUT ENG PWR OOOI ST RETURN TO

OPEN UNARMED REL -------

IN

Figure 59

ACARS OOOI STATUS 3/3

120254 120138 120200 171407 120633

1R

1L

2R

2L

3R

3L

180123 173403 120119

4R

4L

5R

5L

6R

6L

PRINT*

DOOR INPUTS AFT AV DR OPEN FWD AV DR OPEN L FWD PSG OPEN L AFT PSG OPEN R FWD PSG OPEN OPEN R AFT PSG CLOSED FWD CARGO OPEN AFT CARGO OPEN BULK CARGO RETURN TO

162745 131132 133753 055246 131132 133916 194601 162745 131132

1R

PRINT*

6R

2R 3R 4R 5R

ACARS OOOI Pages

______________________________________________________________________________________________________________________________________________________________________________________________



Page 426

MNG TECHNIC



VHF /SAT STATISTICS Pages The ACARS VHF STATISTICS page is accessed by pushing the LSK 2R of the miscellaneous menu page. This page displays statistics concerning transmissions and receptions of the VHF used by the ACARS. When LSK 6R is pushed (PRINT prompt), then the MU attempts to print the ACARS VHF STATISTICS page on the cockpit printer. Actuation of LSK 6L (RETURN TO ACARS MENU prompt) returns the user to the ACARS Main menu. The ACARS SAT STATISTICS page is accessed by pushing the LSK 3R of the Miscellaneous menu page. This screen displays statistics concerning transmissions and receptions of the SATCOM used by the ACARS. When LSK 6R is pushed (PRINT prompt), then the MU attempts to print the ACARS SAT STATISTICS page on the cockpit printer. Actuation of LSK 6L (RETURN TO ACARS MENU prompt) returns the user to the ACARS Main menu.

PARAMETER PAGE The ACARS PARAMETER DISPLAY page is accessed by pushing the LSK 5R of the Miscellaneous menu page. This page allows the user to display the value of any parameter in the parameter table.The user selects the parameter by entering a three digit number, representing the parameter index, into the scratchpad and pushing LSK 1L. If the index entered by the user is non-numeric (greater than three digits in length) or exists outside of the range of valid table indexes (000-255), the entry will be discarded and the INVALID ENTRY indication will be displayed in the scratchpad. If the index is valid, the contents of the parameter will be displayed on the page. Actuation of LSK 6R (PRINT prompt) attempts to print the ACARS PARAMETER DISPLAY page on the cockpit printer. Actuation of LSK 6L (RETURN TO ACARS MENU prompt) returns the user to ACARS Main menu. Example of parameter number 008 Aircraft Registration 011 FMC Flight Number 036 Fuel on board 122 UTC 144 Parking brake 219 Company route

______________________________________________________________________________________________________________________________________________________________________________________________



Page 427

MNG TECHNIC


ACARS MISC
1L

1R

VHF

2L

OOOI 3L

STATISTICS> SATELLITE

STATISTICS>

4L

2R 3R 4R

PARAMETER 5L

DISPLAY>

5R

MAINT>

6R

RETURN TO 6L

ACARS PARAMETER DISPLAY 1L

PARAMETER NO.

ACARS SAT STATISICS 1R

1L

2R

2L

3L

3R

3L

4L

4R

4L

5R

5L

2L

PARM DATA = --

5L

RETURN TO 6L

PRINT*

6R

6L

Figure 60

OUT RX 0 TX 0 NAKS RX DUP TX

OFF ON IN 0 0 0 0 0 0 0 NAKS TX 0 0 INC RX 0

NUMBER OF RETRIES 0 1 2 0 0 0 RETURN TO

PRINT*

ACARS VHF STATISTICS 1R

1L

2R

2L

3R

3L

4R

4L

5R

5L

6R

6L

RX TX NAKS DUP

OUT 0 3 RX TX

OFF ON IN 27 0 2 52 4 3 2 NAKS TX 8 1 INC RX 3

NUMBER OF RETRIES 0 1 2 34 1 0

2R 3R 4R 5R

RETURN TO

1R

PRINT*

6R

ACARS Statistics and Parameter Pages

______________________________________________________________________________________________________________________________________________________________________________________________



Page 428

MNG TECHNIC



ACARS Frequenz The ACARS DATA FREQ page is accessed by pushing the line key 1L of the ACARS MISC page. This page presents a menu of the regions of the world that can be selected by the crew. Each region is associated with a VHF data frequency. The associated data frequency is displayed on the second line of this page. The first character indicates the following modes: A for Automatic Frequency Management mode M for Manual mode (frequency selected by the Pilot) R for Remote mode (frequency selected by the Airport) S for Automatic search mode (scanning) D for Automatic search mode for data. The left arrow or right arrow will point to the current service provider region determined by the aircraft position data. Action on line key 6L will return the user to the ACARS Main Menu. Automatic frequency management mode When line key 6R is pushed, the MU sets the automatic frequency management mode and blanks the star adjacent to line key 6R. The data frequency displayed on the second line of this page will reflect the frequency determined by the MU automatic frequency management mode. Manual mode The manual frequency management mode is entered by selecting a line key with a frequency defined. The star adjacent to the name of the frequency in use is blanked in order to indicate the last selection mode. While in Manual mode, if the selected frequency is different from the frequency determined by the aircraft position data, the MU will display the AUTO/MAN FREQ MISMATCH indication in the MCDU scratchpad. The arrow arrow will be displayed next to the line key selected by the automatic frequency management. While in manual mode, the arrow will not be displayed if the MU is not receiving aircraft position.

______________________________________________________________________________________________________________________________________________________________________________________________



Page 429

MNG TECHNIC



ACARS PREFLT MENU 1/2-> 1L

ACARS MISC
1L

1R

VHF

STATISTICS>

2L

OOOI

SATELLITE

3L

STATISTICS>

4L

INITIAL

2L

3L

REQUEST>

5L

2R 3R

LOAD DATA

1R

WEATHER

REFUELING

4L

6L

ATIS

REQUEST>

4R

VOICE

CONTACT> RECEIVED

MESSAGES>

5R 6R

2R 3R 4R

PARAMETER 5L

DISPLAY>

5R

MAINT>

6R

RETURN TO

6L

ACARS DATA FREQ 1L

M131.725

1R

EUR/OTHERS <-

2L

*ASIA/AUS

3L

*JAPAN

N-AMERICA*

2R 3R

4L

4R

5L

5R

RETURN TO 6L

AUTOMATIC*

6R

Figure 61

ACARS Data Frequency Page

______________________________________________________________________________________________________________________________________________________________________________________________



Page 430

MNG TECHNIC



ACARS MAINTENANCE The ACARS MAINTENANCE page is accessed by pushing the line key 6R of the ACARS MISC page. The user can get access to other maintenance pages through the following MCDU line keys: 1L: PART NUMBER prompt on ACARS PART NUMBER page 2L: STATUS prompt on ACARS STATUS page 3L: TEST prompt on ACARS TEST page 4L: COMM prompt on ACARS COMM Status page 6R: hidden prompt on ACARS DEBUG page Actuation of line key 6L (RETURN TO ACARS MENU prompt) will return the user to the ACARS MENU page. PART NUMBERS The ACARS PART NUMBER page can be accessed via the ACARS MAINTENANCE page.The ACARS PART NUMBER page displays the following data: MU P/N: ACARS MU hardware part number MU S/N: MU serial number CORE SW P/N REV: Core software part number APP SW P/N REV: Application software part number DB P/N: Data base part number. Actuation of line key 6R (PRINT prompt) will attempt to print the ACARS PART NUMBER page on the cockpit printer. Actuation of line key 6L (RETURN TO MAINT MENU prompt) will return the user to the ACARS MAINTENANCE page.

______________________________________________________________________________________________________________________________________________________________________________________________



Page 431

MNG TECHNIC



1L

ACARS MISC DATA
1R

VHF

2L

OOOI 3L

STATISTICS> SATELLITE

STATISTICS>

4L

2R 3R 4R

PARAMETER 5L

DISPLAY>

5R

MAINT>

6R

RETURN TO 6L

1L

ACARS MAINTENANCE
1R

2L

2R

3L

3R

4L

4R

5L

5R

RETURN TO 6L

6R

ACARS PART NUMBERS 1L 2L 3L 4L 5L 6L

MU HW P/N 965-0728-003 CORE SW P/N 998-1383-501.A APP SW P/N REV 998-1686-501 DB P/N 998-1647-503 DISKETTE P/N REV 963-0005-002.A RETURN TO

Figure 62

MU S/N 0384

1R 2R 3R 4R 5R

PRINT*

6R

ACARS Partnumber Page

______________________________________________________________________________________________________________________________________________________________________________________________



Page 432

MNG TECHNIC



ACARS TEST The ACARS TEST menu is selected via the ACARS MAINTENANCE menu. The first page of the ACARS TEST menu enables the user to exercise basic features of the ACARS system. The second page of the ACARS TEST menu enables the user to perform a loop back test on LRUs. Test functions are activated by pushing the MCDU keys. Actuation of LSK 6L (RETURN TO MAINT MENU prompt) returns the user to the ACARS MAINTENANCE menu. Note: The LRU names displayed on the second page are dependent upon which LRUs are installed. VHF Link Test The LINK TEST function attempts to downlink a message to the ground network. Success of the test is determined by whether or not the downlink is acknowledged by the service provider. The data field displayed alongside the LSK indicates the status/results of the test, initially displaying the INITIATE indication. When the function is first selected, the status changes to TEST to indicate that the test is active. The asterisk (*) alongside the LSK also changes to a blank. Attempts to initiate the test while the asterisk is missing are ignored. If the downlink is acknowledged by the ground station, the status will change to PASS and the asterisk is shown again. If not, the status changes to FAIL and the asterisk is shown again. Five seconds after completing the test, the status changes to INITIATE.

MCDU Test The MCDU test function causes the MU to display the ACARS MCDU SCRN TEST page. From this page, the user may select LSK 6L (RETURN TO TEST MENU prompt) to return to the ACARS TEST menu, or select LSK 6R (PRINT prompt) to print the page. Satellite Link Test The SAT LINK test function is disabled because the SDU is not installed, so the MU displays the NO SDU indication and the asterisk is blanked. SDU Test The SDU test function is disabled because the SDU is not installed, so the MU displays the NO SDU indication and the asterisk is blanked. RAM Test The RAM test performs a simple write/read test over portions of RAM. All data stored in RAM is saved. If no errors are detected, then the status field displays the PASS indication for 5 seconds. If errors are detected, then the status field displays the FAIL indication for 5 seconds. While the test is performed, the status field displays the TEST indication.

Printer Test The PRINTER TEST function allows to print all characters on the cockpit printer. The data field displayed alongside the LSK indicates the status/results of the test, initially displaying INITIATE indication. When the function is first selected, the status changes to TEST to indicate that the test is active. The asterisk (*) alongside the LSK also changes to a blank. Attempts to initiate the test while the asterisk is missing are ignored. If the message containing the test pattern is determined to be undeliverable, the status changes to FAIL and the asterisk is shown again. If the data transfer is successful, the status changes to PASS and the asterisk. Five seconds after completing the test, the status changes to INITIATE.

______________________________________________________________________________________________________________________________________________________________________________________________



Page 433

MNG TECHNIC



LRU X Test The LRU X TEST function, on page 2/2 of the ACARS TEST page, performs a simple loop back analysis to evaluate the status of the displayed LRU. The data field alongside the prompt LSK displays initially the INITIATE indication. Actuation of the function changes the field to TEST, and the asterisk alongside the LSK changes to a blank. Attempts to initiate the test while a file transfer is in progress results in displaying the BUSY indication while a file transfer is in progress, then a return to INITIATE. If the TEST is correct, then the MU changes the status field to PASS and displays the prompt asterisk. Any other results causes the MU to display the FAIL indication. Five seconds after completing the test, the status changes to INITIATE. If a LRU specified for file transfer is not installed, then the MU: - blanks the star (”*”) - displays NO LRU instead of INITIATE - ignore actions on keys. If the LRU is not testable, then the MU: - blanks the star (”*”) - displays the NO TEST instead of the INITIATE indication. - ignore actions on keys. If the LRU activity logic indicates that the displayed LRU is inactive, then the MU: - blanks the star (”*”) - displays the INACTIVE indication instead of the INITIATE indication. - ignore actions on keys. When the LRU activity logic indicates that the displayed LRU is active, then the MU displays: - the star (”*”) - the INITIATE indication.

ACARS TEST VHF LINK

1L

*INITIATE

2L

*INITIATE

4L

NO SDU SDU

PRINTER

3L

NO SDU MCDU TEST* RAM TEST

RETURN TO

1L

NO TEST

3L

NO TEST CABIN 1

NO LRU

4L

2R 3R

5R 6R

2/2

FMC ACMS

2L

INITIATE*

ACARS TEST

1R

4R

5L 6L

1/2

SAT LINK

1R

CFDIU

NO TEST

2R

CABIN 2

NO LRU

3R 4R

5L

5R

RETURN TO 6L

Figure 63

6R

ACARS Test Pages

______________________________________________________________________________________________________________________________________________________________________________________________



Page 434

MNG TECHNIC



ACARS STATUS Page The ACARS STATUS menu is accessed via the ACARS MAINTENANCE menu. Actuation of line key 6L (RETURN TO MAINT MENU prompt) will return the user to the ACARS MAINTENANCE MENU page. ACARS ERROR LOG Page The ACARS ERROR LOG page is accessed via the ACARS STATUS menu. The ACARS ERROR LOG page allows access to submenus which provide detailed information about faults detected by the MU. ACARS CLASS 1 AND 2 FAULTS Page Actuation of line key 1L (CLASS 1 AND 2 FAULTS prompt) will display the CLASS 1 AND 2 FAULTS page, on the MCDU. Access to this page will be prevented if no faults exist for that category. ACARS CLASS 3 FAULTS Page Actuation of line key 2L (CLASS 3 FAULTS prompt) will display the CLASS 3 FAULTS page on the MCDU. Access to this page will be prevented if no faults exist for that category. ACARS GROUND FAULTS Page Actuation of line key 3L (GROUND FAULTS prompt) will display the GROUND FAULTS page on the MCDU. Access to this page will be prevented if no faults exist for that category. The data fields adjacent to each prompt indicate the number of fault entries residing in memory for that category. Actuation of line key 6R (PRINT prompt) will attempt to print the ACARS ERROR LOG screen on the cockpit printer. Actuation of line key 6L (RETURN TO STATUS MENU prompt) will return the user to the ACARS STATUS menu. Anomalies given by the CLASS 1 and 2 FAULTS menu and those given by the CLASS 3 FAULTS menu are detected and recorded by the software while the aircraft is in flight. Anomalies given by the GROUND FAULTS menu are detected and recorded by the software while the aircraft is on the ground. The operation of all FAULTS Menus is the same. The most recent error is displayed as the first page. Actuation of the next page function key and down-arrow function key will allow the user to advance to less recent entry pages.

______________________________________________________________________________________________________________________________________________________________________________________________



Page 435

MNG TECHNIC



1L

ACARS MAINTENANCE
1R

2L

2R

3L

3R

4L

4R

5L

5R

RETURN TO 6L

1L

ACARS STATUS
1R

2L

2R

3L

3R

4L

4R

5L

5R

6R

RETURN TO 6L

ACARS ERROR LOG 1L 2L 3L

CLASS 1 AND 2

1

2L

0

3R

3L

4R

RETURN TO 6L

1R 2R

5L

PRINT*

6R

ACARS ERROR LOG 1L

2

4L

4L

5R

5L

6R

6L

CLASS 1 AND 2 FAULTS A/ REG=. D-AIRA LEG DATE UTC ATA CLS 011 JUL08 0515 232434 1 ACARS MU (1RB) ACARS SW LOGIC ADDR 451C : 00FA RETURN TO

Figure 64

COUNT 1

PRINT*

1R

ACARS ERROR LOG

1/2

CLASS 3 FAULTS A/ REG=. D-AIRA DATE UTC ATA JUL13 0713 232434

CLS 3

1L

1R

2R

2L

LEG 002

3R

3L

FQIC (3QT) /ACARS MU (1RB)

4R

2R 3R

FQIC BUS 4L

5R

5L

6R

6L

ADDR 58D7 : 0C8D* RETURN TO

COUNT 1

PRINT*

4R 5R 6R

ACARS Error Log Pages

______________________________________________________________________________________________________________________________________________________________________________________________



Page 436

MNG TECHNIC



RCV 429 DATAS The ACARS RCV 429 DATA page is accessed via the ACARS STATUS menu. This page allows the user to inspect specific broadcast values received from other LRUs.The following information is displayed: A/C REG. number, UTC, DATE. Data displayed is actual broadcast value. For example, if UTC data is ”124500”, then that will be displayed until a different value is received. Actuation of LSK 6L (RETURN TO STATUS MENU prompt) returns the user to ACARS STATUS menu. Actuation of LSK 6R (PRINT prompt) attempts to print the ACARS RCV 429 DATA page on the cockpit printer.

ACARS TX 429 DATA The ACARS TX 429 DATA page is accessed via the ACARS STATUS menu. The purpose of the transmit 429 data page is to display the status of the 429 words broadcast by the MU. The bits of labels 030, 172, 270 and 377 are displayed on this page. LSK 2L is used to select the 429 word to be displayed. The initial value is label 172. Every time that LSK 2L is pushed, the label changes to the next value in the sequence (030, 172, 270, 377). When label 377 is displayed and LSK 2L is pushed, then label 030 is displayed. Bits 32 to 9 inclusive of the selected label are displayed on line 9, adjacent to LSK 4L and 4R.

ACARS SEL 429 RCV PAGE Actuation of LSK 5R (SELECT prompt) causes the MU to display the ACARS SEL 429 RCV page. The purpose of this page is to allow the user to view the last value received for a 429 broadcast word specified by either the core or the application. LSK 1L is used to select the LRU. LRUs that are installed, from which 429 Broadcast data is processed, are selectable. Each time LSK 1L is pushed, the MU selects the next LRU on the list. The LRUs are listed in order of the input channel they utilize. FMGEC CMC FWC L1-1 FWC L2-1 SDAC L1-1 SDAC L2-1 SDAC L1-2 DSAC L2-2 ACMS LSK 2L is used to enter the 429 label. Any value within the 1-377 octal range may be entered. If less than 3 characters are entered, then the entered data is right-justified and zero-filled. LSK 3L is used to select the value for SDI. Every time that LSK 3L is pushed, the next value from the list is selected. The order of values in list is: XX (don’t care), 00, 01,10,11.

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1L

ACARS STATUS
1R

2L

2R

3L

3R

4L

4R

5L

5R

RETURN TO 6L

6R

ACARS RCV 429 DATA 1L

A/C REG . D-AIRA

2L

UTC BC 07 17 54 DATE BC 14 JUL 95

3L 4L

2R 3R 4R

5L 6L

1R

RETURN TO

SELCT>

5R

PRINT*

6R

ACARS TX 429 DATA

ACARS SELCT 429 RCV DATA 1L 2L 3L 4L 5L 6L

LRU FMGEC LABEL

1R

1R

LABEL

SDI XX 333222222222211111111111 210987654321098765432109 -----------------------RETURN TO

1L

PRINT*

2R

2L

3R

3L

4R

4L

5R

5L

6R

6L

*172

2R

BIT 333222222222211111111111 210987654321098765432109

3R

011000000000000000100011 RETURN TO

5R

Figure 65

PRINT*

4R

6R

ACARS RCV/TX 429 Data Pages

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ACARS LRU STATUS The ACARS LRU STATUS page is accessed via the ACARS status page. The ACARS LRU status page displays the state of the activity logic for each LRU. The name of each LRU is displayed in columns 1-10 and its status in columns 12-24. The ACTIVE indication is displayed for each active LRU and the INACTIVE indication is displayed for each inactive LRU. If an LRU is not installed, according to either the aircraft configuration broadcast words, the Application, database, or the default LRU configuration specification, then the corresponding field shows the NOT INSTALLED indication. If an LRU is installed but no 429 broadcast words are processed, the NO 429 DATA indication is displayed. If an LRU name is not supplied, then neither the name or status will be displayed. Actuation of LSK 6R (PRINT prompt) attempts to print the ACARS LRU STATUS page on the cockpit printer. Actuation of LSK 6L (RETURN TO STATUS MENU prompt) returns the user to the ACARS STATUS menu.

ACARS DESCRETES The ACARS DISCRETES pages are accessed via the ACARS STATUS page. The status of the ACARS discrete inputs and outputs are displayed and updated at 1 second intervals. The character ”1” represents an open circuit and the character ”0” represents ground on the MCDU page generated by the MU.

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1L

ACARS STATUS

1R

2L

2R

3L

3R

4L

4R

5L

5R

RETURN TO

6L

ACARS LRU STATUS 1L 2L 3L 4L

FMGC 1 PRINTER DMU SPARE FMGC 2 MCDU 1 FWC 1 SDAC 1

1/3

ACTIVE ACTIVE ACTIVE NOT INSTALLED ACTIVE ACTIVE ACTIVE ACTIVE

5L 6L

RETURN TO

PRINT*

ACARS DISCRETE 1R

1L

2R

2L

3R

3L

4R

4L

5R

5L

6R

6L

AIRCRAFT TYPE

MU FAULT ANNUN MU ACTIVE / STDBY 2L ADL SELECT CAPTAIN /1L 1ST OFFR 3L

RETURN 2L TO

4L
ACARS LRU STATUS 1L 2L 3L 4L

SPARE MCDU 2 CABIN 1 CFDIU SDU 1 MCDU 3 CABIN 1 SPARE

2/3

NOT INSTALLED ACTIVE NOT INSTALLED ACTIVE NOT INSTALLED NOT INSTALLED NOT INSTALLED NOT INSTALLED

5L 6L

RETURN TO

PRINT*

ACARS LRU STATUS 1R

1L

3/3

ACTIVE NOT INSTALLED NO 429 DATA ACTIVE

FQIS SPARE ADL ADIRS

3R

3L

3R

4R

4L

4R

5R

5L

6R

6L

4L

6L 2R

3L

Figure 66

PRINT*

3R

1R

4R

2R

5R

3R

6R

4R

ACARS DISCRETE 6R PRINT*

1R 2R 1R 3R 4R3/3

716 PUSH TO TALK 5R TP14K 2L TP05N 4LRETURN TO VHF DATA KEY LINE RM VOX / DATA SEL PRINT* TP05J6R
5R

RETURN TO

2R

5R

1L
1R

2L

1R

3L

5L

2R

0 1 1 0 1 0 1 1

PRINT*

2LTO RETURN 6L

1/5

TP14A TP14B TP14C TP14D NP09J TP02J TP06E TP03J

1L

6R

6L

2R 1R 3R

1 1 4R 1 0 5R 0 0 6R 1 0 1

PRINT*

2R 3R 4R 5R 6R

6R

ACARS LRU AND Descretes Status Pages

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ACARS COMM The ACARS COMM STATUS page is selected by pushing LSK 4L on the ACARS MAINTENANCE menu. Data field 2L contains COMM when the VHF communications link is available and NOCOMM when the link is not available. Data field 3L contains DATA when the VHF link is in data mode and VOICE when the VHF link is in voice mode. Data field VHF unsent downlinks displays the number of unsent downlinks in the VHF queue. Data field 2R contains COMM when the SAT communications link is available and NOCOMM when the link is not available. Data field SAT unsent downlinks displays the number of unsent messages in the SAT queue. Data field ROUTER unsent downlinks displays the number of unsent messages in the router queue.

ACARS COM AUDIT The VHF and satellite communication AUDIT function is controlled via the ACARS COMM AUDIT page. This page is accessed via the ACARS TEST 1/2 page. LSK 1L (VHF AUDIT prompt) toggles VHF audit on and off. LSK 2L (UPLINKS prompt) enables/disables printing of uplinks addressed to this aircraft. LSK 3L (DOWNLINKS prompt) enables/disables printing of downlinks from this aircraft. LSK 4L (UPLINK TRAFFIC prompt) enables/disables printing of uplinks to other aircraft. LSK 5L (DOWNLINK TRAFFIC prompt) enables/disables printing of downlinks from other aircraft. LSK 1R (SAT AUDIT prompt) toggles satellite channel audit on and off. LSK 2R (LABEL FILTER prompt) enables/disables filtering of uplinks and downlinks. LSK 3R (LABEL prompt) enters label to be used for label filter. Access to each audit function can be separately disabled by the application.

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1L

ACARS MAINTENANCE
ACARS TEST 1L

1R

2L

2R

2L

3L

3R

3L

4L

4R

4L

5L

6R

6L

1R

1L

2L

COMM

3L 4L 5L

(OP) SAT

NO COMM

DATA

2R

2L 3L

3R

UNSENT

DOWNLINKS

VHF ROUTER SDU 000

*INITIATE

000

000

1/2

SAT LINK

INACTIVE SDU

1R

INACTIVE

2R

MCDU TEST*

3R 4R

INITIATE*

5R 6R

ACARS COMM AUDIT

ACARS COMM STATUS VHF (OP)

PRINTER

RETURN TO

1L

*INITIATE

5L

5R

RETURN TO 6L

VHF LINK

4R

4L

5R

5L

VHF AUDIT

SAT AUDIT

OFF

OFF

1R 2R

UPLINKS

OFF DOWNLINKS

OFF

FILTER

LABEL

OFF

3R

LABEL

4R 5R

RETURN TO 6L

6L

6R

Figure 67

PRINT*

6R

ACARS COM Status Page

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ACARS MU RESET To reset the ACARS MU, do the following: Type ’SAM’ into the scratchpad and press LSK 6R The ACARS DEBUG 1/2 page is activated Goto page 2/2 (press next page) The ACARS DEBUG 2/2 page is activated Type ’RESET MU’ into the scratchpad and press LSK 6R The ACARS MU will perform a reset. Attention: During reset the QTY BEFORE value on the ACARS REFUELING REPORT page is cleared. Type in old value after MU RESET.

ACARS REFUELING REPORT Page The ACARS REFUELING REPORT page is accessed by pushing the line key adjacent to the REFUELING REPORT indication on the ACARS MENU page. It is used to enter data for the calculation of fuel data for billing purposes only and displaying the results. QTY BEFORE: The MU will display the remaining fuel quantity, in metric tons, adjacent to line key 1R. The crew can enter a remaining fuel quantity value. Note: This value is erased when a MU Reset is performed.

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1L

ACARS MAINTENANCE
2L 3L 4L

ACARS PREFLT MENU 1/2-> 1R

2R

2L

3R

3L

4L

4R

5L

5L

5R

RETURN TO 6L

ATIS

INITIAL

1L

TYPE ’SAM’ INTO THE SCRATCHPAD AND PRESS LSK 6R

6L

6R

1L

ACARS DEBUG
2L

3L

4L

5L

1/2 FMC>

1R

SPARE2>

2R

ACMS>

3R

SPARE5>

4R

CABIN1>

5R

CABIN2>

6R

2L 3L

1L

ACARS DEBUG
2/2

4L 1R

2L

2R

3L

3R

4L

WEATHER

REQUEST>

REFUELING

4R

CONTACT> RECEIVED

2R 3R

LOAD DATA

1R

MESSAGES>

5R 6R

ACARS REFUELING REPORT 1L

RETURN TO 6L

FPL DATA

REQUEST>

5L 6L

SUPPLY VOL

(

) UNITS

LT

DENSITY

0.889

SUPPLIER

( )

FUEL TYP

TYP/QTY BEFORE

---- / 6.1 SUPPLIED

----.-

FOB

6.1 FUEL DIFF

----.-

2R 3R 4R 5R

JETA1 RETURN TO

1R

PRINT*

6R

4R

5L

5R

RETURN TO

TYPE ’RESET MU’ INTO THE SCRATCHPAD AND PRESS LSK 6R

6L

6R

RESET MU

Figure 68

ACARS RESET Function

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FRONT PANEL TEST After installation, correct operation of the ACARS MU can be checked by using the following controls located on the MU frontface: PASS / FAULT indicator lights In normal operation, the lights indicate the actual status of the ACARS MU. In test mode, after a 3s indicator light test, the lights indicate the test result. 7 segment LED If a test is activated and a failure is detected , the 7 segment LED shows a fault code (see chart). Display 9. 8. 7. 6. 5. 4. 3. 2. 1. 0. 9 8 7 6 5 4 3 2 1 0

Description

TEST pushbutton When pressing the test pushbutton switch: - a ACARS MU ”Cold Start” is initiated. - the LRU PASS and FAULT lights come on for 3s approximatly to indicate correct operation. - The red FAULT light goes off to indicate, that no fault is present. If the red FAULT light remains on and the green light goes off, the system is faulty and shows a fault code on the LED. - the decimal point of the BCD display blinks to indicate correct operation

Internal Hard / Software Fail

A / C Registration missed A / C Type Mismatch Bad H / W Part Number Power Down

Display: 3 After ”Cold Start” the MU did not receive the A/C registration. 2 The A/C type defined in software is different to the pin programm information 1 the MU partnumber format is incorrect.

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TEST

ACARS MANAGEMENT UNIT

Figure 69

ACARS MU Front Panel Test

______________________________________________________________________________________________________________________________________________________________________________________________



Page 446

MNG TECHNIC


MU ACARS

Figure 70

ACARS Location 80 VU

______________________________________________________________________________________________________________________________________________________________________________________________



Page 447

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A UPPER ECAM DU

C MCDU

B

D

RMP1 ( 2/3 )

PRINTER

Figure 71

ACARS Location Cockpit

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Page 448

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23-73

CIDS

CIDS PRESENTATION For higher flexibility in changing cabin layouts , the Cabin Intercommunication Data System (CIDS) is designed to accommodate these demands without the need for complex and costly hardware changes. Most cabin systems are interfaced with one of two microprocessor controlled data busses. Digitized audio control and command signals are transmitted along the bus from a central control unit called the ’Director’. CIDS reduces these requirements: additional cable bundles, Terminal blocks, function and relay boxes, connectors. These are normally associated with the installation of optional systems and cabin re-arrangements. When you have to change the cabin layout, only the controlling software is modified. The existing PAX equipment such as loudspeakers and lighting units remain as before. This software is centrally stored in the Cabin Assignment Module (CAM) and you can modify it aboard the aircraft or in the workshop. The CAM data also determines whether certain options are available. For example you can change the appropriate data in the CAM to accompany all Passenger Address (PA) announcements with chimes. The basic CIDS provides these system functions: the passenger address, the passenger call, the passenger lighted signs, the general cabin illumination control, the cabin and flight crew interphone, the lavatory smoke warning, the escape slide bottle pressure monitoring, the door bottle pressure monitoring, the service interphone (partially integrated into the CIDS), the extended emergency lighting test, the work light test, the passenger reading lights (control and test), the temperature indication of cabin compartment zones.

the boarding music pre-recorded announcements (PRAM) There is a large number of cabin loudspeakers, lighting units, passenger lighted signs, and passenger call buttons including lamps. They are connected to a smaller number (26 or 32) of locally installed driver units, called Decoder Encoder Units (DEU). These DEUs connect to one of two data bus lines, installed along each aircraft side. A second bus system with different DEUs interfaces crew related systems and components. The director units, also connected to the busses, control the individually addressed DEUs. All other attendant control equipment, cockpit equipment and avionics compartment equipment are interfaced directly to the director. The director converts the different types of input and output signals into low level digital data. The program controls this digital data. The majority of system reconfiguration work needed for installation of options, or CIDS upgrades is reduced to software changes. A removable memory cassette, the Onboard Replaceable Module (OBRM), plugged into the front face of the director, contains the software. On major CIDS software changes the OBRM is normally replaced with a new preprogrammed unit. A second plug-in memory cassette (the CAM) fits into the programming and test panel This is installed at the forward attendant station. The CAM defines many of the system properties and all cabin layout information. Also whether chimes should accompany PA announcements and whether each loudspeaker is for attendant or passenger announcements. BITEs allow the CIDS to detect faults both in connected systems, and within the CIDS unit themselves. Optional systems such as passenger entertainment video, advanced passenger services, extended emergency lighting system testing, etc. are also provisioned for in the basic installation. Controls for the cabin systems are centrally provided, for example on the forward attendant panel. The CIDS has sufficient flexibility to accomodate extra sets of controls at other locations. Attendant handsets allow communication over the interphone system and are used for PA announcements. An integrated keypad is used to establish different types of calls and announcements. An associated Attendant Indication Panel (AIP) provides attendants with PA/Interphone dialling and calling information. It is used for displaying certain system warnings. The activation of colored fields on the Area Call Panels (ACP) give long range visual indications of the CIDS for the attendants.

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POWER SUPPLY

POWER SUPPLY FWD ATTND PANEL

D I R E C T O R 2

D I R E C T O R

O B R M PROGRAMMING AND TEST PANEL

O B R M

1 PRAM

CAM

TOP LINE MIDDLE LINE

DEUs A

DEUs B

Passenger related items: - Call - Loudspeaker - Cabin lighted signs - Cabin lights Figure 72

Cabin attendant related items: - Handset - Attendant Indication Panel - Area Call Panel , Add. Attendant Panel - Slide/Door pressure monitoring - Emergency Power Supply Unit , Drain Mast Heating CIDS Schematic

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CIDS Components The CIDS is made up of a number of principle components which connect to two identical control units. The ’active’ director 101RH and the ’hot-standby’ director 102RH. The principle components are the Onboard Replaceable Modules (OBRM) 101RH1 (102RH1). CIDS Directors For redundancy, two identical directors are provided. The director 102RH is normally in hot-standby. It must receive the same inputs and respond to them in the same way as the active director 101RH. The only exception is that its outputs are normally disabled. Each director contains an OBRM module .The director connects only indirectly to the large amount of cabin equipment, via Decoder Encoder Units (DEU). ARINC links and discrete lines connect the director to individual controls, cockpit equipment and other systems. DEU Type A DEUs type A 200RH are installed along each side of the passenger cabin. To each DEU type A 3 PSUs and PIUs may be connected. The DEUs type A connect to the directors via a top-line twisted pair data bus. For redundancy purposes, the physical form of this top-line bus are two twisted pairs along each side of the cabin. They connect alternate DEUs. This means that a break in one top-line twisted pair would disable only every other DEU type A along one side of the cabin. A resistor terminates each top-line data bus cable for cable impedance matching. Each DEU type A is identical, which allows the interchange of any DEUs type A. The DEU mount include coding switches. This gives each DEU location a different address. DEU Type B DEUs type B 300RH are installed in their DEU-mounts on both cabin sides (A320) or on both sides of the cabin centerline (A321). They are located near to the exit doors. The DEUs type B connect to attendant and safety equipment. DEUs type B connect with discrete lines to this equipment: the area call panels, the attendants handsets, the slide and door pressure sensors, the emergency power supply units, The aft attendant panel receives and transmits serial data, also it is connected to the DEU type B with discrete lines. A serial link transmits data to each AIP too. Discrete connections provide AIP power and reception of AIP BITE status.

Not all inputs and outputs are used on each DEU, however, it depends on the cabin layout. The DEUs type B connect to the directors via a middle-line twisted pair data bus. One twisted pair cable on each aircraft side or the cabin centerline connects to all DEUs type B on that side. A resistor or a bus termination unit (BTU) terminates each middle-line data bus cable for cable impedance matching. Each DEU type B is identical. Coding switches in each DEU mount are used to define a different address for each DEU B location. Forward Attendent Panel (FAP) The forward attendant panel 120RH transmits to the director via a serial link which connects to both directors in parallel. For transmission of data to the forward attendant panel, however, two separate ARINC links are provided, one from each director. Separate discrete lines from the panel connect to the power supply units of the reading lights, the attendant work lights and lavatory lights. Programming and Test Panel (PTP) The programming and test panel 110RH transmits to the director via a serial link which connects to both directors in parallel. For transmission of data to the panel, however, two separate ARINC links are provided, one from each director. The Cabin Assignment Module (CAM) 115RH plugs directly into the front face of the panel. Interaction between director and CAM is via the programming and test panel ARINC links. Prerecorded Announcment and Boarding Music (PRAM) The function of the Prerecorded Announcement and Boarding Music (PRAM) Reproducer is to play prerecorded messages. It also plays boarding music programs on a cassette tape to the passengers through the aircraft passenger address system. The PRAM is controlled by the audio module, which is a part of the Fwd Attnd panel . It is installed in the cabin at the forward attendant station. The PRAM is controlled through the Cabin Intercommunication Data System (CIDS) director to receive and transmit control data.

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MNG TECHNIC


CIDS DIRECTOR 2

CIDS DIRECTOR 1

DEU A

DEU A

DEU A

DEU A

DEU A

DEU A

DEU A

DEU A

DEU A

PASSENGER LIGHTED SIGNS CALL LOUDSPEAKERS

OBRM

CABIN LIGHTS

CONTROLS AND INTERFACES

DEU A

SERVICE INTPH JACKS COCKPIT HANDSET

MOUNT

PTP CAM

DEU A

DEU A

MOUNT AREA CALL PANELS

AREA CALL PANELS

PRAM

MID EPSU

DRAIN MAST SLIDE/DOOR PRESS SENSOR LH AND RH

SLIDE/DOOR PRESS SENSOR LH AND RH HANDSET

DEU B

Figure 73

HANDSET SLIDE/DOOR PRESSURE SENSOR LH AND RH

AREA CALL PANELS

DRAIN MAST AFT EPSU HANDSET

ADD. ATTND. PANEL DEU B

DEU B

ATTND. IND PANEL

ATTND. IND PANEL

ATTND. IND PANEL

ATTND IND PANEL

HANDSET

MID EPSU

HANDSET

A321 only

SLIDES PRESS SENSOR LH and RH MID EPSU

MID EPSU

FWD EPSU FAP

DEU B

MOUNT

MIDDLE LINE

AFT ATTND PANELS AREA CALL PANELS

ATTND. IND PANEL DEU B

CIDS Detailed Schematic

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DESCRIPTION AND INTERFACES SDCU The Smoke Detector Control Unit (SDCU) provides the directors with the lavatory smoke detection signal to activate the audio and visual warnings in the cockpit and in the passenger cabin. ECAM The two Flight Warning Computers (FWCs) and the two System Data Acquisition Concentrators (SDACs) are the main computers of the ECAM. In case of CIDS malfunctions or CIDS related systems malfunction, a message is sent to the ECAM. NOTE: the buzzer sound for the Crew Call System is generated by the FWCs. The cabin door position and slide armed information given by the SDACs is sent to the directors to control the seat row lighting and FAP indications (A321). SFCCs The Slat Flap Control Computers (SFCCs) provide signals to the directors to control the FASTEN SEAT BELT and NO SMOKING signs in automatic mode. LGCIU The Landing Gear Control Interface Units (LGCIUS) are also used by the directors to control the NO SMOKING and FASTEN SEAT BELT signs in automatic mode. A signal from the LGCIUs switches on the Service Interphone System ten seconds after landing. Passenger Address HANDSET The cockpit handset is directly connected to the directors. NOTE: cockpit mounted handset has priority over all attendant passenger address announcements and over the Passenger Entertainment System music. AMU The Audio Management Unit (AMU) is used to establish and reset cockpit interphone operation. ENGINE OIL PRESSURE SWITCH / COCKPIT DOOR SWITCH When the engines are running and the cockpit door is open, the forward left entry light goes automatically to 10% lighting intensity.

With cockpit door open, the forward attendant station loudspeaker volume will decrease by 10 dB (PA from cockpit). CABIN PRESSURE SWITCHES In case of cabin depressuration signals are sent to the CIDS directors to control the following items: Cabin lights (full bright). Exit lights (via EMLS). NO SMOKING and FASTEN SEAT BELT signs. NOTE: the RETURN TO SEAT signs are not affected. SERVICE INTERPHONE The service interphone connects the handsets to the eight service interphone plugs. The eight service interphone plugs are located around the aircraft for maintenance purposes. COCKPIT CALL PANEL The cockpit call panel provides call facilities between flight crew and attendant stations, and enables emergency calls to all attendant stations. ANN. LIGHT CONTROL BOX The control box is used to test and dim the CIDS related illuminated pushbuttons. CFDIU The CFDIU is used as an interface between the CIDS and the MCDUs, for testing and trouble shooting. NOTE: selecting CIDS on the MCDU main menu permits access to the same menu, as on the PTP. NO SMOKING / FASTEN SEAT BELTS SWITCHES The NO SMOKING and FASTEN SEAT BELTS switches are directly connected to the diretcors for manual and automatic control of the signs.

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ECAM

SDCU POWER SUPPLY

FWC 1 FWC 2

SDAC 1

SFCC 1

SDAC 2

SFCC 2

CDIS DIRECTOR 2

FWD ATTND PANEL **** * *** * ** * * ** ** ** * * * ****

TO TOP LINE BUSES

PA HANDSET

AMU

LGCIU 2 CIDS DIRECTOR 1

* * * * * * PROGRAMMING AND TEST PANEL

LGCIU 1

TO MIDDLE LINE BUSES ENGINE OIL/ CABIN PRESS SWITCHES

COCKPIT DOOR SWITCH

SERVICE INTERPHONE

COCKPIT CALL PANEL

* CAM *

ANN. LIGHT CONTROL BOX

EMLS

Figure 74

CFDIU

NS/FSB SWITCHES

PRAM

CIDS Interfaces

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MNG TECHNIC



TYPE A DECODER ENCODER UNIT General The Decoder-Encoder Unit (DEUs) are located in the left-hand and right-hand sides of the cabin ceiling. Each type A DEU connects to a CIDS top line data bus. Each data bus takes the form of a shielded twisted pair cable. Functions Each type A DEU interfaces: up to three Passenger Service Units (PSUs) two loudspeakers four flourescent strip lights which a part of the cabin light system. Top Line Data Bus Two top line data buses on each side of the passenger cabin connect the type A DEUs to the director 6 (8) of them are connected to the top line number 1 and 7(8) to the top line number 2. A resistor is located on the last DEU A mount of each line for impedance matching. A broken top line can effect no more than half ot the DEUs installed on one cabin side

Fail Safe Operation In the event of a data bus failure the DEU maintains the current status of the discrete cabin systems output for a certain time. After this delay the outputs are switched to a pre-defined fail safe state, that means the four fluorescent strip lights come on with full brightness and all other items go off. All audio inputs/outputs are immediatley switched off. Emergency Functions All DEUs operate in emergency mode when the DC service bus is no longer powered. The DEUs are then supplied from DC essential bus. The type A DEU passenger address circuits and the type B DEU interphone circuits remain operational.

Coding Switches A coding switch in each DEU mount gives each DEU a unique address. This methode enables removal, interchange and replacment of DEUs without having to consider their adress. Note: In the event of mount change it is necessary to select the same code as used before. CIDS Power Up When the CIDS is powered-up or reset, the director follows a power up routine. This includes the initialization and testing of each DEU and connected equipment. The test results are transmitted to the director which compares them with its programmed data to decide on their status. At least 95% of possible DEU failures are automaticly detected.

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Page 455

MNG TECHNIC


POWER SUPPLY CABIN LIGHT


PSU

PSU FSB

FSB

FSB

NS

NS

NS


DEU A MOUNT

FROM DIRECTOR FROM DIRECTOR 28V SERVICE 28V DC ESSENTIAL

DEU A MOUNT

PSU

J3 J2

J2

DEU A MOUNT

J1

DEU A


CODING SWITCHES DEU MOUNT TOP LINE 1

TOP LINE 2

Figure 75

DEU-A Schematic

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TYPE B DECODER ENCODER UNIT General The Decoder-Encoder Unit (DEUs) are located in the left-hand and right-hand sides of the cabin ceiling or the cabin centerline. Each type B DEU connects to a middle line data bus. Each data bus takes the form of a shielded twisted pair cable. Functions Each type B DEU interfaces the following components and systems: Aft Attendant Panel, Add. Attendant Panel Emergency Power Supply Unit Slide/Door Pressure Monitoring. Passenger Address/Interphone Handset Attendent Indication Panel Area Call Panel Drain Mast Heating Monitoring

CIDS Power Up When the CIDS is powered-up or reset, the director follows a power up routine. This includes the initialization and testing of each DEU and connected equipment. The test results are transmitted to the director which compares them with its programmed data to decide on their status. At least 95% of possible DEU failures are automaticly detected. Emergency Functions All DEUs operate in emergency mode when the DC service bus is no longer powered. The DEUs are then supplied from DC essential bus. The type A DEU passenger address circuits and the type B DEU interphone circuits remain operational.

Middle Line Data Bus One middle line data bus on each side of the passenger cabin or cabin centerline connect the 2(3) type B DEUs to the director A resistor or a bus termination unit (BTU) located in the DEU A mount (end of line) terminates each middle line data bus for impedance matching. A319/320:Two additional mounts already connected to the middle line data bus are installed near to the forward right hand door and to the left emergency exit. A321:Three additional mounts already connected to the middle line data bus are installed near to the forward right hand door and to the right and left emergency exit. Coding Switches A coding switch in the DEU mount gives each DEU a unique address. This methode enables removal, interchange and replacment of DEUs without having to consider their adress. Note: In the event of mount change it is necessary to select the same code as used before.

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SLIDES/DOORS PRESSURE

PA/INTERPHONE HANDSET

AREA CALL PANEL EMERGENCY POWER SUPPLY UNIT

ATTND INDICATING PANEL

ADD/AFT ATTND PANEL J2

J3

J3 J3

DEU B DRAIN MAST HEATING J1

MONITORING

BTU(A321)

CODING SWITCHES DEU MOUNT

MIDDLE LINE FROM DIRECTOR

GND

28V DC SERVICE 28V DC ESSENTIAL

Figure 76

DEU-B Schematic

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DECODER ENCODER UNIT MOUNT The mount for type A and type B DEUs are similar. But due to indexing pins it is not possible to install a type A DEU on a type B DEU mount. The mounts for the type A DEUs have the indexing pins on the outer and those for the type B DEUs on the inner side. On each DEU mount there is an Adress Coding Switch. In case of a mount change the old code must be selected. Note: A table giving the adress code is placed close to the mount.

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ADRESS CODING SWITCHES

INDEXING PINS

DEU A DEU B

CONNECTOR J1

Figure 77

DEU Mount

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PROGRAMMING AND TEST PANEL General The Programming and Test Panel (PTP) is located at the FWD Attendant station behind a hinged access door. For correct Cabin Intercommunication Data System (CIDS) operation, the Cabin Assignment Module (CAM) must be plugged in.

The CAM contains the cabin layouts 1, 2, 3 and M. In the basic configuration, only layout 1 is programmed to the airline request. Only layout M can be modified via the PTP.

Functions The functions of the Programming and Test Panel are as follow : To monitor the failure status of the CIDS and certain connected systems. To activate CIDS component tests and readout of the results. To examine in detail the fault data held in the director BITE memory. To program the CIDS properties and cabin layout information into the CIDS directors, which are copied from the CAM. To onboard reprogram: - CAM data, - activation of the provisioned CIDS extra functions, - change cabin layout, - implement cabin zoning. Description The PTP has an alphanumeric display with four rows of twenty characters. The display is used to present messages, test results and selection menus. There are keys at each end of the display rows. They are labelled on the display with ”< “ or ” >“ characters. There is no power supply switch. The Programming and Test Panel is automatically supplied if the DC service bus is supplied. The DISPL ON pushbutton is used to switch on the display. The display is automatically switched off if the panel is not used for 10 minutes. A keypad is provided for entry of numerical data. The Programming and Test Panel contains two pushbuttons and two annunciator lights for testing the emergency light system. The CAM defines all of the modifiable system properties and layout information for the CIDS. It contains four cabin layouts.

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SYSTEM STATUS
TEST EMER LIGHT

CONT> 1

2

3

4

5

6

BAT

BAT OK

7

8

9

SYS

SYS OK

CLR

0

.

CAM-MODULE Figure 78

Programming and Test Panel

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FORWARD ATTENDANT PANEL General The Forward Attendant Panel (FAP) is in modular form with a master circuit board and sub-panels which connected to the master board. The master board contains all parts which are common to all configurations of the Forward Attendant Panel. This includes the power supply, ARINC 429 transmitter and ARINC 429 receiver. Light Panel The cabin light panel comprises control pushbuttons for the different cabin lighting systems. There are controls for the entrance areas and the different cabin sections. In addition, power switches provide the power for the lavatory lights, attendant work lights and the passenger reading lights. All pushbuttons, except for MAIN ON and MAIN OFF, have integral lights for visual confirmation of the pushbutton activation. Audio Panel The audio panel allows centralized control over passenger entertainment, boarding music and prerecorded announcements. Water and Miscellaneous Panel The Water and Miscellaneous Panel is installed at the bottom of the Forward Attendant Panel Note: For the water and waste panel description refer to the related system. Emergency Light pushbutton EMER LIGHT is a red guarded pushbutton with a integral light which is used to switch the emergency lighting on and off. Slides Armed light (A319/321) The SLIDES ARMED light is used to indicate the slide status. If all slide are armed the light is on. If not all slides are armed, the light is flashing and if all slides are disarmed, the light is off. This signal is received from the SDAC. Doors Closed light (A319/321) The DOORS CLOSED light is used to indicate, that all cabindoors are closed. This signal is received from the SDAC.

Cabin Ready pushbutton (A319/321) Cabin Ready is a pushbutton with a integral light which is used to activate the CABIN READY indication on the ECAM when the button is switches to on. Lavatory Smoke light The LAV SMOKE light is used to warn of lavatory smoke. A command from the Smoke Detection Control Unit (SDCU) can only reset the indication when the smoke has gone. Reset pushbutton When the RESET pushbutton is depressed in case of lavatory smoke warning, it resets only aural and visual indications in the passenger compartment, without affecting the LAV SMOKE indication at the FAP. Panel Light Test pushbutton The PNL LIGHT TEST pushbutton is used to switch on all the Forward Attendant Panel lights (Bulb check). CIDS Caution illuminated pushbutton The CIDS CAUT light is activated, when a CIDS Class1 or Class 1CAB failure occurs (see Chapter Fault Isolation and Bite). The CIDS CAUT light is resetable in flight but comes on again on ground when the landing gear is down and locked. The light cannot be reset on the ground. When a CIDS CAUT occurs, the respective failure message is displayed on the Programming and Test Panel.

AFT /ADD ATTENDANT PANEL General The AFT ATTND panel is installed in the aft entrance area of the aircraft. The ADD ATTND (A321) panel is installed in the middle cabin area of the aircraft. Functions The cabin light panel comprises control pushbuttons for the different cabin lighting systems. There are controls for the different entrance area and cabin section. All pushbuttons, except for MAIN ON and MAIN OFF, have integral lights for visual confirmation of the pushbutton activation. The RESET pushbutton resets the lavatory smoke warnings (same than FAP).

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A321 only

ÈÈÈÈ ÈÈÈÈ A319 and 321 only Light Panel ÈÈÈÈÈÈÈÈÈÈÈÈÈÈÈÈÈÈÈÈÈÈÈÈÈ ÈÈÈÈÈÈÈÈÈÈÈÈÈÈÈÈÈÈÈÈÈÈÈÈÈ ÈÈÈÈÈÈÈÈÈÈÈÈÈÈÈÈÈÈÈÈÈÈÈÈÈ ÈÈÈÈÈÈÈÈÈÈÈÈÈÈÈÈÈÈÈÈÈÈÈÈÈ ÈÈÈÈÈÈÈÈÈÈÈÈÈÈÈÈÈÈÈÈÈÈÈÈÈ ÈÈÈÈÈÈÈÈÈÈÈÈÈÈÈÈÈÈÈÈÈÈÈÈÈ ÈÈÈÈÈÈÈÈÈÈÈÈÈÈÈÈÈÈÈÈÈÈÈÈÈ ÈÈÈÈÈÈÈÈÈÈÈÈÈÈÈÈÈÈÈÈÈÈÈÈÈ ÈÈÈ ÈÈÈÈÈÈÈÈÈÈÈÈÈÈÈÈÈÈÈÈÈÈÈÈÈ ÈÈÈ ÈÈÈ Add. Attendant FWD

Audio Panel

ÈÈÈÈ ÈÈÈÈ ÈÈÈÈ SLIDES DOORS ARMED CLOSED

ÈÈ ÈÈ ÈÈ CABIN READY

Figure 79

Water and Miscellaneous Panel

FWD

AFT

FWD

AFT

FWD

AFT

AFT

Panel

Aft. Attendant Panel

FWD and AFT and ADD ATTND Panel

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ATTENDANT HANDSET General Each attendant station is equipped with a handset for public announcement, interphone dailing and communication. The handset rests in a cradle. Function The handset and cradle provide the following functions: the hook off sensing the Push to Talk (PTT) switching the PA announcement the single-key (A320) or double key (A321) call-activation via integral keypad and telephone conversation. To reset any dailing procedure, press the RESET key. A320: For PA announcement, press the PA ALL key on the handset. To make a announcement use the PTT switch. For Captain call, press the key CPT on the handset. A captain call procedure with aural and visual indication in the cockpit is initiated. The telephone conversation is accomplished as soon as the called handset is hooked off. For Cabin interphone, press the key related to the station (1 L/R,3L,3R) on the handset. A attendant call procedure with aural and visual indication in the cabin is initiated. The telephone conversation is accomplished as soon as the called handset is hooked off. For Service Interphone, press the SERV INT key on the handset. If the aircraft is on ground or the SERV INT OVRD pushbutton is on, telephone conversation is accomplished with headset plugged in at any Service Interphone Jack. A319/321: For PA announcement, press the PA and the ALL key on the handset. To make a announcement use the PTT switch. Pressing the PTT button without key selection activates the ”DIRECT PA” mode of operation with PA announcement in the whole cabin. For Captain call, press the key (CPT, EMER CALL) on the handset. A captain call procedure with aural and visual indication in the cockpit is initiated. The telephone conversation is accomplished as soon as the called handset is hooked off.

For Cabin interphone, press the INTPH key and the key related to the station ( (FWD, MID, EXIT, AFT) on the handset. A attendant call procedure with aural and visual indication in the cabin is initiated. The telephone conversation is accomplished as soon as the called handset is hooked off. For Service Interphone, press the INTPH key on the handset twice. If the aircraft is on ground or the SERV INT OVRD pushbutton is on, telephone conversation is accomplished with headset plugged in at any Service Interphone Jack.

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PA ALL

A-319 (without MID and EXIT)

Figure 80

Attendant Handset

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ATTENDANT INDICATION PANEL One Attendant Indication Panel is fitted on each of the Attendant Stations. Each one comprises an alphanumeric display with 2 lines of 16 characters. It displays communication information on the upper line and cabin system information in the lower line. It is provide with a pink and a green indicator light used as attention getters. Interphone Display Example: FWD attendant calls AFT left Attendant Step 1: When the Forward Attendant handset is unhooked, a symbol appears and a dailing tone is heard. Step 2: 3 L (or AFT ATTN) the designation of the desired station is selected on the handset. A confirmation message appears on the AIP. Step 2a: The designation of the calling station is displayed on the called AIP and a green indicator light comes on steady. Step 3: When the AFT left hand handset is unhooked, the symbol disappears from the Forward AIP. Step 3a: On the AFT left AIP, the CALL indication disappears and the green indicator light goes off.

Indication on all AIPs When a Emergency Call is initiated from the cockpit, the EMERGENCY CALL indication is shown on all AIPs., A319/321: When a All Call is initiated from the cockpit, the CAPTAIN CALL indication is shown on all AIPs.,

Step 4: The AFT left attendant interphone is engaged. A busy tone is heard. To disconnect a handset from the interphone system, put the handle into the cradle or press the RESET pushbutton on the handset.

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Indication on FWD ATTND AIP

Indication on AFT L ATTND AIP

STEP 1

#

STEP 2

# 3L

STEP 2a

CAL L 1 L

STEP 3

3L

STEP 3a

1L

STEP 4

BUS Y 3 L

A319/321 : 1 = FWD ATTN 2 = MID ATTN 3 = EXIT ATTN 4 = AFT ATTN

Information displayed on all AIPs EMERGE NCY CAL L

CA PTA I N CAL L

Figure 81

Attendant Indication Panel

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Passenger Address Display Step 1: When the AFT left attendant handset is unhooked, a symbol appears and a dailing tone is heard. Step 2: When the PA ALL pushbutton is pressed on the handsets keyboard, a confirmation message appears on the AIP. Step 2a: PA ALL IN USE appears on all other AIPs. Step 3: If the PA CALL call is impossible due to the priority of a call already in progress the word BUSY appears on AIP. Priority List: 6. from the cockpit 7. any attendant station 8. the prerecorded announcement system 9. the entertainment PA sources (Boarding music).

Cabin System Displays System information is displayed on all AIPs. There are indications of: Smoke detection in lavatory

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Indication on AFT L AIP STEP 1

#

STEP 2

#

STEP 3

PA AL L

Indication on all other AIPs

STEP 2a

PA ALL IN USE

BUSY PA ALL IN USE

Cabin System Information displayed on all AIPs SMOKE L AVATORY A

Figure 82

Attendant Indication Panel

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Lufthansa Technical Traini


AREA CALL PANEL General The Area Call Panels are installed in the left and right center ceiling, at each end of the cabin and in the middle cabin area. Each panel consists of individually controlled lighted fields. The fields are used in steady or flashing mode.

Attendant Call The green field comes on at the ACP when a Attendant call is initiated from a attendant handset. Attendant calls are accompained by one low chime on the attendant loudspeakers.

Crew Call Any normal call from the cockpit to a attendant station is accompained by a pink steady field on the ACP. Whenever the EMERGENCY call is initiated either from the cockpit or the cabin, the pink field on the ACP flashes. The normal call is accompanied by one high/low chime and a EMERGENCY call by 3 high/low chimes on the attendant loudspeakers. Passenger Call A call from a passenger to the cabin attendant results in lighting the steady blue field on the ACP of that side of the forward, middle or aft section from where the call was initiated. Passenger calls are accompained by one high chime on the attendant loudspeakers. Note: Simultaneously, on the PSU, the corresponding call pushbutton comes on and the seat row numbering sign comes on steady if all the passenger doors are closed or flashing with at least one passenger door open. Lavatory Call A call from the lavatory results in lighting the amber field on the ACP allocated to the lavatory from where the call was initiated. A lavatory call is accompained by one high chime on the attendant loudspeakers. Note: Simultaneously, the lavatory call pushbutton comes on. Lavatory Smoke The amber field of the corresponding ACP will flash whenever smoke is detected in a lavatory. A smoke warning in lavatories is accompained by three low chimes on the attendant loudspeakers.

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Amber

Green Pink Blue

Figure 83

Area Call Panel

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POWER SUPPLY The CIDS is continuously energized (on ground and in flight) when the essential busbar 401PP and service busbar 601PP are energized. In normal CIDS operation, the essential busbar 401PP supplies: the active director. all the circuitry in DEUs type A 200RH which is necessary for PA operation, all the circuitry in DEUs type B 300RH which is necessary for PA and interphone operation. The service busbar 601PP supplies: the hot-standby director 102RH, the remaining (nonessential) circuitry of the DEUs, - the aft and add attendant panel 126RH and 128RH (via DEUs type B), - the area call panels 340RH (via DEUs type B), - the attendant indication panels 320RH (via DEUs type B). the programming and test panel 110RH, the forward attendant panel 120RH, When essential bus power is unavailable, circuitry in the directors and DEUs switches the respective essential circuits to the service bus. This ensures full CIDS capabilities except for emergency mode operation. When electrical power from the service bus is unavailable, this equipment is inoperative: the standby director 102RH (A320 only as long as DIR1 is operative), the nonessential DEU circuits, the programming and test panel 110RH, the forward attendant panel 120RH, The aft and add attendant panel 126RH and 128RH, the attendant indication panels 320RH and area call panels 340RH are also inoperative, because DEUs type B supply these panels (with service bus power). An emergency situation causes a different system to remove electrical power from the service bus to reduce power consumption. The CIDS director further reduces power consumption. It goes into emergency mode operation. It also disconnects the DEU A essential power as long as no PA announcements are made. PA announcements are possible when the top line cut-off relay 106RH is deenergized.

A320: If the active director 101RH becomes faulty, control of the top-line cut-off relay 106RH transfers to the standby director 102RH. The power transfer relay 105RH connects essential power to the standby director 102RH, which then takes over control of the CIDS. When the A/C loses main power (the service bus and the essential bus), and the emergency exit lights switch is set to ARM or ON, all CIDS units , which are supplied by essential bus, are switched automatically to the hot battery bus. A319/321: If the active director 101RH becomes faulty, the standby director 102RH takes over control of the top-line cut-off relay106RH via a parallel connection to the relay. When the A/C loses main power (the service bus and the essential bus), the complete CIDS system is powerloss.

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log 1 = PWR

Figure 84

CIDS Power Supply Schematic A320

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Figure 85


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log 1 = PWR

Figure 86 CIDS Power Supply Schematic A321

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401 PP

401 PP

Figure 87


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PASSENGER ADDRESS SYSTEM General The CIDS director accepts audio signals from the various PA sources in the aircraft. It assigns priorities to each source. It transmits the signal in digital form via the four top line data busses to the DEUs type A. These send it to the cabin loudspeakers for broadcasting. You can perform PA announcements from different sound sources. The director can route the broadcasts to any combination of six audio channels. A separate amplifier in each DEU type A drives each cabin loudspeaker. It is programmed on CIDS initialization for gain and assignment to one of the six audio channels. Three channels are reserved for PAX announcements. This lets you assign each cabin loudspeaker to one of up to three zones. A PA chime capability is also given. The PA system remains available during aircraft emergency conditions. It is powered down as long as no PA announcements are in progress. Volume Control When an engine is running, the PA volume is increased automatically by +6 dB. The PA volume is also increased by +6 dB in the event of cabin depressurization. A separate audio amplifier in the CIDS DEUs type A drives each loudspeaker independently. PA Announcements from Cockpit Handset The handset is mounted at the cockpit center pedestal and contains an integral PTT switch. Press the PTT switch to key the PA system. It overrides lower priority PA sources, and broadcasts the speech over all PA loudspeakers. Sidetone is heard over the handset earpiece. A PA ALL IN USE indication at all attendant indication panels in the cabin accompany announcements which use the handset. PA Announcements from other Cockpit Audio Equipment Any of these cockpit audio equipment may be selected to make a PA announcement: the boomset microphone the hand microphone (with internal PTT switch), the oxygen mask integral microphone. To select the PA system, the rectangular PA button must be pressed and held. It connects the microphone audio to the PA system. The integral PTT switch in

the hand microphone is pressed to key the PA system with the respective audio. When the boomset or oxygen mask is used, pushing the rectangular PA button and the combined PA volume control/PA sidetone switch, switches the sidetone audio to the boomset or headset earpieces. The knob adjusts the sidetone volume. You can monitor the PA sidetone at any time when you only select the volume/sidetone switch. When the PA selector switch is activated, PA ALL IN USE is displayed at all attendant indication panels. PA announcements from the selected cockpit audio equipment are broadcast over all PA loudspeakers. They immediately override PA audio from any other source. PA Announcements from Attendant Handset An attendant handset is mounted at each attendant station. An AIP is installed near to each handset for display of PA-in use information. When the handset is lifted, a dialling tone (440 Hz) is heard. The top line of the AIP displays a number symbol. When you press the ’PA ALL’ key (A320) or PA and ALL (A319/321), a confirmation message is displayed on the AIP. If the PA call is impossible due to the priority of a call already in progress, then the word BUSY appears on the AIP display. To press the keypad RESET key always clears any handset operation and lets you make a new key selection. The CAM also contains a priority list which is divided into priority levels. It starts with: the cockpit as the highest priority, the attendant stations, the prerecorded announcement system, the entertainment PA sources Once a PA ALL call is established, the bottom line of all AIPs displays the respective PA IN USE message. To press the PTT switch keys the PA system. Sidetone audio is fed to the handset earpiece. When the PA announcement is over, put the handset into the cradle or press the RESET button to disconnect the handset from the PA system. A319/321: Pressing the PTT button on handset witout any selection activates the ”Direct PA” mode or operation. PA announcments are possible now via all cabin loudspeakers

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BOOMSET OR HEADSET / HANDMIKE

OXY MASK

AMU

AUDIO CONTROL PANEL

A321

EMER

CALLS FWD

EMER MID

EXIT CALL ON

MECH

ALL

AFT

MIDDLE LINES FWD ATTND PANEL . .. . .. .. . . . .. .. .. .. .. . . . . .. .

PTT

CALLS

TOP LINES

DIRECTOR 1/2

COCKPIT HANDSET A319/320

DEU A

C HL I O MG EI C

HIGH HIGH/ LOW LOW

DEU B

.. . . .

.. .. .. .. PROGRAMMING AND TEST PANEL CAM

KEYBOARD

ANN.LIGHT TEST/DIM BOX

Figure 88

CFDIU

PA

CAPT SVCE INTPH

1L/R 3L

PRAM

A319/321

A320 PA ALL

ENG. OIL PRESS. SW. CABIN PRESS.

ATTENDANT INDICATION PANEL

PTT

3R

RESET

1

INTPH 2 MID

FWD 3

ALL

4 EXIT

AFT CAPT

RESET

PA System Schematic

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CABIN INTERPHONE Description The CIDS director accepts audio signals from the various interphone sources in the aircraft and assigns priorities to each source. The director performs telephone exchange switching and call functions from cockpit call switch settings or the attendant handset keypad entries. All this is done with reference to the parameters defined in the CAM. Chimes are transmitted via the top line data bus and the PA loudspeakers. In the cockpit, integral lamps in the call switches annunciate interphone calls. In the cabin, the ACPs and AIPs are used for annunciation. Cabin Interphone The cabin interphone system offers different communication modes. Each mode can lead to different call activations which are assigned in the CAM. All communication modes are handled with respect to the predefined priorities listed below: Emergency Call. Call from cockpit including: - an all call from cockpit,- a normal call from cockpit. call from cabin station. Additionally, all interphone sources have interrelated priorities, as assigned in the CAM. There are eight priorities which can be individually assigned to the different interphone sources. If more than one interphone source requests the same communication mode, the source with the higher priority will have precedence. If they have the same priority, the interphone source which was dialled first will be given preference. If an interphone source requests a station which is engaged in another communication link: a busy indication at the AIP appears, a busy tone via the sidetone output will be transmitted, if this interphone source has equal priority to, or lower priority than the existing link. If this interphone source has a higher priority, then following action will be executed. The existing communication link will be interrupted and the new link will be established. The audible and visual calls will be activated as assigned for this communication mode. An ”OVER” indication will be displayed at the AIP of the station.

Operation from Cockpit Each application of the cabin and flight crew interphone system starts with a dial procedure. In the cockpit, special keys (Call buttons) are available for dialling the desired interphone function. The interphone equipment (for example cockpit boomset) must be connected to the interphone system. The interphone function at the audio selector panel must be selected. Indicator lights give visual feedback, which are also activated when an attendant station calls the cockpit. A call indicator is activated at the audio selector panel when an attendant calls the cockpit. This indicator is integrated into the transmit button on the audio selector panel. It must be switched if communication is to be established after a call has arrived (or has been activated) from the cockpit. A reset key is installed at the audio selector panel. When this key is pressed, an activated call function from the cockpit is reset or an arriving call from the cabin is cancelled. A normal call from the cockpit activates a steady pink light on the associated ACP, a high/low chime on all cabin speakers and a CAPTAIN CALL indication on the associated AIP. A emergency call from the cockpit activates a flashing pink light on all ACPs, 3 high/low chime on all cabin speakers and a EMERGENCY CALL indication on all AIPs. A319/321 only: A all call from the cockpit activates a steady pink light on all ACPs, a high/low chime on all cabin speakers and a CAPTAIN CALL indication on all AIPs. Operation from Attendant Station To push one key (1L/R, 3L, 3R) of the keyboard on A320 or the INTPH and one key (FWD,MID,EXIT,AFT) on A319/321, which is integrated into the handset, initiates a call function in the aircraft. If an attendant station is called, visual indication is given. A transmitted chime via the cabin loudspeakers assigned to the attendants station and/or attendants area performs audible call signals. All attendant stations in the cabin are equipped with a reset key to reset the interphone function, and permits a new dial procedure. A normal call from the cabin activates a steady green light on the associated ACP, a high/low chime on the associated attendant speakers and a ”calling station” indication on the associated AIP.

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OXY MASK


AMU AUDIO CONTROL PANEL

ATT

DEU A TOP LINES MIDDLE LINES

DIRECTOR 1/2

DEU B

CAB

AREA CALL PANEL A319/320 A321

CALLS

EMER

CALLS FWD

EMER MID

EXIT CALL ON

MECH

ALL

AFT

C L HIGH HO I G HIGH/ M I LOW EC LOW

PROGRAMMING AND TEST PANEL CAM A320 PA ALL

KEYBOARD

AVIONICS BAY

Figure 89

CABIN

PA

CAPT SVCE INTPH

1L/R 3L

FWC COCKPIT

ATTENDANT INDICATION PANEL A319/321

3R

RESET

1

INTPH 2 MID

FWD 3

ALL

4 EXIT

AFT CAPT

RESET

Cabin Interphone Schematic

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SERVICE INTERPHONE Service Interphone The service interphone system provides the telephone communication on the ground between the flight crew and the ground service personnel. Eight service interphone jacks are installed at different locations on the aircraft. The service personnel use those to talk to each other, the cockpit and the attendant stations through handsets. The service interphone system is made up of the components listed below. eight interphone jacks, a service interphone OVRD switch, with an integral indicator light. The audio lines from the interphone jacks are routed to the amplifiers in both Cabin Intercommunication Data System (CIDS) directors. The system control and functional status indication is prepared through discrete lines. The amplifier of the service interphone system is part of the CIDS director and is therefore energized when the Cabin Intercommunication Data System (CIDS) is switched ON. The service interphone system is switched on automatically when the landing gear squat switches are compressed for at least 10 s. The interphone OVRD switch switches the system on manually and the integral light of the service interphone system OVRD is then illuminated. Operation from the Cockpit The acoustic equipment in the cockpit transmits the audio signals. The audio signals are fed to the interphone amplifier in the CIDS director through the Audio Management Unit (AMU). The audio signals are transmitted to the attendant stations through the amplifier of the service interphone system. Alternatively the audio signals are transmitted to the service interphone jacks through the audio lines. Operation from the Cabin Attendant Stations This is done when you push the service interphone key (A320) or the interphone key twice (A319/321) on the attendant handset. Communication is done through the mid buslines to the director and the audio lines to the service interphone jacks and the audio lines from the director through the AMU to the cockpit acoustical equipment. The ”service system in use” sign comes into view on the Attendant Indication Panel (AIP) 320RH at each attendant station. The AIP indicates as soon as the service interphone system is active and at least one boomset is connected to the service interphone jack.

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A320 OXY MASK


PA ALL

CPT

AMU

3R

RESET

INT

CAB

1

2 MID

FWD 3

ALL

4 AFT

EXIT

CAPT

RESET

KEYBOARD

DEU B

DIRECTOR 1/2

AUDIO CONTROL PANEL

A319/321

INTPH

2x SERV INT

1L/R 3L

PA

MIDDLE LINES GROUND SERVICE JACK (8)

ATTENDANT INDICATION PANEL

OVERHEADD PANEL SVCE INT OVRD

LGCIU ON

Figure 90

Service Interphone Schematic

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PASSENGER LIGHTED SIGNS Description The state of input signals to the director control these signs: the NO SMOKING and FASTEN SEAT BELT lighted signs in the cabin, the RETURN TO SEAT signs in the lavatories, the EXIT signs. Manual commands are received from the cockpit NS and FSB switches. These are grounded inputs NO SMOKING COMMAND or NO SMOKING AUTO, and grounded input FASTEN SEAT BELT COMMAND respectively. The director interprets open circuited NS inputs and open circuited FSB input as NS OFF and FSB OFF commands respectively. Director software uses data from the CAM to activate signs which opened on: the cockpit commands, the discrete inputs LANDING GEAR DOWN LOCKED, an EXCESSIVE ALTITUDE, the FSB AUTO input, the inputs SLATS and FLAPS. The director also provides FSB AUTO input for the optional FSB AUTO cockpit commands. The interface to the EXIT signs is via a connection from the director to the emergency lighting system. DEUs type A drive and interface all other signs. The director addresses each sign independently. Switching on of any signs is signalled to the SDAC via the respective director output NO SMOKING or FASTEN SEAT BELT. Switching on of any NS sign is also signalled to the EMLS.

Operation of FASTEN SEAT BELT/RETURN TO SEAT Signs The FSB and RTS lighted signs in the cabin and lavatories respectively are all switched on under any of these conditions: Cockpit FASTEN SEAT BELT switch in overhead panel is switched on. The FSB signs are also activated in the event of excessive aircraft decompression, when the FSB switch is in AUTO or OFF position. When the FSB switch is in the AUTO position - causes the signs to be operated when the landing gear is down and locked. - causes the signs to be operated when the slats > 21 or Flaps > 19 . The PAX lighted signs are deactivated in the event of an aircraft emergency.

Operation of NO SMOKING/EXIT Signs All NS and EXIT signs are switched on under any of these conditions: Cockpit NO SMOKING switch in overhead panel is switched to ON. Excessive aircraft decompression, irrespective of the NS switch position. Landing gear down and locked when the cockpit NS switch is in the AUTO position.

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PSU LIGHTED SIGNS NS/FSB

LAVATORY LIGHTED SIGNS RTS

CABIN LOUDSPEAKERS

DEU A TOP LINES

DIRECTOR 1/2

MIDDLE LINES

LGCIU FWD ATTND PANEL

SFCC

CABIN PRESSURE SWITCH

ECAM

SDAC

DEU B

ATTENDANT INDICATION PANEL C H I M L E O G I C

HIGH HIGH/ LOW LOW

CFDIU

COCKPIT

EPSU TEST

EXIT ON

PROGRAMING AND TEST PANEL CAM

AVIONICS BAY Figure 91

EPSUs

EXIT EXIT SIGNS

EXIT CABIN

PAX lighted Signs Schematic

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PASSENGER / LAVATORY CALLS

CALL INDICATION IN THE CABIN

Description The equipment interface to the directors via DEUs type A as follows: one PAX call pushbutton and the seat row numbering light at each passenger seat row, one PAX call pushbutton with an integral lamp in each lavatory. Each DEU type A can interface up to three PAX call pushbuttons and lights. Each pushbutton and light is separately addressable. The CAM data assigns each pushbutton and light to a LH or RH cabin zone. First activation of a PAX or lavatory pushbutton activates a chime. Visual indications come on. A second activation of a PAX or lavatory pushbutton reset the visual indications.

Call to Attendants from PAX Seat When a passenger seat PAX call pushbutton is pressed: the associated call light seat row number with a L/H or R/H reference comes on or flashes (aircraft on ground and at least one passenger door open) a high 1 chime is broadcast over loudspeakers, the steady illumination of a blue light in the respective ACP (FWD or AFT, and RH or LH). the seat row number with a LH/RH reference is shown on the AIPs. The AIPs are related to the PAX call zones. Call to Attendants from LAV When a lavatory seat PAX call pushbutton is pressed: the call light initgrated in the pushbutton comes on. a high 1 chime is broadcast over loudspeakers, the steady illumination of a amber light in the respective ACP (FWD or AFT, and RH or LH). the pink light on the related AIP comes on. the number and the location of the related lavatory is shown on the AIP. The BITE status in the DEUs type A signals defective PAX call lamps to the director. Faults may be examined via the PTP.

TONE

SPKR ATTND

AIP

ACP ZONE ALL

PAX Call

1x High

X

--

Z

PAX Call LAV

1X High

X

X

Pink

Z

CAPT ATTND

1X High/ Low

X

X

Pink

Z

Pink steady

CAPT ALL

1X High/ Low

X

X

Pink

A

Pink steady

CAPTEMER CALL

3X High/ Low

X

X

Pink

A

Pink flashing

ATTND ATTND

1X High/ Low

X

Green

Z

Green steady

LAV Smoke

3X High

X

X

FSB NS

1X Low

X

X

CALL

SPKR PAX

INDIV Resp. Pax row Light LAV Call Light Amber steady

LAV Call Light Amber flashing FSB NS Signs

A

--

ACP LIGHT Blue steady L/H R/H Amber steady L/H R/H

Amber flashing L/H R/H

--

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SEAT ROW NUMBERING LIGHT NON SMOKER SIGN (A319/321) PASSENGER SERVICE UNIT A

B

C

14

CABIN LOUDSPEAKERS LAVATORY LIGHTED CALL PUSHBUTTON

X

14 TOP LINES

SDAC

DIRECTOR 1/2

DEU A

MIDDLE LINES FWD ATTND PANEL

CL HIGH HO I G HIGH/LOW MI LOW EC

LAVATORY CALL LIGHT DEU B AIP

PROGRAMMING AND TEST PANEL

CAM

Figure 92

AREA CALL PANEL

PAX and LAV Call Schematic

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PROGRAMMING AND TEST System Testing and Programming When the CIDS is energized, the directors perform a comprehensive hardware and software self-test. The CIDS top and middle line data busses, the PTP with the CAM, the FAP and the DEUs with the connected loads are tested. Programming and Test Modes The first appearing menu is the main menu. It shows the three main modes of the PTP: the SYSTEM STATUS mode, the SYSTEM TEST mode, the PROGRAMMING mode. Access Regulation The SYSTEM STATUS, the SYSTEM TEST and the ZONING (part of the PROGRAMMING mode) can be entered without access code. For the CABIN PROGRAMMING mode you have to enter a 3 digit access code. It prevents the CABIN PROGRAMMING (part of the PROGRAMMING mode) against unauthorized access. Access Code Entering Three digits are necessary for the access to the CABIN PROGRAMMING mode. There appears a * symbol on the display after input of each digit. The complete access code will be accepted by selecting the displayed ENTER function via the labelled key. Entering an incorrect access code initiates the PTP display message: ”USER AUTHORIZATION FAILURE”. A new entry can be started after activation of the displayed RET function. The aircraft is delivered with access code 333.

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CIDS
TEST EMER LIGHT

1

2

3

4

5

6

BAT

BAT OK

7

8

9

SYS

SYS OK

CLR

0

.

Figure 93

PTP Menu

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SYSTEM STATUS General The SYSTEM STATUS mode monitors the current status of the CIDS. This includes the directors, the DEUs, the data buses, the CAM, the FWD ATTND PANEL, the PTP and the interfaces to other systems. For detailed failure description the mode MAINTENANCE, which is part of the SYSTEM STATUS mode, can be selected. For support of the maintenance/cabin crew, the status of the following systems is also monitored: Lavatory Smoke Detection Slides Bottle Pressure Doors Bottle Pressure Drainmasts After selection of this mode, in case of no failure, the following messages are displayed on the PTP: CIDS OK LAV SMOKE SYS OK SLIDES PRESS OK DOORS PRESS OK DRAINMASTS OK A failure in one of these systems causes an annunciation. The SYSTEM STATUS mode is displayed automatically on the PTP, except when the PTP is in the MAINTENACE, the SYSTEM TEST or the PROGRAMMING mode. In case of failure the respective following messages are:
Detailed Messages Messages in case of
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2

1 CIDS

1

SYSTEM STATUS

MAINTENANCE

CIDS OK

CONT>

CONT>

SLIDES PRESS LOW

LAV SMOKE SYS OK

FWD L

DOOR PRESS OK

DRAINMASTS OK

CONT>

< SLIDES - STS - DOORS>

SLIDES STATUS

CONT>

FWD L=DISARM R=ARM MID L=DISARM R=DISARM

2

A319/321 only Figure 94

PTP System Status

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MAINTENANCE General A MAINTENANCE mode for detailed CIDS failure description is part of the SYSTEM STATUS mode. Failures are written into the directors BITE ground/ flight memory and can be read via this mode and via the CFDS in the cockpit. The last occured failure is the first for reading. For failures of continuously monitored systems the BITE ground memory will be updated when the failure has been cancelled. No entries into the BITE memory are made in case of emergency conditions (normal power is not available or the CIDS is working with the mandatory layout). No subsequent related failures will be entered into the memory after the original failure has been entered: Examples: If one DEU-A fails, no further entries into memory are made fo the associated components. If one DEU-B fails, no further entries into memory are made for the associated slides bottle pressure, but the system status on the PTP displays SLIDES PRESS LOW.

LRU IDENTIFICATION Messages for LRU Identifications: DIRECTOR 1 DIRECTOR 2 CAM Note: M-COUNT = modification count of layout M PROG AND TEST PNL

FAULT DATA This FAULT DATA mode includes the flight leg, the date, the time (UTC), the number of occurences (max 4 counts, for intermittent failures) and coded trouble shooting data for internal director and DEU failures. Present failures on ground are marked with GND, failures of the last leg with LEG -00, and failures of the previous legs with -01, -02 and up.

CLASS 3 FAULTS

MAINTENANCE via CFDS and PTP The maintenance mode via the CFDS-MCDU is available with the SYSTEM REPORT/TEST mode. All failures, which are written in the CIDS director BITE ground/flight memory can be read via this mode. The maintenance via CFDSMCDU follows the same procedure as the maintenance via the PTP. A test procedure is selectable via the MCDU. A CIDS director 2 test is also available, the Emergency Light Battery/System tests are not available.

Some CLASS 3 FAULTS are only detected and written into the BITE memory at director power on or after a manual test activation via PTP. On ground, after cancelling a class 3 fault, it disappears from the memory after the next director power on or after a new test activation. Some CLASS 3 FAULTS are detected and written into the BITE memory due to continuous monitoring. Such a class 3 faul disappears from the memory after the failure itself has disappeared.

LAST LEG REPORT

GRND SCAN

In flight, this report is called CURRENT LEG REPORT. It is the only displayed and accessable item within the MAINTENANCE mode in flight. The LAST/ CURRENT LEG REPORT contains class 2 + 1 failures of the last/current flight leg. The report includes the date, the time and the ATA chapter for each failure. There are no entries for flight legs without failures but the flight leg counts are incremented.

The GND SCAN indicates all class 1 + 2 failures which are present on the ground. For the continuously monitored systems, the ground memory is updated when the failure is cancelled, for other systems, the ground memory is updated after a director power on or after a test activation via PTP.

PREV LEGS REPORT The PREV LEGS REPORT contains 1 + 2 failures of the last 64 flight legs. The PREV LEGS REPORT has all data, which are stored in the LAST LEG REPORT.

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MAINTENANCE

1

2

CONT>

2

1
LAST LEG REP APR12

PREV LEG -30 MAR13

DEU-A 200RH15

DIRECTOR 2

UTC 0700 ATA23-73-46

UTC 0712 ATA23-73-34

CONT>

CONT>

CONT>

LAST LEG REP APR12

PREV LEG -38 MAR05

CAM

FWD ATTND PNL

UTC 0700 ATA23-73-46

UTC 0540 ATA23-73-12

Figure 95

PTP Maintenance Menu

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3 MAINTENANCE
DIRECTOR

1: Z010H0002110

3

LRU IDENTIFICATION

CONT>

CONT>

LRU IDENTIFICATION

DIRECTOR

2: Z010H0002110

CONT>

CONT>

LRU IDENTIFICATION

LRU IDENTIFICATION

CAM M-COUNT=005

PROG AND TEST PNL

Z050H00000343

Z020H0000110

Figure 96

CONT>


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Page 495

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MAINTENANCE
CONT>

5

4
FAULT DATA

4

LEG

UTC

N

CODE

5

-03

1340

4

240A07

CONT>

FE27

CLASS 3 FAULTS ATA 23-73-20 SIGN LAMP 02L, 05R
6 6

GRN SCAN NO FAILURE

Figure 97


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Page 496

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6

5

Figure 98

CIDS MCDU BITE Menu

______________________________________________________________________________________________________________________________________________________________________________________________



Page 497

MNG TECHNIC


1

2

4

3

Figure 99

CIDS MCDU BITE Menu

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SYSTEM TEST MODE General The general SYSTEM TEST mode is one of the main modes. It initiates the test except: when the aircraft is in flight, normal power is not available (PTP not powered), the mandatory layout is in use. These devices can be tested: Director 1 (the currently active director is marked, example DIR1 (ACT), the passive one can only be tested via the MCDU). Director 2 (see Director 1) CIDS BUS DEUs, Type A DEUs, Type B Programming and Test Panel (membrane switches are not checked) CAM FWD ATTND Panel (the pushbutton and membrane switches are not checked) AFT ATTND Panel (membrane switches are not checked) ATTND Indication Panels Loudspeakers (only operational test) Sign Lamps PAX Call Lamps Area Call Panels (only operational test) Reading/Work Lights Emergency Lighting Battery Drainmasts Additionally, there is a RESET function. It initiates a general CIDS power on reset including the power on test of the complete system. If failures still exist, these can be read from the automatically displayed SYSTEM STATUS/MAINTENANCE mode on the PTP. The complete power on test is only performed if one or more cabin doors are open (same as for CIDS power on).

Initiation of System Tests If you push the device related membrane switch, the test of this device is activated. The flashing message -WAIT FOR RESPONSE appears on the display. When the test is finished, the ATA chapter and the message TEST OK comes on. An old failure message in the director’s BITE ground memory is cancelled. In case of a failure, the respective result is written into the directors BITE memory and the failure message appears on the PTP display. The TEST mode is not available in flight.

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1 CIDS
1

SYSTEM TEST
RESET>

CONT>

CONT>

DIRECTOR 1

ATA 23-73-34

TEST OK

CONT>

CONT>

Figure 100

CONT>

PTP Test Menu

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Page 500

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Figure 101

CIDS MCDU BITE Test Menu

______________________________________________________________________________________________________________________________________________________________________________________________



Page 501

MNG TECHNIC




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PROGRAMMING MODE General The programming mode is one of the main modes. It is divided into: the zoning (needs no access code), the cabin programming (needs a 3 digit access code, A/C delivered with access code 333).

ZONING Cabin Zones For CABIN ZONES programming, enter the seatrow number of the end of the zone. The adjacent zone adapts automatically. NS Zones NS ZONES starts is accordance with the programmed CABIN ZONES. Each cabin zone starts with a NS zone. For programming, enter the seatrow number at the end of the NS zone. Entering a 0 (zero) deletes the NS zone in the related cabin zone. If the layout of the CABIN ZONES is changed, the NS zones layout follows automatically. The number of seatrows of each NS zone remains constant, except the cabin zone is smaller than the NS zone. If the cabin zone is extended again, the previous NS zone length is realized. The class divider separates the cabin zones. The CLASS DIVIDER programming is only used for BITE related current sensing reasons: installing a passive divider (without sign lamps) instead of an active divider (with sign lamps) or the reverse. removing an active divider (the layout of the cabin zones are not changed). Note: On A321 the Cabin lighting is independent from the Cabin programming. The FWD zone controls the area between door 1 and door 2, the AFT zone the area between door 2 and door 4. On A321 the Call assigment (PAX call) is independent from the Cabin programming.

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2

1 CIDS

NS END FWD 07 >

CABIN ZONES

END FWD

09 >

NS END MID 12 >

END MID

16 >

NS END AFT

21 >

insert 06 :A321 only

PROGRAMMING

CABIN ZONES

END FWD

09 >

END MID

16 >

push

06

3 CLASS DIVIDER

ZONING

1

2

DIV>

3

PARAMETER SAVED

-PUSH FOR CHANGE

MODIFICATION 019

Figure 102

FWD ACTIVE >

PTP Zoning Programming

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CABIN PROGRAMMING CAM Layout Selection The Cabin Programming is access code protected. After entering the 3 digit access code and selecting the CAM LAYOUT selection mode, the respective menu comes on. The programmed layouts are marked with a < or > sign. The number of active layout is flashing. A new layout is selected by pushing the related membrane switch. After selection, a new layout is marked and down loaded into the director. The system is updated automatically and CIDS works with this layout until a new is selected and loaded.

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1 CIDS

ENTER ACCESS CODE

ENTER>

insert 333 PROGRAMMING

1

CAM LAYOUT SELECTION <1

LAYOUT

2

3

LAYOUT

M>

LAYOUT 1 LOADED

Figure 103

CAM LAYOUT SELECTION

PTP Layout Selection

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FAULT ISOLATION AND BITE Failure Detection and Transmission There are 3 types of test available in the CIDS for failure detection: the power on test (activated after every power connection to the CIDS), the continuous test (automatic periodical system test) the manually activated test (via PTP, CFDS ...) The status is transmitted to the CFDS (via ARINC 429) and ECAM (via discret outputs to the SDAC). The failure indication is possible on: the FAP (CIDS caution light), the PTP, the CFDS/MCDU display, the ECAM displays. In flight, it is possible to reset the illuminated caution light on the FAP. After landing, if the failure still exists, the light comes on again and the SYSTEM STATUS mode is displayed. The failures are divided into 4 failure classes, 1, 1CAB (cabin), 2 and 3. The relation of failure classes and the transmission to the indicators are shown on the Failure Transmission list.

CIDS Power-Up Test Time that the computer must be de-energized: A/C on ground with engines stopped and 1 or more cabin doors open : 10 sec Progress of Power-Up Test: Duration: approx. 120 sec Cockpit repercussions directly linked to power-up test accomplishment (some other repercussions may occur depending on the A/C config. but these can be ignored): ECAM MAINTENANCE STATUS: - CIDS 1 on, after approx. 30 sec off - CIDS 2 on, after approx. 80 sec off Audio Control Panel: - ATT light flashes NOTE : Depending on the customized cabin config and the position of the NS/FSB switch the following repercussions can be observed in the cabin: cabin lights are off for approx. 1 sec call/ seatrow numbering lights and signs flash two chimes are heard after approx. 120 sec after power-up test initialisation Results of Power-Up Test: Test pass : - none NOTE : : On the PTP the CIDS OK message comes on. Test failed : if CIDS 1 or CIDS 2 failed - ECAM MAINTENANCE STATUS : CIDS 1 or CIDS 2 comes on if CIDS 1 and CIDS 2 failed - MASTER CAUTION light comes on - ECAM WARNING : COM CIDS 1 + 2 FAULT comes on - ECAM INOP SYS : CIDS comes on

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Figure 104

CIDS Failure Classes

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CIDS WARNINGS ON ECAM AND FAP General All BITE results are stored in the director BITE dedicated memory with failure indication shown on the CIDS caution light, the ECAM Engine/Warning Display, the Status Page on the ECAM System display, the MCDU and the PTP. A failure in the initialization progress in director 1 causes : Display of the STS reminder on the Engine/Warning Display. Display of the CIDS 1 message on the STATUS page (on request). Note: The STATUS page does not appear automaticlly after warning. Class 1 Fail ECAM Warning If a class 1 fail occurs, the single chime sounds, the ECAM MASTER CAUT and the CIDS CAUT light comes on. The failure messege (CIDS 1+2 FAULT) is shown on the ECAM Engine/Warning Display, and on request on the STATUS page. The CIDS caution on ECAM occurs if: both director fail, more than 50% of all DEU A fail more than 20% of adjacent DEU A zonewise fail, all DEU B with connected handset fail. Class 2 Fail ECAM Warning When a CIDS caution occurs on the ECAM, the detailed failure message is memorized in the PTP and available on request (on ground only). The caution messages on the ECAM are the STS reminder on the Engine/ Warning page, and the maintenance message (Class 2) on the STATUS page which is displayed on request. The DIRECTOR continues to signal CIDS caution to the SDAC until the fault is corrected. The CIDS caution on ECAM occurs if: one director fail, no crosslink from other director

CIDS Caution light on FAP Some CIDS caution or failure signals (Class1, 1CAB, 2) activate also the CIDS CAUT light on the FAP. Note: The CIDS CAUT light is resetable in flight, but comes on again on the ground (landing gear down and locked). The light cannot be reset on ground. When a CIDS CAUTION occurs, the respective failure message is displayed on the PTP. The CIDS CAUT light comes on if: both directors fail, more than 50% of all DEU A fail, 20% of adjacent DEU A zonewise fail 50% of all NS-FSB signs fail and PA fail, all DEU B with connected handset fail. no data from SDCU channel 1+2, the lavatory smoke detector fails, the heater of drain mast fails, the control unit of the drain mast fails, the slide/door pressure is low, DEU-A or DEU-B fails, CIDS TOP or MID BUS fails, FAP or PTP fails or no data, CAM fails.

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CIDS Class1 Warnings

FLIGHT PHASE INHIBITION

1

2

MASTER WARN

5 6 7

3 4

CIDS Class1CAB or Class2 Warnings

8 9 10 MASTER CAUT

STATUS

COM CIDS 1+2 FAULT

single chime STATUS INOP SYSTEM CIDS 1

CIDS

FWD ATTND PANEL

LIGHT EMER

SMOKE LAV

RESET

FWD ATTND PANEL

CIDS PNL LIGHT TEST

Figure 105

CAUT

LIGHT EMER

SMOKE LAV

RESET

CIDS PNL LIGHT TEST

CAUT

CIDS ECAM and FAP Warnings

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Page 510

MNG TECHNIC


Figure 106

CIDS MCDU BITE Test Menu

______________________________________________________________________________________________________________________________________________________________________________________________



Page 511

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Figure 107

CIDS Failure List

______________________________________________________________________________________________________________________________________________________________________________________________



Page 512

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Figure 109

CIDS Failure List

______________________________________________________________________________________________________________________________________________________________________________________________



Page 513

MNG TECHNIC


Figure 110

CIDS Failure List

______________________________________________________________________________________________________________________________________________________________________________________________



Page 514

MNG TECHNIC


Figure 111

CIDS Failure List

______________________________________________________________________________________________________________________________________________________________________________________________



Page 515

MNG TECHNIC


Figure 112

CIDS Failure List

______________________________________________________________________________________________________________________________________________________________________________________________



Page 516

MNG TECHNIC



Figure 113

CIDS Failure List

______________________________________________________________________________________________________________________________________________________________________________________________



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LOCATION

Figure 114

CIDS Location Cockpit

______________________________________________________________________________________________________________________________________________________________________________________________



Page 518

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Figure 115

CIDS Locations DEUs and Directors

______________________________________________________________________________________________________________________________________________________________________________________________



Page 519

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Figure 116

CIDS Location DEUs A321

______________________________________________________________________________________________________________________________________________________________________________________________



Page 520

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Figure 117

CIDS Location DEUs A319

______________________________________________________________________________________________________________________________________________________________________________________________



Page 521

MNG TECHNIC


Figure 118

CIDS Location Cabin

______________________________________________________________________________________________________________________________________________________________________________________________



Page 522

MNG TECHNIC


Figure 119

CIDS Location Cabin

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23-32

PRAM

PRAM PRESENTATION The function of the Prerecorded Announcement and Boarding Music (PRAM) Reproducer 10RX is to play prerecorded messages. It also plays boarding music programs on a cassette tape to the passengers through the aircraft passenger address system. The PRAM is controlled by the audio module, which is a part of the Fwd Attnd panel 120RH. It is installed in the cabin at the forward attendant station. The PRAM is controlled through the Cabin Intercommunication Data System (CIDS) director to receive and transmit control data.

SYSTEM OPERATION AND CONTROL The Prerecorded Announcement and Boarding Music (PRAM) system is controlled during normal operation from the audio module in the Fwd Attnd panel 120RH. The system initialization is made automatically when the reproducer receives 115 V AC. During the initialization the LED display on the audio module is not shown. The least significant digit LED of the announcer display will show ’0’ when the initialization is complete. The operation procedures to program and release the announcements are as follows: Announcement Control Keys ENTER When the ENTER pushbutton is pushed, the cursor moves into the next MEMO position (MEMO 1, if no cursor is present). The required message is keyed-in on the keyboard and appears on the MEMO 1 display. When the ENTER pushbutton is pushed, the keyed-in data is accepted (cursor moves to the next MEMO position). The READY light comes on when the PRAM has found the corresponding announcement. The required messages for the MEMO 2 and MEMO 3 displays are keyed-in the same as for the MEMO 1 display. CLEAR When the cursor has moved into the related position (MEMO 1, MEMO 2 or MEMO 3) and the CLEAR pushbutton switch is pushed, the display clears. START NEXT

When the START NEXT pushbutton switch is pushed and the READY light is on, the message shown on the MEMO 1 display moves up to the ’ON ANNOUNCE’ display. The MEMO 2 display message then moves up to the MEMO 1 display. The message shown on the MEMO 3 display moves up to the MEMO 2 display. START ALL All messages keyed on MEMO 1, MEMO 2 and MEMO 3 will be announced continuously until the last announcement has finished (number 0). In this continuous mode, it can add another message for announcement after the messages that you have keyed-in. STOP When the STOP pushbutton switch is pushed, the message announcement stops immediately. Emergency Announcement Test To test the emergency announcement, bring the cursor to MEMO 1 and press ”701 ENTER 701” (E-P is displayed in the ON ANNOUNCE display). Boarding Music Control Keys ON/OFF When the ON/OFF pushbutton switch is pushed, the light in the pushbutton switch comes ON. The channel 1 is automatically displayed on the BGM channel display. When the ON/OFF pushbutton is pushed again, the light goes OFF. SEL When the SEL pushbutton switch is pushed, the system selects one of the available channels. These are displayed in a numerical ascending code (four channels in the mono mode, two in the stereo mode). Note: LH uses two mono music programs (special mode of operation). Channel 1+2 is boarding music one and channel 3+4 is boarding music two. VOLUME The LEDs on the volume display show the volume level (2 dB steps). (-) When the (-) pushbutton switch is pushed the volume level decreases. (+) When the (+) pushbutton switch is pushed the volume level increases.

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Figure 120

PRAM Control Panel

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PRAM DESCRIPTION (SPECIAL OPERATION)

FRONTPANEL CONTROL (SPECIAL OPERATION)

The Prerecorded Announcement and Boarding Music (PRAM) Reproducer is installed in the aft avionics compartment in a 4 MCU (ARINC 600) box. It has a total of four cassette decks, but only the tape in announcement tape deck A position is used (the other tape decks are disabled) . This cassette is used for the prerecorded announcement reproduction and for the boarding music reproduction. Up to 32 prerecorded announcements are stored on the first two tracks on that tape and two boarding music channels are available on the other two tracks. The PRAM has the capability to produce an emergency announcement in the event of a rapid cabin decompression and up to four other stored announcements. The emergency announcement and the four other once are stored in a Solid State Stored Voice (SSSV). A ground signal from a rapid decompression triggers the emergency announcement. The other announcements can be triggered by separate discret-inputs. All functions are remotely controlled from the audio module in the Fwd Attnd panel (120RH) (except the output level of normal and emergency announcements). They are adjustable at the front of the reproducer. The announcement part of the PRAM uses track 1+2 of tape A with up to 16 announcements on each track. Messages are searched and located by detecting and counting the inter-message blank portion (8sec.) The current tape position is memorized, a subsequent message search on the same track is possible. If track change is necessary, a completely tape rewind is made. Messages can only be searched and played one after the other (no continuous play) The boarding music part of the PRAM uses track 3+4 of tape A. Each track is a complete program and will be repeatedly played (rewinding tape after end of tape). The audio module in the Fwd Attnd panel 120RH controls the prerecorded announcements and the Boarding Music (BGM). The reproducer and Fwd Attnd panel have two ARINC 429 data bus lines (transmit and receive bus) controlled through the CIDS.

The PRAM can be controlled and tested from the frontpanel. Music Channel Selector no effect. TEST Selector For normal use of the PRAM, the switch must by in the NORMAL position. all other positions are for testing the announcement tape and SSSV emergency announcements. In the DECK A1 and A2 position it is possible to check the announcement audio of track 1+2 from tape A via the monitor output connector. In the DECK A3 and A4 position it is possible to check the boarding music audio of track 3+4 from tape A via the monitor output connector. In the DECOMP position its possible to check the emergency announcements stored as Solid State Stored Voice (SSSV) in a memory, when a decompression occurs. In the SSSV 1(2,3,4) position its possible to check the four different announcements stored as Solid State Stored Voice (SSSV) in a memory (if available). The TAPE INITIAL position has no effect on Special Mode of Operation (a initialization routine is not necessary).

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Music Channel Selector Test Selector FAP Audio panel

Annoncement TAPE B Position (spare tape on LH) DEU-A Annoncement TAPE A Position Music Tape A + B (not LH) Director 1

DEU-B

Audio Cabin Decomp.

PRAM

Figure 121

PRAM Schematic and Frontpanel

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FRONTPANEL CONTROL (NORMAL OPERATION)

PRAM DESCRIPTION (NORMAL OPERATION) The Prerecorded Announcement and Boarding Music (PRAM) Reproducer is installed in the aft avionics compartment in a 4 MCU (ARINC 600) box. It has a total of four cassette decks. Two are for the prerecorded announcement reproduction and the other two are for the boarding music reproduction. Up to 256 prerecorded announcements are stored on two tapes in the prerecorded announcement reproducer (each minimum 30 s). The PRAM has the capability to produce an emergency announcement in the event of a rapid cabin decompression and up to four other stored announcements. The emergency announcement and the four other once are stored in a Solid State Stored Voice (SSSV). A ground signal from a rapid decompression triggers the emergency announcement. The other announcements can be triggered by separate discrete-inputs. All functions are remotely controlled from the audio module in the Fwd Attnd panel (120RH) (except the output level of normal and emergency announcements). They are adjustable at the front of the reproducer. The announcement part of the PRAM uses two identical tape decks with up to 256 announcements on each tape. The two cassette tapes are used alternately. When one is playing a message, the other is in search mode for the next message to be announced ( if necessary) to give continuous play. The boarding music part of the PRAM uses two identical tape decks with four channels. Two cassette tapes are used alternately. When one is playing the other rewinds to give continuous play. The audio module in the Fwd Attnd panel 120RH controls the prerecorded announcements and the Boarding Music (BGM). The reproducer and Fwd Attnd panel have two ARINC 429 data bus lines (transmit and receive bus) controlled through the CIDS.

The PRAM can be controlled and tested from the frontpanel. Music Channel Selector In the REMOTE Position, the PRAM is remote controlled from the audio part of the FAP. In the 1 (2,3,4) position, only the music track 1 (2,3,4) is played continuously. In the AUTO position, the different music tracks are played one after the other automatically and continuously. TEST Selector For normal use of the PRAM, the switch must by in the NORMAL position. all other positions are for testing the different announcement tapes and SSSV emergency announcements. In the DECK A1 (2,3,4) or DECK B1 (2,3,4) position it is possible to check the announcement audio of each track from tape A or B via the monitor output connector. In the DECOMP position its possible to check the emergency announcements stored as Solid State Stored Voice (SSSV) in a memory, when a decompression occurs. In the SSSV 1(2,3,4) position its possible to check the four different announcements stored as Solid State Stored Voice (SSSV) in a memory (if available). In the TAPE INITIAL position, a special mode is activated to read initialization data from the announcement tapes A and B in order to store this data in memory. NOTE: The TAPE INITIAL position must be used when the PRAM access door was open (for check or tape change) in order to initialize the PRAM tapes for at least 6 seconds.

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Figure 122

PRAM System Schematic

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LOCATION

Figure 123

PRAM Location 80 VU

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Figure 124

PRAM Location Cabin

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23-71

COCKPIT VOICE RECORDER

CVR PRESENTATION The cockpit voice recorder (CVR) is designed to record crew conversations and communications on a magnetic tape, in flight and on ground, when at least one engine is running or up to five minutes after the second engine is shut down irrespective of which engine is shut down first. The system can also operate in manual mode on the ground. The recorder is a four-track system and all tracks are recorded simultaneously. The recording tape is of the magnetic loop type which allows 30 minutes of recording time. When the tape is fully recorded, the system progressively erases recordings made in the previous 30 minutes and simultaneously records new information ; thus only information recorded in the last 30 minutes of operation is retained. The recorded information can be intentionally erased when the aircraft is on the ground with the parking brake control handle pulled up, locked and electrically powered. Bulk erasure is also possible during manual operation of the system. Recording of conversations and communications must comply with standards specified by the FAA.

ACTIVATION AND CONTROL Power Supply The CVR is automatically supplied with 115VAC when the aircraft is in one of the configurations given below : in flight with engines running or stopped. on the ground with at least one engine running on the ground during the first five minutes following energization of the air craft electrical network. on the ground up to five minutes after second engine shutdown. Manual selection of power supply to the CVR allows the functions given below with the aircraft on the ground and both engines shutdown : To test the CVR for correct operation. To erase tape information if required. To record the beginning of the check list before the first engine starts running. For manual selection of power supply to the CVR press the GND CTL pushbutton on the control panel.

CVR Erase Circuit The Recorder ERASE pushbutton must be pressed in for a minimum of two seconds to prevent inadvertent erasure. The erause head erases the previously recorded information on all 4 channels simultaneously, before a new recording is made. The ERASE pushbutton enables complete erausre of the tape by activation of a magnetic field. ERASE is only posible when aircraft is on ground (R and L main gear shock absorber compressed) and parking prake applied. Note: if the engines are shut down, the CVR must first be energized. CVR Test Circuit The CVR TEST is initiated by pressing the CVR TEST pushbutton on the RCDR panel. A 600 Hz test tone is applied sequentially during 0.8s to each of the four tracks. This signal should then be heared through the loudspeakers (provide the aircraft is on ground, R and L main gear shock absorber compressed and parking prake applied) and through the headset conected to the maintenance panel 50 VU in the cockpit. Note: prior to this test, the CVR must be energized.

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POWER SUPPLY INTERLOCK

Figure 125

CVR Schematic

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DESCRIPTION The cockpit voice recorder (CVR) is designed to record crew conversations and communications on a magnetic tape, in flight and on ground, when at least one engine is running or up to five minutes after the second engine is shut down irrespective of which engine is shut down first. The system can also operate in manual mode on the ground. The recorder is a four-track system and all tracks are recorded simultaneously. The recording tape is of the magnetic loop type which allows 30 minutes of recording time. When the tape is fully recorded, the system progressively erases recordings made in the previous 30 minutes and simultaneously records new information ; thus only information recorded in the last 30 minutes of operation is retained. The recorded information can be intentionally erased when the aircraft is on the ground with the parking brake control handle pulled up, locked and electrically powered. Bulk erasure is also possible during manual operation of the system. Recording of conversations and communications must comply with standards specified by the FAA.

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Figure 126

CVR Detailed Schematic

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Figure 108

CIDS Failure List

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POWER SUPPLY INTERLOCK Aircraft in flight The LGCIU1 5GA1 provides a ground signal to the relay 6RK. The relay 6RK is supplied directly with 28VDC from the bus bar 801PP via the circuit breaker 4RK. The CVR is supplied with 115VAC via the normally-open contacts of the relay 6RK from the bus bar 801XP through the circuit breaker 2RK. The normally-closed contacts of the relay 8RK provide a parallel path with the normally-open contacts of the relay 6RK for the supply of 115VAC to the CVR. Aircraft on ground With one or both engines running, no ground signal is fed to the relay 10RK. Since the relay 10RK controls the energized or de-energized state of relay 8RK, the relay 8RK remains de-energized. The 115VAC is connected from the bus bar 801XP via the circuit breaker 2RK through the normally-closed contacts of the relay 8RK to the CVR. During the first five minutes of energization of the aircraft electrical network, with both engines shutdown, a ground signal is fed to the time-delay relay 10RK. The relay 10RK is supplied with 28VDC from the bus bar 801PP via the circuit breaker 4RK. When energized this relay starts its timing function. During this timing function, the relay 8RK remains de-energized and the CVR is supplied with 115VAC. After 5 minutes, a ground signal is sent via the normally-open contacts of the time-delay relay 10RK to energize the relay 8RK which cuts off the supply of 115VAC to the CVR. Up to five minutes after second engine shutdown, a ground signal is sent to the time-delay relay 10RK. The relay 10RK is supplied with 28VDC from the busbar 801PP via the circuit breaker 4RK. When energized, this relay starts its timing function. During this timing function, the relay 8RK remains de-energized and the CVR is supplied with 115VAC. After 5 minutes, a ground signal is sent via the relay 10RK to energize the relay 8RK which cuts off the supply of 115VAC to the CVR. NOTE : In the cases described in Para. 2 (b) and (c) above, the relay 6RK is de-energized (A/C on the ground) and the recording function stops.

Manual power supply Manual selection of power supply to the CVR allows the functions given below with the aircraft on the ground and both engines shutdown : To test the CVR for correct operation. To record the beginning of the check list before the first engine starts running. To erase tape information if required. With both engines shutdown, a ground signal is sent to the relay 12TU. Pressing the RCDR/GND CTL pushbutton switch 11TU supplies the 28VDC to the relay 12TU via the normally-open contacts of the pushbutton switch 11TU and the normally-closed contacts of the relay 13 TU. The relay 12TU is energized. A ground signal is sent to the blue ON legend which comes on. A ground signal is also sent to the relay 6RK which energizes and supplies the CVR with 115VAC power. When the RCDR/GND CTL pushbutton switch is released, the 28VDC is applied via relay 12TU and pushbutton switch 11TU to the relay 13TU which is energized. If one engine is started, the ground signal to relay 12TU is removed and the relay 12TU is de-energized. Thus the blue ON legend goes off and the relay 6RK is de-energized.

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Figure 127

CVR Power Supply Schematic

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LOCATION

Figure 128

CVR Location Stabilizer Comp.

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Figure 129

CVR Location Cockpit

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ATA 25

EQUIPMENT/FURNISHING

25-XX

EMERGENCY LOCATER TRANSMITTER (ELT)

GENERAL The Automatic Fixed three-frequency Emergency Locator Transmitter (ELT) is installed in the upper aft section of the fuselage. It is designed to transmit a digital distress signal to satellites that are part of the COSPAS/SARSAT system. These satellites relay the received signal directly to reception stations on the ground. This signal is transmitted on 406 MHz and is used to locate and identify the ELT. The ELT also transmits a 121.5 MHz and 243 MHz homing signal for close-in aircraft location determination. The ELT can be manually activated from the Control Panel on the Pilots Overhead Panel or automatically activated by means of an internal acceleration sensor (G-switch). The ELT can be reset after automatic or manual activation by the ELT control panel switch. ELT activation is indicated by an ON LED at the control panel. The system components are: one ELT (with Mount/Programming Unit) one external antenna one controlpanel

COMPONENTS -The Mount/Programming Module stores the airplanes tailsign and countrycode (for Germany it is 218). The Mount/Programming Module is programmed by the manufacturer or by Lufthansa WF2. The Mount/Programming Module automatically downloads the 406 MHz programming information to the ELT when the ELT is installed. -The ELT is mounted on the mount/programming module, which is located in the aft cabin ceiling, left of centerline. The ELT receives datas from the mount/ programming module and stores it in its own memory. -The ELT Antenna, an external blade type, three-frequency antenna, is mounted on the upper aft section of the fuselage.

-The ELT Control Panel, located on the Overhead Panel, contains a two position guarded, toggle ON/AUTO switch and a TEST/RESET pushbutton.

OPERATION With the switch in the AUTO position, the ELT can be automatically activated by the G-Switch. Placing the switch in the ON position activates the ELT, which will begin transmitting distress signals after a self-test of approximately 150 seconds duration. Pushing the TEST/RESET pushbutton stops the ELT transmission and/or initiates a self-test of the ELT, Programming Unit, and verifies proper programming information transfer. The ELT then returns to the AUTO mode. A successful self-test is indicated by appearance of the ELT ON LED within 15 seconds, remaining uninterrupted for approximately 10 seconds. WARNING: EXCEPT IN THE FIRST 5 MINUTES OF THE HOUR, NEVER PLACE THE ELT SWITCH IN THE ON POSITION!! RESCUE ACTIONS WILL BE UNDERTAKEN!! ANYHOW, IF YOU DID SO, CALL IMMEDIATELY THE NEXT RESCUE CENTER FOR CANCELLATION OF THE ALARM! AREA PHONE NUMBER Frankfurt/Main, Stuttgart, Saarbrücken, Nürnberg 069 690 78801 Munich 089 978 0330 Berlin, Leipzig, Dresden, Erfurt. Laage 030 6951 2488 Düsseldorf, Köln, Münster. Paderborn 0211 4216254 Bremen, Hamburg. Hannover 0421 5372120 Installation of the ELT When the programming unit is connected to the ELT, data is automatically transferred to the ELT to code it with the aircraft information. Upon completion, verification is made of the validity of parameters transferred to the ELT. In case of a failure, the red TX indicator on the programming unit will flash for 10 seconds. If verification is successful, the red TX indicator will illuminate continuously for 10 seconds.

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Figure 130

ELT Schematic

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LOCATION

AUTO/OFF/ON Switch Transmit Light

Figure 131

Emergency Locator Transmitter (ELT)

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Location on A321

ETL Antenna

ELT Control Panel (Flight Compartment)

Figure 132

ELT Components

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Training Manual A 319/320/321 ATA 31 Indicating/Recording 31-50 31-60 31-21

Central Warning Systems Electronic Instrument System

Electrical Clock

ATA Spec. 104 Level 3

Lufthansa Issue: January 1998 Technical Training GmbH For Training Purposes Only Book No: ALL A319/320/321 31-50 LEVEL 3 E Lufthansa Base Lufthansa 1995 ______________________________________________________________________________________________________________________________________________________________________________________________



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Lufthansa Technical Train


ATA 31

INDICATING/RECORDING SYSTEMS

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COCKPIT PHILOSOPHY All the aircraft and system controls are arranged to be within easy reach of the two crew members. The concentration of system controls on the overhead panel is achieved by extensive use of illuminated pushbuttons directly installed on the system synoptic panel. In normal operation, no lights are illuminated in the cockpit. This is called the ”lights out philosophy”.

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Figure 1

Electric Panel

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Lufthansa Technical Train


PUSHBUTTON PRINCIPLE General Status and failure indications are integrated whenever possible into the relevant illuminated pushbuttons which must be operated for corrective action.

Pushbuttons with two stable Positions Most of the illuminated pushbuttons have two stable positions: pressed in and released out; each position corresponding to a control signal sent to a system. Pressed in (Recessed): Normaly used system activation (AUTO or ON) Temporarily used system activation (ON) System activated for maintenance operation (ON) or override (OVRD). Pushbuttons with one stable Position Some pushbuttons have only one stable position: released out. When pushed they send a control signal to the system. Released out (flush with the panel): System deactivation (OFF) Manual activation of a system (ON) Activation of an alternate system (ALTN).

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Typical Pushbutton with two stable Positions 1. Pressed in

ADR1 PRESSED IN NO LIGHT SYSTEM ACTIVATED CORRECT OPERATION

ADR1

PRESSED IN FAULT LIGHT ON SYSTEM ACTIVATED FAULTY CONDITION

2. Released out

ADR1 RELEASED OUT OFF LIGHT ON SYSTEM DEACTIVATED

Figure 2

Pushbutton Principle

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COLOUR PHILOSOPHY The illuminated pushbutton and annunciator lights are of different colours according to their function. In normal operation, only green lights and sometimes blue lights are illuminated.

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RED IS USED FOR A FAILURE NEEDING IMMEDIATE ACTION

AMBER IS USED FOR A FAILURE NEEDING AWARENESS BUT NO IMMEDIATE ACTION

WHITE IS USED TO INDICATE A PUSH BUTTON IN AN ABNORMAL POSITION OR MAINTENANCE OPERATION

GREEN IS USED TO INDICATE NORMAL OPERATION OF A BACK UP SYSTEM

BLUE IS USED TO INDICATE NORMAL OPERATION OF A TEMPORARILY USED SYSTEM

Figure 3

Colour Philosophy

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31-60

ELECTRONIC INSTRUMENT SYSTEM (EIS)

Modul: EIS GENERAL EIS PANELS The EIS (Electronic Instrument System) presents Data for: Electronic Flight Instrument System ( EFIS ) Electronic Centralized Aircraft Monitoring ( ECAM ). The 6 Display Units ( DUs ) are identical and interchangeable.

EFIS The Primary Flight Display presents all the flight parameters necessary for short term aircraft control. The Navigation Display presents navigation and radar information. The EFIS Displays are: PFD : Primary Flight Display ND : Navigation Display ECAM The Engine and Warning Display presents engine indications, fuel quantity, and Flaps/Slats position. The lower display presents either system pages synoptics or status messages. The ECAM displays are: Engine and Warning Display ( E/WD ) System or Status Display ( SD ).

EFIS Controls The EFIS displays are controlled by an EFIS Control Panel and PFD/ND Transfer Pushbutton. Two EFIS control panels are provided. A PFD/ND Transfer Pushbutton is also fitted on each side.

ECAM Controls The ECAM displays are controlled by an ECAM Control Panel. The ECAM Control Panel and various switching controls are located on the Center Pedestal.

Reconfigurations The displays can be transferred automatically if a system failure is detected. It is also possible to transfer them manually.

Attention Getters The warning messages are accompanied by either a MASTER WARNING, or a MASTER CAUTION and an aural warning. The ”MASTER WARN” light flashes red for any red warning. The ”MASTER CAUT” light comes on amber for level 2 amber warnings. Aural warnings are broadcast by two loudspeakers.

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Attention Getters EFIS Control Panel

EFIS Control Panel

PFD/ND Xfr

PFD/ND Xfr

o

o E/WD PFD

ND

ND

PFD

SD Loudspeaker

Loudspeaker

Transfer Selector Switches ECAM Control Panel

Figure 4

Instrument Panel

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31-50

CENTRAL WARNING SYSTEMS

Modul: EIS Displays ECAM DISPLAYS ( Electronic Centralized Aircraft Monitoring ) Engine / Warning Display ( E/WD ) The Engine / Warning Display is normally presented on the upper ECAM display unit. It presents engine parameters, fuel on board, flap / slat position. The lower part is dedicated to Warning and Memo messages. Two symbols can be displayed: STS

indicates that the Status Page is not empty. For the Status Page on the lower ECAM display see following pages.

ADV

indicates an Advisory when the ECAM is in Mono display. Mono display mode means that only one ECAM display is operating.

System Display ( SD ) 1. Cruise Page The Cruise Page is displayed during cruise on the ECAM System Display. It comprises information from the Engine and Air Pages.

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ENGINE 5

5 81.5 5 670 92.5 3500

N1 % EGT c N2 % F.F KG/H

IGNITION SEAT BELTS NO SMOKING

F.USED

81.4

FLX

5

84.6 % 35 oC

1530

KG

115

OIL QT

FOB: 18000 KG

665

1560

VIB 0.8

( N1 ) 0.9

115

VIB 1.2

( N2 ) 1.3

S FLAP F 92.7 3

AIR

APU AVAIL

CKPT 20

3500 ADV

LDG ELEV AUTO FWD oC AFT 22 23

TAT +19 C SAT +18 C

STS

2500

FT

CAB V/S

FT/MN > O CAB ALT FT 3050 GW 60300 KG

23 H 56

A 320-211 Engine / Warning Display (on upper ECAM Display)

Figure 5

Cruise Page (on lower ECAM Display)

E/WD and Cruise Page

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System Display (continued) 2. Status Page The Status Page contains the summary of the aircraft operational status after a failure. This page is also automatically called when slats > 2 ( in approach ).

3. Permanent Data Permanent Data is displayed at the bottom of the System or Status Display. Total Air Temperaure ( TAT ) and Static Air Temperature ( SAT ) are displayed in green. The Greenwich Mean Time ( GMT ), synchronized with Cockpit Clock is displayed in green. The Gross Weight ( GW ) is shown in green. It is inhibited before Flight Phase 2 and after Flight Phase 9. Two items of information can be displayed one at a time on the display above GMT: The Load Factor ( G LOAD ) is displayed in amber when the value is out of limits ( above 1.4 g or below 0.7 g ). The Altitude selected on the FCU is displayed in green when the metric unit is selected, provided G LOAD parameter is not displayed.

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Stucture of the Status Page STATUS

CAT 1 ONLY FLAPS SLOW

INOP SYS G+B HYD CAT 2 B ELEC PUMP G RSVR L+R AIL SPLR 1+3+5 L ELEV SLATS AP 1+2 ENG 1 REV NORM BRK NW STEER

CANCELLED CAUTION NAV IR 2 FAULT

MAINTENANCE APU AIR COND

APPR PROC DUAL HYD LO PR .IF BLUE OVHT OUT: -BLUE ELEC PUMP....AUTO .IF GREEN OVHT OUT: -GREEN ENG 1 PUMP..ON -PTU..............AUT O -L/G........GR VTY EXTN -LDG SPD INCREM.....10 KT -LDG DIST...........X 1.8

TAT +19 c SAT +18 C

G LOAD 1.5 23 H 56

STATUS INOPERATIVE SYSTEMS APPROACH PROCEDURES PROCEDURES ( blue ) LIMITATIONS ( green ) CANCELLED CAUTIONS

GW 60300 KG

MAINTENANCE

Permanent Data ( on lower part of lower ECAM display )

Status Page (on lower ECAM Display)

Figure 6

Status Page and Permanent Data

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Modul: ECAM DISPLAYS STRUCTURE OF THE ENGINE/WARNING DISPLAY The upper part of the E/WD is dedicated to the Engine Control parameters. On the right part of the EW/D the fuel on board is indicated. A word, a number and a specific symbol indicates the Slats and Flaps position. On the lower part of the E/WD the Memo Messages present the aircraft systems or functions temporary selected. Normal check lists like Take Off or Landing are displayed. As soon as a failure is detected a special message is displayed which has priority on the Memo Page.

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PRIMARY ENGINE PARAMETERS FUEL SLATS AND FLAPS WARNINGS AND MEMO

ENG 1 FIRE -THR LEVERS........IDLE -P ARKING BRK.........ON -ENG MASTER 1.......OFF -ENG 1 FIRE P/B ...PUSH -AGENT 1..........DISCH -ENG MASTER 2.......OFF -ACT VHF1........NOTIFY

STATUS AND SYSTEM PAGE

overflow arrow

PERMANENT DATA

MEMO MESSAGES: -A/C system functions selected -or check lists. WARNING AND CAUTION MESSAGES Note: An overflow arrow appears when the text of warning and caution messages exceeds the capacity of the display.

Figure 7

Engine / Warning Display

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SYSTEM DISPLAY EXAMPLE The Status and System Page presents one of the twelve aircraft System Pages. Here you can see the hydraulic one.

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PRIMARY ENGINE PARAMETERS FUEL SLATS AND FLAPS WARNINGS AND MEMO

HYD GREEN

0

BLUE

0

PSI

YELLOW

PSI

0

PTU

ELEC OVHT

RAT LO

LO

LO


PERMANENT DATA

TAT +19 C SAT +18 C

23 H 56

G.W. 60300 KG C.G. 28 1 %

Hydraulic Page

Figure 8

System Display Example: Hydraulic Page

______________________________________________________________________________________________________________________________________________________________________________________________



Page 562

MNG TECHNIC



STATUS PAGE EXAMPLE After a failure the Status and System Page provides the operational summary of the aircraft systems. The left part of the Status Page displays in blue the limitations and the postponable procedures, in green the landing capability and some reminder information. The cancelled cautions are diplayed at the bottom. The right part indicates the inoperative systems and the maintenance status. Note: When the Status Page disappears, a message STS appears on the Engine and Warning Display to indicate that the Status Page is no more empty. On the lower part of the Status and System Page some data are displayed.

______________________________________________________________________________________________________________________________________________________________________________________________



Page 563

MNG TECHNIC


PRIMARY ENGINE PARAMETERS FUEL SLATS AND FLAPS

WARNINGS AND MEMO

STATUS MIN RAT SPD.........155 KT APPR PROC DUAL HYD LO PR .IF BLUE OVHT OUT: -BLUE ELEC PUMP....ON -L/G............GR VTY EXTN -LDG SPD INCREM......10 KT -LDG DIST............X 1.8


CAT 1 ONLY FLAPS SLOW CANCELLED CAUTION NAV IR 2 FAULT

TAT +19 C SAT +18 C

PERMANENT DATA

Figure 9

23 H 56

INOP SYS G+B HYD CAT 2 B ELEC PUMP G RSVR L+R AIL SPLR 1+3+5 L ELEV SLATS AP 1+2 ENG 1 REV NORM BRK NW STEER

MAINTENANCE APU AIR COND

G.W. 60300 KG C.G. 28 1 %

Status Page Example

______________________________________________________________________________________________________________________________________________________________________________________________



Page 564

MNG TECHNIC



OVERFLOW, STATUS AND ADVISORY INDICATION Overflow Indication When the text of warning/caution messages exceeds the capacity of the display, a green overflow arrow appears below the grey stripe. This arrow concerns only warning messages and does not deal with memo messages. This arrow remains displayed on the screen as long as there are texts still waiting for display.

STS Message STS appears below the grey ribbon at the same location as the overflow arrow provided that this one is away. This message indicates to the pilot that the status page is no more empty.

ADV Message ADV appears only in single display configuration to signal to the pilot that a parameter drifts in an A/C system. As the corresponding system page cannot be displayed on the lower ECAM display unit, the pilot has to fetch the information on the ECAM control panel: the associated pushbutton switch flashes to indicate which system is concerned.

______________________________________________________________________________________________________________________________________________________________________________________________



Page 565

MNG TECHNIC


Status Reminder

Avisory Reminder

Overflow Indication AMM 31-66-00

Figure 10

E/WD with ADV, STS and Overflow Indication

______________________________________________________________________________________________________________________________________________________________________________________________



Page 566

MNG TECHNIC



ECAM COLOUR PHILOSOPHY The colour of the indications and messages shown on the Engine / Warning Display and on the System Display depends on their meaning: RED Faults and flight situations which require immediate action. AMBER Faults an flight situations which require attention, but no immediate action. GREEN Aircraft and system indications within normal range. WHITE Titles and remarks. BLUE ( CYAN ) Instructions and limitations. MAGENTA Special messages.

______________________________________________________________________________________________________________________________________________________________________________________________



Page 567

MNG TECHNIC



5

5 81.5 5 670

N1 % EGT c

N2 % F.F 5070 LBS/H 92.5

81.4 5

FLX

84.6% 35 oC

FOB: 39600 LBS

Blue (Cyan)

White Amber

White Blue (Cyan)

665 S FLAP F 92.7

Blue (Cyan)

3 5070

Red

ENGINE 2 FIRE

Blue

-THR LEVER 2.................IDLE -ENG MASTER 2...............OFF -AGENT 2........AFT 10S DISCH

Amber

Green

White

White 3 S

Figure 11

ECAM Colour Philosophy

______________________________________________________________________________________________________________________________________________________________________________________________



Page 568

MNG TECHNIC



ATTENTION GETTERS A set of attention getters is installed in front of each pilot. It consists of a Master Warning light and a Master Caution light. When a warning occurs, the Master Warning light flashes continuously. The crew may cancel it, in most cases, as well as the associated aural warning by pushing the Master Warning Light. When a caution occurs, the Master Caution light comes on and stays on associated with a single chime. The Master Caution Light extinguishes when it is pushed.

______________________________________________________________________________________________________________________________________________________________________________________________



Page 569

MNG TECHNIC


Master Warning Light

Master Caution Light

Figure 12

Attention Getters

______________________________________________________________________________________________________________________________________________________________________________________________



Page 570

MNG TECHNIC



ALERT LEVELS Messages have different classes and levels. The level depends on the importance of the message for flight safety. Level 3: warnings ( highest priority ) Level 2: cautions Level 1: cautions Status messages LEVEL 3 Level 3 messages ( warnings ) have highest priority. Level 3 warnings are caused by real emergency situations which require action by the crew. Typical causes for level 3 warnings are Aircraft in dangerous flight situation ( e. g. stall or overspeed ) System faults which concern safety ( e. g. excessive cabin altitude or engine fire ) Level 3 warnings are connected with an aural warning ( continous repetitive chime or special call out ) and the flashing master warning light. If there is a system page for the system concerned it will be displayed on the system display.

LEVEL 1 Level 1 messages point to a system which is faulty but not directly necessary for the flight, e. g. PSCU 1 fault or EFIS DMC 3 fault. Level 1 messages appear on the E/WD without any chime. If there is a system page for the system concerned it will be displayed on the system display.

LEVEL 2 A level 2 message is given when a system fault does not directly affect flight safety. It requires awareness of the crew, but no immediate action. A typical level 2 message is ”IDG 1 OIL LO PR”. Level 2 messages are connected with a single chime and the master caution light. If there is a system page for the system concerned it will be displayed on the system display.

______________________________________________________________________________________________________________________________________________________________________________________________



Page 571

MNG TECHNIC


flashing

Level 3

Level 2

Level 1

Figure 13

Class 1 Failures

______________________________________________________________________________________________________________________________________________________________________________________________



Page 572

MNG TECHNIC



FAILURES WITHOUT ECAM WARNINGS Failures which cannot be left uncorrected until the next scheduled maintenance check appear on the right side of the status page. They do not cause any warning or caution message. On ground, after engine shut down STS appears on the E/WD when class 2 failures are stored. For further information the crew has to select the status page manually. The faults are listed under the heading ”MAINTENANCE”.

______________________________________________________________________________________________________________________________________________________________________________________________



Page 573

MNG TECHNIC


Status

Figure 14

Class 2 Failure

______________________________________________________________________________________________________________________________________________________________________________________________



Page 574

MNG TECHNIC



TYPES OF FAILURES The ECAM indicates three types of failures. Independent Failure Primary Failure Secondary Failure The failure type is indepedent of the alert level.

Independent Failure An independent failure is a failure of a LRU or a system which does not concern any other system. Example: Flight Warning Computer 1 failure. Independent failures are indicated on the left side of the E/WD with their title underlined.

Primary Failure A primary failure is a failure of a LRU or a system which concerns other systems. Example:A failure of the blue hydraulic system will result in the failure of some spoilers. Primary failures are indicated on the left side of the E/WD in a frame.

Secondary Failure A secondary failure is a result of a primary failure. Example:F/CTL ( some spoilers fail if the blue hydraulic system is lost ). Secondary failures are indicated on the right side of the E/WD.

______________________________________________________________________________________________________________________________________________________________________________________________



Page 575

MNG TECHNIC


Primary Failure (red or amber)

Independent Failure (red or amber)

System Pages (amber) corresponding to Secondary Failure

Figure 15

Types of Failures

______________________________________________________________________________________________________________________________________________________________________________________________



Page 576

MNG TECHNIC



Modul: AUDIO WARNINGS AURAL WARNINGS General This is the list of the various audio signals generated by the Flight Warning computers and the manner of cancellation. All aural warnings may be cancelled by pressing the EMER CANC pushbutton on the ECAM Control Panel.

Warning and Callouts Generation All Aural Warnings and Synthetic Voice Callouts are generated in both FWCs. By means of a discrete audio synchronization signal between the FWCs the audio signals are synchronized, this means that the ”faster” FWC suppresses the other one.

______________________________________________________________________________________________________________________________________________________________________________________________



Page 577

MNG TECHNIC



AUDIO WARNINGS AURAL WARNING

WARNING

SINGLE CHIME CONTINUOUS REPETITIVE CHIME

amber warning

automatic

red warning

EMER CANCEL MASTER WARN.

AP disconnection

CAVALRY CHARGE

CANCELLATION

MASTER WARN.or second push on take over pb

land.capability change

automatic

CRICKET

stall

nil

BUZZER

call (SELCAL or cabin)

reset on ACP

TRIPLE CLICK

AUTO CALL OUT

radio height

C CHORD

altitude alert ACP:Audio Control Panel Figure 16

List of Aural Warnings

______________________________________________________________________________________________________________________________________________________________________________________________



Page 578

MNG TECHNIC



FLIGHT PHASES A flight is divided into ten flight phases. The phases depend on parameters shown on the drawing below. The flight phases have influence on internal BITEs ECAM warning inhibits automatic selection of system pages.

______________________________________________________________________________________________________________________________________________________________________________________________



Page 579

MNG TECHNIC



Figure 17

Flight Phases

______________________________________________________________________________________________________________________________________________________________________________________________



Page 580

MNG TECHNIC



AUTOMATIC MODE Flight Phases and System Pages Generally, the System Pages are selected automatically according to the Flight Phase. The Flight Phases are generated in each Flight Warning Computer ( FWC ) and transmitted to the Display Management Computers ( DMCs ) which display the System Page according to the momentary Flight Phase. For example, when the aircraft is supplied with electrical power (Flight Phase 1) the Door Page is displayed. When Flight Phase 2 starts the Wheel Page is displayed: after Engine Start the crew has to know if the Aircraft is ready for taxiing , i. e. if the wheels are ok. If the APU is switched on during Flight Phase 1 or 2 the APU Page has higher priority and will be displayed. This System Page selection mode is called Automatic Mode. In some documents it is also called Flight Phase Mode. A manual page call can replace the current display at any time.

______________________________________________________________________________________________________________________________________________________________________________________________



Page 581

MNG TECHNIC


AMM 31-60-00

Figure 18

Flight Phase Explanation

______________________________________________________________________________________________________________________________________________________________________________________________



Page 582

MNG TECHNIC



Modul: ECAM FLIGHT DISPLAYS GENERAL A manual page call can replace the current display at any time. The APU or engine system pages are displayed in priority if they are started. They remain displayed 10 sec. after APU AVAIL or at the end of ENG START. The flight phases are computed by the FWCs.

Flight Phase 1 Door / oxygen Page. This Page appears as soon as the aircraft is supplied with electrical power. Flght Phase 2 The wheel page is displayed only when engine start has been completed. The flight control page replaces the wheel page for 20 sec. when either sidestick is moved or when rudder deflection is above 22 degrees. Flight Phases 3-5 Engine Page. During this phase, most warnings are inhibited. ”TO. INHIBIT” is displayed on the E/WD. Flight Phase 6 The cruise page is only displayed in flight. It contains both engine and air information. The cruise page appears as soon as slats are in and the engines are no longer at take off power. It disappears when the landing gear is selected down ( wheel page back ). The ”T.O. INHIBIT” message disappears. Flight Phases 7 and 8 Wheel page. During this phase, most warnings are inhibited. ”LDG INHIBIT” is displayed on the E/W display. Note: Ground spoilers are displayed extended only after touch down. Flight Phase 9 Wheel page. The ”LDG INHIBIT” message disappears. Flight Phase 10 Door / oxygen page. Five minutes after the 2nd engine shutdown, the Flight Warning Computers start a new flight leg in phase 1.

______________________________________________________________________________________________________________________________________________________________________________________________



Page 583

MNG TECHNIC



ECAM SYSTEM DISPLAY

OR STATUS PAGE 5MN AFTER

2ND ENG SHUT DOWN

80 KTS

TOUCH DOWN

800 FT

1500 FT

LIFT OFF

80 KTS

1ST ENG TO PWR

1ST ENG STARTED

ELEC PWR

SYSTEM PAGE

TAT +19 c SAT +18 C

23 H 56

GW 132000 LBS

FLIGHT PHASE NUMBER

Figure 19

Flight Phases (general)

______________________________________________________________________________________________________________________________________________________________________________________________



Page 584

MNG TECHNIC


ECAM SYSTEM DISPLAY DOOR OXY

OXY 1850 PSI

CARGO SLIDE

5MN AFTER

2ND ENG SHUT DOWN

80 KTS

TOUCH DOWN

800 FT

1500 FT

LIFT OFF

80 KTS

1ST ENG TO PWR

1ST ENG STARTED

ELEC PWR

CABIN

SLIDE

CABIN 23 H 56

GW 132000 LBS T

TAT +19 c SAT +18 C

Flight Phase 1

Figure 20

Flight Phase 1

______________________________________________________________________________________________________________________________________________________________________________________________



Page 585

MNG TECHNIC


ECAM SYSTEM DISPLAY

ECAM SYSTEM DISPLAY

ENGINE

WHEEL 2530

F.USED 2525 LBS

20 OIL 20 0 115 QT 0 115 10 1

oC

12 REL 2

13 3

oC

12 REL 4

23 H 56

GW 132000 LBS

O.O

O.O

VIB ( N2 ) O.O

O.O

100 PSI 100 0 55 0 55 oC 52 50 NAC oC 20 20

AUTO BRK MAX

TAT +19 c SAT +18 C

VIB ( N1 )

TAT +19 c SAT +18 C

Flight Phase 2

23 H 56

GW 132000 LBS

Flight Phase 3-5

Figure 21

Flight Phase 2 and 3-5

______________________________________________________________________________________________________________________________________________________________________________________________



Page 586

MNG TECHNIC


ECAM SYSTEM DISPLAY ENGINE F.USED

115 AIR 5MN AFTER

2ND ENG SHUT DOWN

80 KTS

TOUCH DOWN

800 FT

1500 FT

LIFT OFF

80 KTS

1ST ENG TO PWR

1ST ENG STARTED

ELEC PWR

1530

( N1 )

0.8

0.9

LBS 1530 OIL

QT

VIB

( N2 )

1.2

1.3

115

LDG ELEV AUTO 2500 FT

o CKPT FWD F AFT 70 70 72 72 TAT +19 c SAT +18 C

VIB

CAB V/S FT/MN O CAB ALT FT

23 H 56

7500

GW132000 LBS

Flight Phase 6

Figure 22

Flight Phase 6

______________________________________________________________________________________________________________________________________________________________________________________________



Page 587

MNG TECHNIC


ECAM SYSTEM DISPLAY

ECAM SYSTEM DISPLAY

D O O R / OXY

WHEEL

OXY 1850 PSI

CABIN CARGO 10 1

oC

12 REL 2

13 3

oC

12 REL 4

AUTO BRK MAX

23 H 56

SLIDE

CABIN GW 132000 LBS

TAT +19 c SAT +18 C

23 H 56

GW 132000 LBS T

TAT +19 c SAT +18 C

SLIDE

Flight Phase 7-9

Flight Phase 10

Figure 23

Flight Phases 7-9 and 10

______________________________________________________________________________________________________________________________________________________________________________________________



Page 588

MNG TECHNIC



Modul: ECAM CONTROL PANEL UTILIZATION ECAM CONTROL PANEL OFF / BRT knobs These knobs control the brightness of each ECAM DU. They are associated with an automatic adjustment of the display intensity depending on the changing light conditions. On the OFF position, the EIS is reconfigurated as for a DU failure. If one ECAM Display is lost the remaining one shows the E/WD. A System Display can be shown by pressing and holding the adjacent Synoptic Key on the ECAM Control Panel.

TO CONFIG When pressed a take-off power application is simulated. If the configuration is correct the ”TO CONFIG NORMAL” message is displayed on the E/WD. This test will trigger a warning if the aircraft is not in TO configuration i.e: Slats or Flaps not in TO configuration Pitch trim not in TO configuration Speed brakes not retracted One door not closed Wheel brake overheat One sidestick not operative EMER CANC When pressed: 1 Any present aural warning is cancelled 2. In case of a red warning, the ECAM message remains displayed. Master Warning Lights will be cancelled. 3. In case of an amber caution, the Master Caution and ECAM messages are cancelled for the rest of the flight. The Status Page may be called with the white ”CANCELLED CAUTION” message and the failure title. Note: The EMER CANC inhibition can be manually restored by pressing RCL for more than 3 sec.

ALL When pressed ( and hold ) all the system pages are displayed successively at 1 second intervalls. It also allowes, by succesive pressing, to display all the system pages one after the other and to stop on the desired one. This is particularly useful in case of ECAM control panel failure because the ALL function remains available.

CLR The light in the CLR pushbutton comes on as long as a Warning / Caution message or a Status message is present on the ECAM DU. As long as the light in the CLR pushbutton is on, pressing it will change the ECAM display. The first of the Warning messages will be deleted. Pressing it again will delete the next Warning message. When the last Warning message is deleted Memo messages come back.

RCL When pressed, the Warning / Caution messages which have been cancelled are recalled.

STS When pressed, the Status Page is displayed. If no Status message is present the ”NORMAL” message is displayed during 5 seconds.

______________________________________________________________________________________________________________________________________________________________________________________________



Page 589

MNG TECHNIC


0

ENGINE vib (n1) 0 0 0

f.used lbs

20 oil 0 115 qt

vib (n2) 0 0

20 0 11 5

100 ps 100 0 0 0 0 c 20 20 nac c 20 20

BLEED 24 C C H 50 C LO HI

RAM AIR

C A B P R ELDG S S ELEV V/S FT/MN P PSI 2 UP 0 0 8 0 0 2 DN

APU

1 IP

HP HP

2 IP

Bleed

2000 FT CAB ALT FT

vent

nlet

ELEC BAT 1 BAT 2 28 V DC BAT 28 V 50 A 50 A DC 1 DC 2 DC ESS TR 1 TR 2 28V ESS TR EMER GEN28 V 150 A 150 A AC 1 AC ESS AC 2 GEN1 GEN2 0 % 0 % APU GEN 0 V 0 V 94 % 116 V 0 hz 0 hz 400 hz 20 C dg 22 dg 11 C 20

10 500

0

SYS1

PSI 30 C 28

30 PSI 27 C

Engine

24 C C H 50 C LO HI

safety

extract

pack 1

pack 2

Cabin Pressure

10

DOOR/OXY avail

o 99 5 7 3 58O

CABIN SLIDE AVIONIC

SLIDE

EGT LOW OIL LEVEL

APU

CABIN

Cond

SLIDE

35558 APU ELEC

ELEC LO

LO

LEFT

CTR

R GHT

Hydraulic F/CTL

Fuel

GBY SPD BRK

L AIL BG

SLIDE CABIN

FUEL LO PR EMER EXIT

0

F U E LBS L F.USED2 0 FOB

1552 12193 8068 12193 1552 +11c c +10 +10 c c +11

WHEEL

OXY 1850 PSI

F.USED1 0

yellow PSI ptu

0

RAT LO

CARGO

FLAP OPEN .C

PSI

AVIONIC BLEED 45PSI

N

0

HYD blue

Electric

apu APUGEN 72 116 V 4OOHZ

green

20 c 20 1 rel2

EMER EXIT CARGO BULK SLIDE CABIN

SLIDE

20 c 20 3 rel4

l ELEV BG

auto brk med

Door/Oxygen

ELAC 1

Wheel

2

SEC1

2

PITCH TRIMGY up RUD GBY

R AIL GB

3 R elev YB

Flight Control

ECAM TO CONFIG

UPPER DISPLAY

OFF

ENG

BLEED

PRESS

ELEC

APU

COND

DOOR

WHEEL

STS

RCL

HYD

FUEL

F/CTL

ALL

BRT

LOWER DISPLAY

OFF

EMER CANC

BRT

CLR

Figure 24

CLR

ECAM Control Panel

______________________________________________________________________________________________________________________________________________________________________________________________



Page 590

MNG TECHNIC



SINGLE DISPLAY UNIT OPERATION

ADVISORY MODE The value of some critical system parameters is monitored by an advisory mode. When the value drifts from its normal range, the corresponding ECAM page is displayed and the affected parameter pulses. For example the PRESS page will be displayed if the cabin pressure increases above its normal value, but is still well below the threshold of the warning. In this case the crew may revert to manual pressure control and prevent warning activation. Note: an advisory may or may not lead to a failure. They are totally independent one from the other.

If a Parameter drifts out of normal Range when there is only one ECAM Display Unit available ( Single Display Unit Operation ), a white ADV message pulses at the bottom of the Engine Warning Display to attract crew attention. As the corresponding system page cannot be displayed automatically on the SD, the pilot has to fetch the information on the ECAM control panel: the associated key light flashes to indicate which system is concerned. The System Pages can be selected only manually by pressing and holding the concerned pushbutton.

FAILURE MODE If a failure occurs which is important to be indicated the corresponding ECAM page is displayed. In the case of two or more failures occuring the same time the FWC decides which one is to be displayed first. With the Failure Mode the pilots automatically get the indications they need in the case of a failure.

______________________________________________________________________________________________________________________________________________________________________________________________



Page 591

MNG TECHNIC


5

5

81.5

N1 %

670

EGT c

5

92.5 5070

N2 % F.F KG/H

IGNITION SEAT BELTS NO SMOKING

FLX

81.4 5

84.6 % 35 oC

FOB: 39600 LBS 665 S FLAP F 92.7 3

5070 ADV

APU AVAIL

STS A 320-211 Figure 25

ADV - Reminder in Single Display Unit Operation

______________________________________________________________________________________________________________________________________________________________________________________________



Page 592

MNG TECHNIC



Modul: ECAM WARNINGS FAILURE PROCEDURE EXAMPLE Normal configuration We are in flight in normal configuration. The ECAM displays indicate that all is correct. The cruise page is displayed.

Note: The Cruise Page is only displayed in Flight and can not be selected manually.

______________________________________________________________________________________________________________________________________________________________________________________________



Page 593

MNG TECHNIC


GREEN

BLUE

PTU

RAT MANON

H Y D 4

6 8 6 8 10 N 1 4 10

2

ENG1 PUMP

ELEC PUMP A FAULT U T OFF O

FAULT OFF

YELLOW

40VU

A FAULT U ELEC PUMP T OFF O ENG2 PUMP H FAULT Y FAULT ON D OFF

2

EGT 5 10 C 490 N2 89 89 F.F 2900 LBS/H 2900 5 10 490

SEAT BELTS NO SMOKING

FOB

LBS

FLAP

ENG A.ICE

VIB F.FLOW

6630 16

OIL

6630 16

LDG ELEV CKPT FMT AFT 7O 72 73 69 TAT +19 .C SAT +18 .C

0.8

23 H 56

0.9

VIB

1. 2

1. 3

AUTO 5OO V/S FT/MN O CAB ALT FT 4150 G.W

132000

TO CONFIG

LBS

ENG

BLEED

PRESS

ELEC

APU

COND

DOOR

WHEEL

CLR

Figure 26

EMER CANC

STS

RCL

HYD

FUEL

F/CTL

ALL

CLR

Failure Example part 1

______________________________________________________________________________________________________________________________________________________________________________________________



Page 594

MNG TECHNIC



Green Hydraulic System has low level An aural warning, the single chime, and a visual warning, the Master Caution, attract your attention. The Engine and Warning Display indicates the title of the failure and the actions to be taken. On the Status and System Display, the Hydraulic Page is called automatically. The CLR pushbuttons come on and as long as the failure is not cleared, they stay on. On the Hydraulic Panel FAULT lights come on, indicating the pushbuttons to release out. The first action you have to do is to press the Master Caution pushbutton.

______________________________________________________________________________________________________________________________________________________________________________________________



Page 595

MNG TECHNIC


GREEN

BLUE

PTU

RAT MANON

H Y D 4

6 8 6 8 10 N 1 4 10

2

EGT 5 10 C 490 N2 89 89 F.F 2900 LBS/H 2900

FOB

OFF


LBS

FLAP MASTER CAUT

Master Caution, FAULT lights and CLR pushbutton on ECAM Control Panel illuminate.

HYDG RSVR LO LVL -PTU.............OFF -GREEN ENG1PUMP..OFF

Hydraulic Page appears


FAULT

40VU

2

5 10 490

Caution Message

ENG1 PUMP

YELLOW

3000 PSI 3000 PSI 3000

TO CONFIG TAT + 19 C .C 23 H 56 GW 102000 LBS SAT + 18 C

ENG

BLEED

PRESS

ELEC

APU

COND

DOOR

WHEEL

CLR

Figure 27

EMER CANC

STS

RCL

HYD

FUEL

F/CTL

ALL

CLR


______________________________________________________________________________________________________________________________________________________________________________________________



Page 596

MNG TECHNIC



On this picture, the Master Caution pushbutton has been pushed, the light is out. Now you have to perform the actions indicated on the Engine / Warning Display. At first, the PTU has to be switched off. Then the green engine 1 pump has to be switched off.

______________________________________________________________________________________________________________________________________________________________________________________________

FRA Revision No : 02TS 84 rw 25.4.94 Issue Date : 21/05/2013


Seite: Page: 52

Page 597

MNG TECHNIC


GREEN

BLUE RAT MANON

ENG1 PUMP

H Y D 4

OFF

6 8 6 8 10 N 1 4 10

2


FAULT FAULT

YELLOW

PTU

40VU

A FAULT U FAULT ELEC PUMP T OFF O ENG2 PUMP H FAULT Y FAULT ON D OFF

2

EGT 5 10 C 490 N2 89 89 F.F 2900 LBS/H 2900 5 10 490

FOB

LBS

FLAP MASTER CAUT

HYDG RSVR LO LVL -PTU.............OFF -GREEN ENG1PUMP..OFF

HYD G RSVR LO LVL -PTU...............OFF -GREEN ENG 1 PUMP..OFF

3000 PSI 3000 PSI 3000

TO CONFIG TAT + 19 C .C 23 H 56 GW 102000 LBS SAT + 18 C

ENG

BLEED

PRESS

ELEC

APU

COND

DOOR

WHEEL

STS

RCL

CLR

Figure 28

EMER CANC

HYD

FUEL

F/CTL

ALL

CLR


______________________________________________________________________________________________________________________________________________________________________________________________



Page 598

MNG TECHNIC



On this picture, the corrective actions have been taken. All the FAULT lights are off. On the Engine / Warning Display, the messages associated with the corrective action have disappeared. On the left hand side of the Engine / Warning Display, the result of the failure appears indicating that it is a primary failure. On the right hand side, the secondary failures are displayed. The next action to be done is to press a CLR pushbutton. By this means you get information about the first secondary failure. In this example, the Flight Control Page will be displayed.

Seite: Page: 54 ______________________________________________________________________________________________________________________________________________________________________________________________ Revision No : 02 Issue Date : 21/05/2013


Page 599

MNG TECHNIC



GREEN

BLUE RAT MANON

ENG1 PUMP

H Y D 6 4

OFF

8 6 8 10 N 1 4 10

2


FAULT

40VU


2

EGT 5 10 C 490 N2 89 89 F.F 2900 LBS/H 2900 5 10 490

HYD G RSVR LO LVL G SYS LO PR

FOB

LBS

FLAP MASTER

PRIMARY FAILURE

* F CTL * WHEEL

SECONDARY FAILURES

HYD G RSVR LO LVL G SYS LO PR 0

YELLOW

PTU

PSI

CAUT

* F/CTL * WHEEL

3000 PSI 3000

TO CONFIG

TAT + 19 C .C 23 H 56 GW 102000 LBS SAT + 18 C

ENG

BLEED

PRESS

ELEC

APU

COND

DOOR

WHEEL

CLR

Figure 29

EMER CANC

STS

RCL

HYD

FUEL

F/CTL

ALL

CLR


______________________________________________________________________________________________________________________________________________________________________________________________



Page 600

MNG TECHNIC



On this picture, CLR was pressed once. The failure disappeared from the left hand part of the Engine / Warning Display and Memo Messages came back. The System Page corresponding to the first secondary failure is displayed. The next action to be done is to press a CLR pushbutton again.



Page 601

MNG TECHNIC



GREEN

BLUE

PTU

RAT MANON

H Y D 4

6 8 6 8 10 N 1 4 10

2

ENG1 PUMP


FAULT OFF

YELLOW

40VU


2

EGT 5 10 C 490 N2 89 89 F.F 2900 LBS/H 2900 5 10 490


FOB

LBS

FLAP MASTER CAUT * F CTL * WHEEL

GBY 1

5

1

5

SPD BRK

L AIL

ELAC 1

GB

L ELEV

GB

2

SEC 1

PITCH TRIM GY UP RUD

GBY

R AIL 2

3

GY

R ELEV

TO CONFIG

BY


ENG

BLEED

PRESS

ELEC

APU

COND

DOOR

WHEEL

CLR

Figure 30

EMER CANC

STS

RCL

HYD

FUEL

F/CTL

ALL

CLR


______________________________________________________________________________________________________________________________________________________________________________________________



Page 602

MNG TECHNIC



On this picture, CLR was pressed a second time. The title of the first secondary failure disappeared. The System Page associated with the next secondary failure is displayed. The next action to be done is to press the CLR pushbutton again.



Page 603

MNG TECHNIC


GREEN

BLUE RAT MANON

H Y D 4

6 8 6 8 10 N 1 4 10

2

ENG1 PUMP FAULT OFF

YELLOW

PTU


40VU


2

FOB

EGT 5 10 C 490 N2 89 89 F.F 2900 LBS/H 2900 5 10 490

LBS

FLAP MASTER CAUT


* WHEEL

NW STEER 30 1

30 REL

2

30 AUTO BRK 3

30 REL

4 TO CONFIG

1 5

1

EMER CANC

5


ENG

BLEED

PRESS

ELEC

APU

COND

DOOR

WHEEL

CLR

Figure 31

STS

RCL

HYD

FUEL

F/CTL

ALL

CLR

Failure Examble part 6

______________________________________________________________________________________________________________________________________________________________________________________________



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On this picture, CLR was pressed a third time. The title of the secondary failure disappeared. The Memo message is back on the right hand part of the Engine / Warning Display. The Satus Page is displayed. The next action to be done is to press the CLR pushbutton again.

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GREEN

BLUE RAT MANON

H Y D 4

6 8 6 8 10 N 1 4 10

2

ENG1 PUMP


FAULT OFF

YELLOW

PTU

40VU


2

EGT 5 10 C 490 N2 89 89 F.F 2900 LBS/H 2900 5 10 490

FOB

LBS

FLAP MASTER CAUT ENG A.ICE


STATUS -L/G.....GRVTY EXTN -LDG DST......“*1.2 CAT 2 ONLY SLATS/FLAPS SLOW

TAT + 19 C .C SAT + 18 C

INOP SYS: CAT 3 GREEN HYD SPLR 1+5 L/G RETRACT NW STEER NORM BRK REVERSER 1 YAW DAMPER

23 H 56

TO CONFIG

EMER CANC

ENG

BLEED

PRESS

ELEC

APU

COND

DOOR

WHEEL

HYD

FUEL

F/CTL

ALL

GW 102000 LBS

CLR

Figure 32

STS

RCL

CLR


______________________________________________________________________________________________________________________________________________________________________________________________



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Lufthansa Technical Training~


On this picture the CLR pushbutton was pressed a fourth time. The CLR pushbuttons are off. On the System Display, the cruise page is back. The warning has been cleared. The Statur Reminder STS indicates that the Status Page is not empty. The RCL pushbutton allows the crew to recall warnings.



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GREEN

BLUE RAT MANON

H Y D 4

6 8 6 8 10 N 1 4 10

2

ENG1 PUMP


FAULT OFF

YELLOW

PTU

40VU


2

EGT 5 10 C 490 N2 89 89 F.F 2900 LBS/H 2900 5 10 490


FOB

LBS

FLAP MASTER CAUT ENG A.ICE

STS

F.FLOW

663O 16

OIL

6630 16

LDG ELEV CKPT FWD AFT .f 7O 72 73 69

VIB O. 8 O. 9 VIB

1. 2 1. 3 AUTO 5OO V/S FT/MN O CAB ALT FT

TO CONFIG

EMER CANC

4150


ENG

BLEED

PRESS

ELEC

APU

COND

DOOR

WHEEL

CLR

Figure 33

STS

RCL

HYD

FUEL

F/CTL

ALL

CLR


______________________________________________________________________________________________________________________________________________________________________________________________



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FLIGHT PHASE INHIBITS Some warnings are suppressed during determined Flight Phases. The reason is that during critical flight situations ( e.g. rotation during T. O. ) the crew is not confronted with minor warnings which they can not handle in this situation. For example: If the inertial reference part of ADIRU 1 or ADIRU 2 fails during Takeoff no Master Caution Light and no ECAM Caution Message appears until Flight Phase 5 is finished ( when altitude 1500 feet is reached ). Then the Master Caution Light and a Single Chime come on and an ECAM Caution Message is displayed on the E / WD. Additional, instructions are displayed in blue. On the following three pages there are excerpts from the AMM. Each table contains the Flight Phase Inhibits for a specific failure. The Flight Phases concerning a faliure can be found in the AMM 31-51-XX . XX is for the ATA chapter concerning the system.

Page: Page: 64 ______________________________________________________________________________________________________________________________________________________________________________________________ Revision No : 02 Issue Date : 21/05/2013


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Figure 34

Flight Phase Inhibit ATA 34

______________________________________________________________________________________________________________________________________________________________________________________________



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Figure 35


______________________________________________________________________________________________________________________________________________________________________________________________



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Figure 36


______________________________________________________________________________________________________________________________________________________________________________________________



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ECAM BITE General The components of the Central Warning System are: 2 Flight Warning Computers ( FWC ) 2 System Data Aquisition Concentrators ( SDAC ) 1 ECAM Control Panel For the Indication the two ECAM Display Units are used. The SDACs send their BITE signals to the CFDIU via their own side FWC.



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to FAC 1

from DUs

from DUs

ECAM CP

DMC 1

DMC 3

DMC 2

SDAC 2

SDAC 1

CFDIU

FWC 1

FWC 2

AMM 31-60-00

Figure 37

CFDIU to ECAM Interface

______________________________________________________________________________________________________________________________________________________________________________________________



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Operational Test The Operational Test for the Central Warning Systems can be found in the AMM 31-50-00. There are some preconditions: Release the ENG/FADEC GND PWR 1 & 2 pushbutton switches to ON. Do the ADIRS start procedure. Access to the ECAM pages is the procedure as below: 1. On the MCDU, get the SYSTEM REPORT/TEST INST menu page item. 2. Push the LSK 1L for ECAM 1. The ECAM 1 page comes into view. The ECAM 1 page is for Flight Warning Computer 1, System Data Acquisition Concentrator 1 and the ECAM Control Panel. The ECAM 2 page is for Flight Warning Computer 2, System Data Acquisition Concentrator 2 and the ECAM Control Panel.



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Figure 38

Central Warning Systems Test part 1

______________________________________________________________________________________________________________________________________________________________________________________________



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The next step is to push LSK 4L for GROUND SCANNING. If the system is good, NO FAILURE will be displayed.



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Figure 39

Central Warning Systems Tests part 2

______________________________________________________________________________________________________________________________________________________________________________________________



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31-60

ELECTRONIC INSTRUMENT SYSTEM ( EIS )

GENERAL The EIS (Electronic Instrument System) presents on Display Units (DUs): EFIS (Electronic Flight Instrument System) information, (i.e. flight parameters and navigation data). ECAM (Electronic Centralized Aircraft Monitor) information. The layout of the 6 DUs and the breakdown of the information displayed on them is presented as follows: - 2 DUs are installed on the center instrument panel, one above the other. They display ECAM information. - 2 DUs are installed side by side in front of each pilot. They display flight and navigation data. On each main instrument panel, in normal configuration, the outer DU will be alloted to the Primary Flight Display (PFD) function, and the inner DU to the Navigation Display (ND) function. Each pilot is given the possibility to display ECAM information instead of navigation information on its inner DU, in order to cover failure cases.

Primary Flight Display ( PFD ) The instrument Primary Flight Display (PFD) displays all the primary flight indications necessary for short-term aircraft control.

Navigation Display ( ND ) The instrument Navigation Display (ND) displays the navigation information necessary as the flight progresses, and in the mode chosen by the pilot : Rose, (with NAV, VOR, ILS submodes), ARC, PLAN.

Information presented on EFIS DUs The EFIS DUs enable display of flight path control and navigation data for the crew. Each pilot has two CRT display units at his disposal, one PFD and one ND, on which is displayed the following information: Attitude Horizontal Situation Mach/Airspeed Altitude Vertical Speed Radio Altitude Weather Radar Information Marker Beacon Indication Flight Mode Annunciaton (autopilot/flight director modes).

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Captain’s PFD

Captain’s ND

Engine/Warning Display

FO’s ND

FO’s PFD

System Display ECAM Control Panel

Figure 40

EIS Cockpit Components

______________________________________________________________________________________________________________________________________________________________________________________________



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ARCHITECTURE General The Electronic Instrument System is an avionic system connected with most of the aircraft systems to perform the EFIS and ECAM functions. The EFIS provides flight information and the ECAM provides system and warning information. The Electronic Instrument System comprises seven computers: three identical Display Management Computers ( DMCs ) two identical Flight Warning Computers ( FWCs ) two identical System Data Acquisition Concentrators ( SDACs ). EFIS Each DMC decodes and processes data from the aircraft systems in such a way to generate images on PFDs and NDs. The three DMCs receive identical information.

ECAM Each DMC uses A/C system data which is processed by the System Data Acquisition Concentrators ( SDACs ) and Flight Warning Computers ( FWCs ) before being presented on E/WD and SD. The SDACs digitalize aircraft system data and transmit it to the DMCs. The DMCs using SDACs outputs, generate aircraft system information for system display on the SD.

The DMCs use the outputs of the FWCs to display aircraft information on the lower part of the E/WD ( messages ). The SDACs receive A/C system information concerning amber cautions and transmit it to the FWCs. The FWCs receive A/C system data concerning red warnings and memos, they generate messages, audio signals and activate attention getters.

Control Panels Two EFIS Control Panels, one ECAM Control Panel and Transfer Selector Switches are provided for EFIS and ECAM controls. Control panels are physical interfaces between the crew members and the EIS.

Redundancy A great redundancy between systems is used to minimize the loss of information. Loss of a SDAC, or a FWC, or a DMC does not affect EIS operation. The system still operates normally with one SDAC, one FWC and one DMC inoperative.

Note: the DMCs receive directly A/C system data for display on the upper part of the E/WD.



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Figure 41

EIS Schematic

______________________________________________________________________________________________________________________________________________________________________________________________



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SWITCHING PANEL ADIRS Source Switching If a primary AIR DATA or ATTITUDE/HEADING source (side 1 or side 2) fails, the relevant flags come into view on the associated PFD and ND (Captain for side 1, First Officer for side 2). In this case, the affected pilot can switch over to the ADIRU 3 source to recover the lost flight parameters, by setting the AIR DATA or the ATT HDG selector switch to CAPT ON 3 for side 1, or to F/O ON 3 for side 2. These selector switches are located on the source switching panel which is aft of the ECAM control panel, on the center pedestal. In case of complete failure of the ADIRU 1 (or 2), both selector switches (AIR DATA and ATT HDG) must be set to the switched position CAPT ON 3 (or F/O ON 3).

DMC Switching This selector switch is used for EIS DMC switching. When rotated out of the vertical position, DMC 3 totaly replaces DMC 1 or DMC 2.

ECAM/ND Transfer This transfer facility enables the crew to display an ECAM image to either ND, in case of failure of ECAM DUs.

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AMM 34-11-00

A

AMM 31-60-00

Figure 42

Switching Panel

______________________________________________________________________________________________________________________________________________________________________________________________



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EFIS CONTROL PANEL The EFIS control panels, which are part of the Flight Control Unit, are divided into PFD controls and ND controls. PFD Controls Both EFIS Control Panels have the same controls: Hg / hPa Selector Knob for QNH units ( in Hg or hPa ). BARO Setting Knob for QNH setting. Pulling the knob selects the standart value. In this case, STD is displayed in the Baro Reference Window. FD Pushbutton Switch for switching of FD Bars resp. FPD-Symbols. LS Pushbutton Switch for switching of LOC and GLIDE Scales and deviation symbols on the PFD. Baro Reference Window is used to display the pressure reference value and the reference used.

ND Controls For the Navigation Display there ara following controls: Mode Selector Switch to use the ND in different modes. In the Rose Modes the aircraft symbol is in the middle of the ND, in the ARC Mode it is at the bottom of the DU. PLAN corresponds to a map displayed on the ND. ENG is a standby mode to display engine parameters on the ND. Scale Selector Switch to selesct the range on the ND. for example, if ”320” is selected the distance between aircraft sysmbol and compass rose corresponds to a distance of 320 nautical miles. ADF/VOR selector: enables ADF or VOR bearing pointers to be selected on the associated ND as well as the corresponding navigation station characteristics in any mode except PLAN mode. Data base display P/Bs: these five P/Bs enable additional data to be displayed on the ND. When pressed these P/Bs respectively display Airports, ADF stations, VOR/DME stations, Waypoints and Constraints.



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Mode Selector Switch

Data Base Display Pushbuttons

BARO Correction Display

Scale Selector Switch

Hg / hPa Selector

ILS

BARO Setting Knob

FD

ILS

ADF / VOR Selector Switches FD Pushbutton Switch

LS Pushbutton Switch

______________________________________________________________________________________________________________________________________________________________________________________________


Figure 43 EFIS Control (Captain) A319/320/321 / TRAININGPanel MANUAL

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LOCATION Flight Warning Computers ( FWCs ) System Data Acquisition Concentrators ( SDACs ) Display Management Computers ( DMCs )

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Figure 44

EIS Location: SDACs, FWCs and DMCs

______________________________________________________________________________________________________________________________________________________________________________________________



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Figure 45

Circuit Breakers (part 1)



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Figure 46

Circuit Breakers (part 2)


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31-61-00 Figure 47

PFD-ND Relay Boxes Location

Seite: ______________________________________________________________________________________________________________________________________________________________________________________________ Revision No : 02 Issue Date : 21/05/2013


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Figure 48

Test Plug Location AMM 31-69-00



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31-64

PRIMARY FLIGHT DISPLAY (PFD)

Modul: PFD GENERAL PRESENTATION Primary Flight Display Structure A grey background is displayed on speed, heading and altitude PFD windows. In case of avionics Ventilation Blower and Extract Fault, the grey background is supressed in order to limit PFD tubes consumption and to prevent them from overheating.

Attitude The aircraft attitude is shown on the central part of the display by a cutsphere shaped window which features a conventional Attitude Display Indicator.

Heading Actual and Selected Heading or Track Information is shown at the bottom of the display.

Speed The Airspeed Scale on the left hand side contains all the data of a conventional Airspeed Indicator plus significant limit protections and Target Speed.

Guidance Flight Director Bars or Flight Director Symbol display guidance orders on the attitude sphere.

Altitude The Altitude Scale on the right side displays the aircraft actual Altitude according to the selected baro setting reference.

Trajectory Deviation Lateral and vertical scales provide trajectory deviation information during an ILS or R NAV approach.

Vertical Speed A green pointer and a numerical value display the aircraft Vertical Speed at the extreme right of the Primary Flight Display.

Flight Mode Annunciator Annunciations and messages regarding Flight Management and Guidance System operation are displayed at the top of the PFD which is devided into 5 columns and 3 lines.



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______________________________________________________________________________________________________________________________________________________________________________________________ Figure 49 PFD Structure



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31-65

NAVIGATION DISPLAY

General The indication on the Navigation Display ( ND ) depends on the position of the Mode Selector Switch on the EFIS Control Panel. The different Modes are shown on the following pages.

ROSE ILS Mode This Mode corresponds to the conventional HSI with Localizer- and Glideslope - Indication. ROSE VOR Mode This Mode corresponds to the conventional HSI with VOR Course and VOR Deviation. ROSE NAV Mode This Mode corresponds to the conventional HSI without VOR Course resp. VOR Deviation but with Flight Plan indications ( from the FMGES) and Weather Radar Indication.

ARC Mode This mode shows a sector of 90 degrees in front of the aircraft. The aircraft symbol is at the bottom of the DU. Flight Plan information and Weather Radar are displayed like in the ROSE NAV mode. PLAN Mode In this mode a map is displayed with North up. In the middle of the display there is a waypoint as a reference point. Depending on the selected range other waypoints and the planned Flight Path are displayed. An aircraft symbol is displayed according to the Present Position.



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Figure 50 ND - ROSE ILS Mode and ROSE VOR Mode ______________________________________________________________________________________________________________________________________________________________________________________________ Revision No : 02 Issue Date : 21/05/2013


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LS

LS

Page: Page: 92 Figure 51 ND - ROSE NAV Mode and ARC Mode ______________________________________________________________________________________________________________________________________________________________________________________________ Revision No : 02 Issue Date : 21/05/2013


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LS

Figure 52

ND - PLAN Mode



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31-50

CENTRAL WARNING SYSTEMS

FLIGHT WARNING SYSTEM STRUCTURE

ALERT SIGNAL FLOW

General On the schematic, all components of the Flight Warning System are shown.

Level 1 When a SDAC receives a failure signal which should result in a level 1 warning it sends a signal to the FWCs. The FWCs generate the caution message for the Engine/Warning Display.

ECAM Control Panel The ECAM Control Panel indirectly controls the ECAM Display Units: via the DMCs for system display selection via the FWCs for the management of messages. When messages are cleared manually, the system display can show an other synoptic or the status page automatically.

DMC Inputs The DMCs receive all data for the Engine part of the Engine/Warning Display directly from the associated systems.Some data for system synoptics are received directly, most data via the SDACs.

Level 2 Most failure signals for level 2 warnings have the same signal flow and result in a caution message. The FWCs additionally illuminate the master caution lights and they generate a single chime. Not all failure signals flow through the SDACs to the FWCs: Some systems are directly connected to the FWCs.

Level 3 All failure signals for level 3 warnings are directly connected to the FWCs. The FWCs generate the warning message for the Engine/Warning Display, illuminate the master warning lights and generate an aural warning.



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AMM 31-50-00

Page: Figure 53 ECAM Structure ______________________________________________________________________________________________________________________________________________________________________________________________ Revision No : 02 Issue Date : 21/05/2013


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31-56

ECAM CONTROL PANEL (ECP)

Description The ECAM Control Panel works digitaly. However, some pushbuttons send a discrete signal, so they can be used in the event of an ECP failure. The discrete outputs are: CLR ( Clear ) STS ( Status ) RCL ( Recall ) EMER CANC ( Emergency Cancel ) ALL.

ECP Pushbutton Lights Control The Pushbutton Lights are controlled by DMC 1 and DMC 2 via a digital bus.

Failures If the ECP power supply fails or the ECP is inoperative the analog connected pushbuttons remain operative. The ECAM-system can still be used.

Interface The pushbuttons for signals to the FWCs are: CLR STS RCL EMER CANC TO CONFIG. The FWCs receive BITE information via the output bus of the ECP. On the same bus, they receive information if the TO CONFIG-pushbutton is pushed. The pushbuttons for signals to the DMCs are: System-pushbuttons ( Digital Bus ) ALL.



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FWC 2

FWC 1

FDIU GPWC

DMU

PVI

ECAM Control Panel DMC 1

DMC 3

DMC 2

ASM 31-61-00

Page: Page: 97 Figure 54 ECAM Control Panel Interconnection ______________________________________________________________________________________________________________________________________________________________________________________________ Revision No : 02 Issue Date : 21/05/2013


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31-55

SYSTEM DATA ACQUISITION CONCENTRATOR (SDAC)

GENERAL Description Two convertible SDACs are installed in the A 320. Both SDACs receive the same information. Each SDAC receives digital and analog data and descrete signals directly from various aircraft systems. These data are concentrated by each SDAC, i. e. the numerous analog and digital inputs are processed and offered to different users in ARINC 429 format. Most data for the System Display are processed in the SDACs. Both SDACs also receive data about level 2 and level 3 malfuctions and failures from most aircraft systems. These data are transmitted to the FWCs. The FWCs generate the corresponding Caution Messages and Procedure Messages for the Engine / Warning Display. The software is memorized on an OBRM ( On Board Replacable Module ) so that software modifications can be done quickly.

Failures If one SDAC fails the receiving systems still are supplied by the other SDAC. All functions are preserved. If SDAC 1 fails, the message ” FWS SDAC 1 FAULT ” appears on the E/WD. ( If SDAC 2 fails ”...SDAC 2...” instead of SDAC 1). If both SDACs fail most of the Caution Messages can not be displayed any more. On some System Pages some parameters can not be displayed any more. ” XX ” is displayed instead. Local Warning Lights are still operative on the Overhead Panel. On the E/WD the following message appears: in Amber FWS SDAC 1 + 2 Fault -MONIT OR OVERHEAD PANEL in Blue ECAM ENG FUEL F/CTL in Blue WHEEL SYS PAGES AVAIL in Blue



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Figure 55

SDAC

Page: Page:4 ______________________________________________________________________________________________________________________________________________________________________________________________ Revision No : 02 Issue Date : 21/05/2013


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SDAC INTERFACE Inputs The systems listet below send digital data to the SDACs: CPC 1 & 2(Cabin Pressure Controller) Pack Temperature Controller 1 & 2 BMC 1 & 2 (Bleed Monitoring Computer) SFCC 1 & 2 (Slat/Flap Control Computer) EVMU (Engine Vibration Monitoring Unit) BSCU (Braking/Steering Control Unit) ECB (Electronic Control Box) BCL (Battery Charge Limiter) EGIU 1 & 2 (Electrical Generation Interface Unit) TPIS Det Unit (Tire Pressure Indicating System Detection Unit. This is an option). Various aircraft systems send different kinds of analog inputs to the SDACs: Discretes Frequencies Resistances Synchro signals Resolver signals. Some hundred Discrete/Analog Inputs are listet in the ASM 31-54-00.



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SDAC 1

SDAC 2

SDAC 1

AMM 31-54-00

SDAC 2

AMM 31-54-00

Pageage:1 Figure 56 SDAC Inputs ______________________________________________________________________________________________________________________________________________________________________________________________ Revision No : 02 Issue Date : 21/05/2013


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Outputs The systems listet below receive digital data from the SDACs: DMC 1, 2 & 3 The Display Management Computers receive data to be displayed on the System Display. FWC 1 & 2 The Flight Warning Computers receive data for Caution Messages and BITE information. CIDS Director 1 & 2 The CIDS Directors receive data for door indication from SDAC 1 only. FDIU The Flight Data Interface Unit receives data to be recorded in the Digital Flight Data Recorder. DMU The Data Management Unit receives data to be used in the AIDS. MU ACARS The ACARS Management Unit receives data to be sent via VHF from SDAC 1 only.



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SDAC 1

SDAC 2

AMM 31-54-00

Figure 57

SDAC Outputs



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31-53

FLIGHT WARNING COMPUTER (FWC)

GENERAL Description Two convertible FWCs are installed. Both FWCs receive the same information. The FWCs work on following tasks: Triggering of Master Warning and Master Caution Lights Triggering of AFS Autoland Lights Generating of Warnings and Caution Messages for the E/WD Generating of Procedures associated to failures Generating of the Status Function Generating of Memo Messages Generating of Aural Warnings and Callouts Generating of Flight Phases ( by using input parameters) Triggering of the DMCs which System Page is to be displayed (in automatic mode) Establishing of the interface to CFDIU for the SDACs Comparison of Heading, Attitude and Altitude Indications Calculation of Overspeed Limits Calculation of Stall Limits. Both FWCs receive all data about level 3 warnings ( and some data about level 1 and level 2 malfunctions ) directly from the aircraft systems in analog or digital form depending on the affected system. Most data about level 1 and level 2 malfunctions are received from both SDACs. Each FWC is connected with the opposite FWC via a data bus. Normally, the FWCs use data from SDAC 1. If this one fails SDAC 2 is used. The software is memorized on an OBRM ( On Board Replacable Module ) so that software modificatins can be done quickly.

Failures If one FWC fails the receiving systems still are supplied by the other FWC. Most functions are preserved. If FWC 1 fails, the message ” FWS FWC 1 FAULT ” appears on the E/WD. On captain’s Master Warning and Master Caution Light the ” Master ”-line, and on copilot’s side the ” Warning ” and ” Caution ”-lines can not illuminate any more because the upper and the lower bulb are triggered from different FWCs. ( If FWC 2 fails ”...FWC 2...” instead of FWC 1 and the Attention Getters accordingly reversed. ) If both FWCs fail the consequences are: no Auto Callouts no Aural Warnings no Memo Messages no status page no Master Warning Light no Master Caution Light On the E/WD the following message appears: in Amber FWS FWC 1 + 2 FAULT -MONIT OR SYS in Blue -MONIT OR OVERHEAD PANEL in Blue



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AMM 31-53-34

Page: Page: 105 Figure 58 FWC ______________________________________________________________________________________________________________________________________________________________________________________________ Revision No : 02 Issue Date : 21/05/2013


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FWC 1

FWC 2

FWC 1

AMM 31-52-00

FWC 2

AMM 31-52-00

Page: Page: 107 ______________________________________________________________________________________________________________________________________________________________________________________________ Figure 59 FWC Inputs Revision No : 02 Issue Date : 21/05/2013


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Outputs The systems listet below receive digital data from the FWCs: DMC 1, 2 & 3 The Display Management Computers receive Messages to be displayed on the Lower Part of the E/WD and Status Messages for the Status Page. CFDIU The Centralized Fault Display Interface Unit receives data for BITE information. FDIU The Flight Data Interface Unit receives data to be recorded in the Digital Flight Data Recorder. DMU The Data Management Unit receives data to be used in the AIDS. MU ACARS The ACARS Management Unit receives data to be sent via VHF from FWC 1 only. Descretes trigger the items listet below: Cockpit Loudspeakers Master Warning Lights Master Caution Lights Auto Land Lights

Page Page: 108 ______________________________________________________________________________________________________________________________________________________________________________________________ Revision No : 02 Issue Date : 21/05/2013


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FWC 1

FWC 2

AMM 31-52-00

Figure 60

FWC Outputs Page: Page: 109

______________________________________________________________________________________________________________________________________________________________________________________________



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COMPARISON General Each FWC works as a comparetor. If it detects differences between the displayed parameters a message is generated. The monitored parameters are: Heading Attitude Altitude.

Heading Comparison If the FWC detects a difference between the heading indications, the Message ” CHECK HDG ” appears on the heading scale of the PFD and on the ND. In this case the pilots have to compare the indications with the Standby Compass to do the appropiate ADIRS switching.



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PFD

ND

Figure 61

Heading Comparison Messages

______________________________________________________________________________________________________________________________________________________________________________________________



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Attitude Comparison If the FWC detects a difference between the attitude indications, the message ” CHECK ATT ” appears on the PFD. In this case the pilots have to compare the indications with the Standby Horizon to do the appropiate ADIRS switching.

Altitude Comparison If the FWC detects a difference between the altitude indications, the message ” CHECK ALT ” appears on the PFD. In this case the pilots have to compare the indications with the Standby Altimeter to do the appropiate ADIRS switching.



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PFD

Figure 62

Attitude and Altitude Comparison Messages

______________________________________________________________________________________________________________________________________________________________________________________________



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MASTER WARNING AND MASTER CAUTION The Attention Getters ( Master Warning and Master Caution ) are triggered by the FWCs. The Attention Getters have two bulbs each. FWC 1 triggers the ” Master ”-line of the left Master Warning Light and the Master Caution Light and the ” Warning ” resp. ” Caution ”-line of the right Attention Getters. ( FWC 2 accordingly reversed ). When an Attention Getter is pushed ground is connected to both FWCs. The Attention Getters are connected to different FWC gates.

IS 84 rw 14.12.92 Page: Page: 114 ______________________________________________________________________________________________________________________________________________________________________________________________ Revision No : 02 Issue Date : 21/05/2013


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ASM 31-52-00

Figure 63

Master Warning and Master Caution Schematic

______________________________________________________________________________________________________________________________________________________________________________________________



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31-63

DISPLAY MANAGEMENT COMPUTER (DMC) - CATHOD RAY TUBE (CRT)

GENERAL Description Three convertible DMCs are installed in the A 320. Each DMC is able to drive one PFD, one ND and either ECAM display unit. In normal configuration, DMC 1 supplies captain’s PFD and ND and the E/WD with data. DMC 2 supplies FO’s PFD and ND and the SD. DMC 3 is active but does not supply any DU (Hot Spare). If the DMC Transfer Switch is turned to ”CAPT” DMC 3 supplies Captain’s DUs and the EW/D instead of DMC 1. The signals from DMC 3 are routed through DMC 1 then. ( In position ”FO” through DMC 2 to FO’s DUs and the SD.) The DMC receive data, process them and send them to the connected DUs.

Failures If a DMC fails the DUs connected show a white stroke. When DMC 3 fails in normal configuration the following message appears on the EW/D: in Amber EIS DMC 3 Fault

Inputs The DMCs receive ARINC 429 from some systems for EFIS indication ARINC 429 for System Display-parameters (mainly via from the SDACs) ARINC 429 from the EFIS control panels ARINC 429 from the ECAM control panel (system synoptic selection) ARINC 453 for weather radar information RS 422 from the FWC message bus for messages to be displayed on the lower part of the EW/D and Status Messages for the Status Page A discrete from the ECAM control panel ( ALL-pushbutton ) A discrete from the annunciator lights switch Discretes for EIS switching.



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Figure 64

DMC Interface

______________________________________________________________________________________________________________________________________________________________________________________________



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DMC / DU INTERCONNECTION DMC 1 and DMC 2 send data to be displayed to the connected DUs via a Dedicated Serial Data Link ( DSDL ). The pictures are generated in the DUs. The weather radar information is sent to the EFIS DUs via 4 extra buses, 3 for colour and one for synchro. DMC 1 sends PFD data to PFD 1. Additionally, the same bus is coonected to ND 1 for transfer purposes. ND data are sent to ND 1 and PFD 1. ECAM data are sent to the EW/D. After an EIS transfer, DMC 3 sends data through DMC 1 or DMC 2.

Schematic In the Schematic, the normal signal flow of DMC 1 is shown: Captain’s PFD uses the PFD Master DSDL and sends a PFD feedback. Captain’s ND uses the ND Master DSDL and the four weather radar Buses and sends a ND feedback. The Upper ECAM Display Unit uses ECAM Master DSDL (with Engine/ Warning Display information) and sends an ECAM feedback.

Feedback Each DU uses a return DSDL to send the data received from the DMC back to the DMC together with status and BITE information. The DMC compares this feedback with its own data input.

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DMC 1

DMC 3

DMC 2

PFD - Cpt

ND - Cpt

ECAM - upper

ECAM - lower

ASM 31-61-00

______________________________________________________________________________________________________________________________________________________________________________________________


Figure 65

DMC/DU Interconnection


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DMC OUTPUTS TO OTHER AIRCRAFT SYSTEMS DMC Output Buses Each DMC sends maintenance data to the CFDIU via an ARINC 429 low speed bus. These buses are accessable through test plugs in the avionic compartment for direct bus signal decoding by means of an appropriate portable bus reader, should the CFDIU be unserviceable. On a different bus (ARINC 429 high speed), DMC 1 and DMC 2 send digital data to the items listet below: FWCs The Flight Warning Computers receive data for comparison function. ECAM CP The ECAM Control Panel receives data for pushbutton illumination. PVI The Paravisual Indicator receives data for indication. FDIU The Flight Data Interface Unit receives data to be recorded in the Digital Flight Data Recorder. DMU The Data Management Unit receives data to be used in the AIDS.

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Figure 66

DMC Outputs

______________________________________________________________________________________________________________________________________________________________________________________________



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DMC Discrete Outputs Some computers receive discrete signals from the DMCs: AMU the AMU receives a signal from the EFIS Control Panels and the DMCs if the ILS pushbotton switch is pressed or if the Mode Selector Switch is in ILS position. This is for switching the DME audio signal from the VOR system to ILS. Captain’s discrete is from DMC 1. F/O’s discrete is from DMC 2. After an EIS DMC transfer the DMC 3 takes over from DMC 1 oder DMC 2. FWC the FWCs receive a valid discrete from each DMC. FMGC the FMGCs receive validity information from the PFDs. The FMGCs need this information to determine the landing category.



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AMM 23-51-00

AMU

Figure 67

ASM 23-51-00

DMC/AMU Discrete Interface

______________________________________________________________________________________________________________________________________________________________________________________________



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FWC INTERFACE Inputs The systems listet below send digital data to the FWCs: CFDIU (Centralized Fault Display Interface Unit ) ECAM Control Panel FMGC 1 & 2 (Flight Management and Guidance Computer) FAC 1 & 2 (Flight Augmentation Computer) FCU (Flight Control Unit) SDAC 1 & 2 (System Data Acquisition Concentrator) LGCIU 1 & 2 (Landing Gear Control and Interface Unit) FQI (Fuel Quantity Indication Computer) SDCU (Smoke Detevtion Control Unit) ECU Engine 1 & 2 (Electronic Control Unit) EIU 1 & 2 (Engine Interface Unit) Radio Altimeter 1 & 2 ILS 1 & 2 ADIRU 1, 2 & 3 (Air Data/Inertial Reference Unit) FCDC 1 & 2 (Flight Control Data Concentrator) DMC 1 & 2 (Display Management Computer) ACARS Management Unit Vaious aircraft systems send different kinds of analog inputs to the FWCs: Discretes Synchro signals The Discrete and Synchro Inputs are listet in the ASM 31-52-00.



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DMC 1

FWC 1 DMC 2

FWC 2

DMC 3

Figure 68

DMC/FWC Discrete Interface

Pa ______________________________________________________________________________________________________________________________________________________________________________________________ Revision No : 02 Issue Date : 21/05/2013


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DMC 1

FMGC 1

DMC 3

FMGC 1

DMC 2

______________________________________________________________________________________________________________________________________________________________________________________________ Figure 69 DMC/FMGC Discrete Interface



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ANNUNCIATOR LIGHT TEST General On the ground, when performing a cockpit light test (by means of the INT LT/ANN LT switch), all the DUs present fixed warning flag patterns, showing flags and annunciators at their proper location. This mode is inhibited in flight.

______________________________________________________________________________________________________________________________________________________________________________________________



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Figure 70

DU Indication during Light Test

______________________________________________________________________________________________________________________________________________________________________________________________



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DMC Discrete Input The test is activated by the DMCs on reception of a ground descrete signal.

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DMC 1

DMC 3

DMC 2

Figure 71

Wiring for Light Test Discrete

______________________________________________________________________________________________________________________________________________________________________________________________



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CRT General Six display units are installed. Theay normally present EFIS and ECAM information. The DUs receive all data from the DMCs. In turn the DUs send back to their driving DMC some feedback signals giving the DU status plus acknowledgement of data received from DMC.

Brightness Control Light sensors are installed on the face of each DU in order to provide automatic adjustment of the display brightness with changing light conditions. This automatic brightness adjustment is combined with the manual brightness control, which keeps always priority.

Overheat Protection If cooling air is lost, the grey background areas disappear from the PFDs and the WX image from the NDs. If the DU internal temperature exceeds a given threshold, the DU is automatically cut off.



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Figure 72

Display Unit

______________________________________________________________________________________________________________________________________________________________________________________________



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31-68

EIS SWITCHING

GENERAL Various reconfiguration possibilities are provided in the Electronic Instrument System (EIS) in order to cope with the operational requirements below in case of failure of a Display Management Computer (DMC), a Display Unit (DU) or a Control Panel (EFIS control section of the Flight Control Unit (FCU) or ECAM Control Panel) : DMC transfer: EIS DMC 1/3 or 2/3 (DMC 3 replacing DMC 1 or 2) PFD-ND transfer ECAM DU transfer: upper DU to lower DU ECAM-ND transfer: ECAM/CAPT ND or ECAM/F/0 ND.

EFIS Switching Each time the PFD/ND pushbutton is pressed, the images displayed on the PFD and ND are interchanged. The image previously displayed on the PFD is displayed on the ND and vice versa. The PFD potentiometer switches the PFD display unit on or off and controls the brightness in conjunction with the automatic brightness control system. In the off position, automatic and manual reconfigurations are possible. The ND inner potentiometer switches the ND Display Unit on or off and enables general adjustment of the ND brightness. The outer potentiometer only adjusts the brightness of the weather radar image.

Some of these transfers are performed automatically.



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Captain

F/O

Brightness Controls and PFD/ND Xfer Pushbuttons

EIS Transfer Selector Switches

Figure 73

EIS/ECAM Reconfigurations

______________________________________________________________________________________________________________________________________________________________________________________________



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Modul: EIS SWITCHING Modul: EIS TRANSFER ( nicht im Buch vorhanden ) Each DMC can supply only one ECAM DU!

PFD1

ND1

E/W

PFD/ND XFR

ND2

PFD2

PFD1

PFD/ND XFR

SYS

DMC 1

DMC 3

ND1

E/W

PFD/ND XFR

DMC 2

DMC 1

ECAM/ND XFR

EIS DMC

NORM

NORM

CAPT

F/O

CAPT 3

Normal Configuration

PFD/ND XFR

DMC 3

NORM F/O 3

PFD2

SYS

EIS DMC CAPT 3

ND2

DMC 2

ECAM/ND XFR NORM

F/O 3

CAPT

F/O

DMC Transfer Captain 3

DMC 1 supplies Captain’s EFIS and E/WD. DMC 2 supplies FO’s EFIS and SD.

Figure 74

DMC 3 replaces DMC 1.

EIS Switching part 1



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PFD1

ND1

PFD/ND XFR

E/W

ND2

PFD/ND XFR

SYS

DMC 1

DMC 3

EIS DMC NORM CAPT F/O 3 3

PFD1

PFD2

ND1

E/W

PFD/ND XFR

PFD/ND XFR

DMC 3

EIS DMC

ECAM/ND XFR NORM CAPT F/O

DMC 2

ECAM/ND XFR

NORM CAPT 3

DMC Transfer FO 3

PFD2

SYS

DMC 1

DMC 2

ND2

NORM F/O 3

CAPT

F/O

ECAM / ND Transfer FO

DMC 3 replaces DMC 2.

SD on FO’s ND. Lower ECAM DU empty.

Figure 75

EIS Switching part 2

______________________________________________________________________________________________________________________________________________________________________________________________



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OPERATION DU Inputs Each display unit comprises one NORMAL and one ALTERNATE input which can be selected by a discrete signal.

Note: Normally, display units use the normal input. If the discrete signal ( ground ) is given to a display unit, it uses the alternate input signal.

______________________________________________________________________________________________________________________________________________________________________________________________



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Display Unit

N

NORMAL INPUT

A

ALTERNATE INPUT

Figure 76

Discrete via relay box (selection of NORMAL or ALTERNATE input)

DU Normal / Alternate Discrete

______________________________________________________________________________________________________________________________________________________________________________________________



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Luftha

DU RECONFIGURATION

DMC TRANSFER

General A PFD/ND transfer can be done manually or automatically. For a transfer, the DUs switch to the other input signal.

Manual PFD / ND Transfer A manual PFD/ND transfer is done when the PFD/ND Transfer Pushbutton is pressed. The transfer is performed by switching on the alternate input of the DUs through a relay box. Two relay boxes are installed in the cockpit of the A 320 / A 321 behind the instrument panel. One relay box is for Captain’s PFD and ND, one relay box is for F/O’s PFD and ND. The relay boxes are supplied with 28 V DC.

Each DMC comprises three channels. Each channel is dedicated to either a PFD, a ND or an ECAM display. Each DMC features switching capabilities for display transfer or DMC transfer. The DMC transfer is performed by a switching inside each DMC. All the signals then come from DMC 3. The dicrete for DMC transfer is given from the EIS DMC Selector Switch. It also activates the relay box.

Automatic PFD / ND Transfer In the case of detected failure of the DU normally displaying the PFD image, there is an automatic PFD / ND transfer. The PFD image is presented on the remaining EFIS DU. This automatic transfer is controlled by the DMC via a discrete to the relay box. If the PFD DU knob is turned to OFF, the relay box is activated and the PFD image is automatically displayed on the other DU. ( PFD priority over ND )

Automatic Transfer from Upper ECAM DU to Lower ECAM DU In the event of upper ECAM DU failure, the Engine/Warning image is displayed on the lower ECAM DU instead of the system page or status page. This switching is automatic: on reception of the upper DU anomaly signal, through the feedback DSDL bus, the DMC 2 ECAM channel processor switches by software to an ECAM single display configuration which privileges the E/W processing. The lower ECAM DU receives the E/W image from the DMC 2, still through its NORMAL input. The same applies when turning the UPPER DISPLAY potentiometer to OFF.



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from DMC 2

Captain EIS DMC Selector Switch to DMC 2

DMC 1

DMC 3

Figure 77

AMM 31-68-00

EFIS Transfers (Captain)

______________________________________________________________________________________________________________________________________________________________________________________________



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ECAM / ND TRANSFER

L

General The transfer between ECAM an ND is performed inside the DMCs.

Position CAPT With the ECAM/ND transfer switch in this position, captain’s ND operates like a lower ECAM DU. DMC 1 can process only one ECAM picture, so the upper ECAM DU will receive no input and display a diagonal stroke. The lower ECAM DU operates as a E/WD now. It still receives its data from DMC 2.

Position F/O With the ECAM/ND transfer switch in this position, FO’s ND operates like a lower ECAM DU. DMC 2 can process only one ECAM picture, so the lower ECAM DU will receive no input and display a diagonal stroke. The upper ECAM DU still operates as E/WD and it still reiceives data from DMC 1.



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ECAM/ND Transfer Selector Switch

PFD

ND

ECAM

DMCs

Figure 78

ECAM / ND Transfer

______________________________________________________________________________________________________________________________________________________________________________________________



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GENERAL SWITCHING The picture below shows the connection of the DUs to the DMCs. Note that each Display Unit has a normal input (N) and an alternate input (A).

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DMC 1

DMC 2

DMC 3

Figure 79

AMM 31-63-00

EIS Configuration

______________________________________________________________________________________________________________________________________________________________________________________________



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31-60

ELECTRONIC INSTRUMENT SYSTEM

POWER SUPPLY Description The table shows the power supply of the EFIS and ECAM LRUs.

Normal Operation The 3 DMCs and the 6 DUs are supplied with 115V/400Hz and with 28V DC. The DC power is needed for the switching.

Emergency Operation If the Emergency Buses are supplied onlyEmergency Operation means that the If only the AC Essential and the DC Essential Buses are supplied with electrical power the following LRUs remain available: PFD 1 ND 1 E/WD ECAM CTL PNL FCU SDAC 1 FWC 1 MASTER CAUTION LIGHTS (half) MASTER WARNING LIGHTS (half) DMC 1 DMC 3 (only when EIS DMC CAPT 3 is selected). DMC Power Supply Switching When DMC 1 fails the Captain selects DMC 3. In the case of AC Bus 1 failure, the relay 26WT switches the power supply to the AC Essential Bus.

______________________________________________________________________________________________________________________________________________________________________________________________



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AMM 31-60-00

Figure 80

AMM: EIS Power Supply

______________________________________________________________________________________________________________________________________________________________________________________________



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31-69

EIS - TEST / BITE

GENERAL The computers of the Electronic Instrument Systems include a BITE ( Built-In Test Equipment ). With this BITE the LRUs of the EIS are monitored and tested. The components of the Electronic Instruments System are: 6 Display Units ( DU ) 3 Display Management Computers ( DMC ) The test of each DMC and its connected DUs is started via the MCDU. The FAC 1 acts as FIDS ( Fault Isolation and Detection System ) for the FCU which includes the EFIS control sections.



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Feedback from DUs

ECAM CP

Feedback from DUs DMC 3

DMC 1

DMC 2

SDAC 1 FWC 2

FWC 1 SDAC 2

CFDIU

Figure 81

EIS Bite Data Flow

______________________________________________________________________________________________________________________________________________________________________________________________



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EIS TEST Via SYSTEM REPORT / TEST one of the three EIS can be selected. In the menu mode various menus are offered: Last Leg Report The Last Leg Report indicates the failures which occured during the last leg. Previous Legs Report The Previous Legs Report indicates the failures which occured during the previous legs. LRU Identification indicates the partnumbers of the DMC selected and the connected DUs. If DMC 3 is selected without DMC switching, no partnumbers of DUs will be indicated. Engines The DMC connected to the upper ECAM DU monitors primary parameter indications of both engines. Should an exceedance occur, the DMC memorizes the maximum value reached during the last flight leg. The ENGINES function is used to indicate the exceedances.

Test The Test-Menu offers three different tests: SYSTEM TEST DISPLAY TEST INPUT TEST: When selected, all inputs are monitored, incuding sources which are not used in the normal configuration ( e.g. FWC 2 ).

Dump BITE Memory This functions offers the possibility to display memorized data.



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Last Leg Report

Previous Legs Report

LRU Identification ( active DMC )

Engines Overspeed / Overtemperature

Figure 82

Dump BITE Memory

DMC Test

EIS Menu

______________________________________________________________________________________________________________________________________________________________________________________________



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EIS System Test The EIS System Test can be found in the AMM 31-60-00. When selected, the DMC performs the functional tests of all the internal functions of the system (DMC, DUs, DMC-DUs links). The procedure to do the EIS System Test is: 1. Do the Job Set-up according to AMM. 2. For DMC 1 and 2, select EIS DMC Switch to NORM. For DMC 3, select CAPT 3. 3. Select SYSTEM/REPORT TEST menu. 4. Push LSK 4R for INST The Instrument page appears. 5. Push the LSK adjacent to EIS 1, EIS 2 or EIS 3. The EIS DMC page appears. On the DUs which are connected to the DMC, the image disappears and MAINTENANCE MODE comes into view. 6. Push LSK 4R for TEST. The EIS DMC Test page appears. 7. Push LSK 2L for SYSTEM TEST. The first EIS DMC System Test Page appears. 8. Push LSK 6L for RETURN. The second EIS DMC System Test page appears. 9. Push LSK 5L for START TEST. The MCDU displays NO RESPONSE-PRESS RETURN. 10.Push LSK 6L for RETURN. The Instrument page appears. The DUs go back to the normal mode. 11. Push the same LSK like in step 5 to select the same EIS. The EIS DMC page appears. The DUs operate in the maintenance mode. 12.Push LSK 4R for TEST. The EIS DMC Test page appears. 13.Push LSK 5L for SYSTEM TEST RESULT. The TEST OK indication comes into view when there is no failure detected. 14.Push LSK 6L for RETURN. 15.Do the Close-up according to AMM.



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Figure 83

System Test

______________________________________________________________________________________________________________________________________________________________________________________________



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Display Test When selected, a test pattern is displayed on the DUs. This test pattern enables the maintenance crew to assesess the condition of all the DUs for ageing and display quality, thus providing at a glance a confidence check of the Display Units. The procedure to do the Display Test is: 1. Select SYSTEM/REPORT TEST menu. 2. Push LSK 4R for INST The Instrument page appears. 3. Push the LSK adjacent to EIS 1, EIS 2 or EIS 3. The EIS DMC page appears. On the DUs which are connected to the DMC, the image disappears and MAINTENANCE MODE comes into view. 4. Push LSK 4R for TEST. The EIS DMC Test page appears. 5. Push LSK 3L for DISPLAY TEST. On the DUs which are connected to the DMC the test pattern appears.

______________________________________________________________________________________________________________________________________________________________________________________________



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Figure 84

Display Test

______________________________________________________________________________________________________________________________________________________________________________________________



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Input Test When selected, all inputs are monitored, including sources which are not used in the normal configuration ( e.g. FWC 2 ). During 5 s, the DMC scans all its ONSIDE buses, i.e. it checks that each bus sends labels with their correct refresh rate. During the next 5 s, the DMC scans all its OFFSIDE buses. At the end of this test, the DMC signals all the buses seen as faulty for display on the I/PTEST RESULTS display. The procedure to do the Input Test is: 1. Select SYSTEM/REPORT TEST menu. 2. Push LSK 4R for INST The Instrument page appears. 3. Push the LSK adjacent to EIS 1, EIS 2 or EIS 3. The EIS DMC page appears. On the DUs which are connected to the DMC, the image disappears and MAINTENANCE MODE comes into view. 4. Push LSK 4R for TEST. The EIS DMC Test page appears. 5. Push LSK 4L for I/P TEST.



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Figure 85

Input Test

______________________________________________________________________________________________________________________________________________________________________________________________



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DUMP OF THE BITE MEMORY The DUMP BITE MEMORY function displays a menu which presents the following failure items: INTERNAL FAULT IN FLIGHT EXTERNAL FAULT IN FLIGHT FAULT ON GROUND The FAILURE COUNTER RESET item may be used to reset the failure counters. Note: The Information is coded in hexadecimal.



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Figure 86

Dump of the BITE Memory

______________________________________________________________________________________________________________________________________________________________________________________________



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31-21

ELECTRICAL CLOCK

Modul: ELECTRIC CLOCK DESCRIPTION GENERAL UTC Centre Display The Centre Display (UTC) indicates the current time (hours, minutes). Periods of 15 seconds are indicated by three horizontal segments. On request, the current date (month, day) is displayed when the SET knob is pressed.

Elapsed Time Display and Elapsed Time Selector A Bottom Display, called the Elapsed Time (ET) indicates: The elapsed time, provided the ET Selector is set to RUN. The ET Display is frozen if the ET Selector is set from RUN back to STOP. Or the year when the UTC Selector is set to MO. The counter is reset to zero and the display goes off when the ET Selector is set to RST (Reset) position.

Chronometer Display and CHR Pushbutton An Upper Display, called CHR (Chronometer) indicates the minutes provided the chronometer pushbutton has been pressed. The seconds are indicated by a sweep hand. Pressing the CHR pushbutton again will stop the chronometer function. To reset the chronometer to zero, the CHR pushbutton has to be pressed a third time.

UTC Selector An UTC selector allows Date or Time updating. MO : to set months and years DY : to set day HR : to set hours MIN : to set minutes RUN : to start the UTC counter The UTC selector must be pressed an turned to set it from RUN to MIN position.

Set Knob A SET knob allows the current date to be displayed if the UTC selector is set to RUN. The SET knob is also used to update time and date according to the UTC selector position. To increment: turn to either side To decrement: push lightly.

Test The Clock is tested from the ANN LT selector. All displays must show 8.



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SET KNOB MO DY

DATE

CHRONOMETER DISPLAY

UTC SELECTOR

HR MIN UT C RUN

SET

50

10

CHR HR

MIN

UTC CENTRE DISPLAY

ELAPSED TIME DISPLAY MO

40

ELAPSED TIME SELECTOR

UTC HR

MIN

DY

20 CHRONOMETER PUSHBUTTON

ET RUN S T

E O

P

T

CHR RST

Figure 87

Electrical Clock

______________________________________________________________________________________________________________________________________________________________________________________________



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POWER SUPPLY

CLOCK INTERFACE

The clock is important for the function of the CFDS. Its electrical connection guarantees at least the supply of the time base. ESS BUS 401PP normaly supplies time base and indication. HOT BUS 701PP supplies the time base when ESS BUS 401 PP is lost. The indication, however, is lost. The Internal Battery supplies the time base for maximal 15 days when ESS BUS and HOT BUS are lost. After main battery change it is not necessary to readjust the clock. By means of a shunt in the plug it is guaranteed that the clock stops working after removal. When the plug is removed the internal power supply via the internal battery is interrupted.

The clock transmitts UTC ( Universal Time Coordinated ) in ARINC 429 format to following computers: Centralized Fault Display Interface Unit ( CFDIU ) Flight Data Interface Unit ( FDIU ) Flight Management and Guidance Computers ( FMGCs )



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Figure 88

Clock - Principle Diagram

______________________________________________________________________________________________________________________________________________________________________________________________



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CFDIU Interface with the Clock The CFDIU is connected to the clock, located on the center instrument panel, which provides: GMT/Date elapsed time chronometer. The clock provides the date and the time at which faults occur. The time associated with the fault messages and the ECAM warnings enables the correlation between the ECAM warnings and the fault messages memorized by the CFDIU. In the event of loss of clock operation or incorrect operation detected by the CFDIU, the CFDIU takes over and calculates the time and the date using its internal clock. This transition does not require reinitialization if there is not a long power cutoff (> 200 ms). If there is a long power cutoff, the crew performs reinitialization using the MCDU. The CLOCK is declared failed or invalid (= incorrect operation) by the CFDIU when one of the labels of the GMT and DATE parameters is not refreshed over 1 minute.

Clock Indication on ECAM The CFDIU permanently acquires GMT and the date and transmits these data on its output buses, for examle to the DMCs. During normal operation, DMC 2 sends data to the lower ECAM DU.



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Clock

DMC CFDIU

MM 31-32-00

Figure 89

Clock-ECAM Interface

______________________________________________________________________________________________________________________________________________________________________________________________



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OPERATIONAL TEST The AMM offers an Operational Test in ATA Chapter 31-21-00. During this Test, the Clock Switches and the Annunciator Lights Switch must be used and the Clock Indications must be monitored.



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Clock

Annunciator Lights Switch

AMM 31-21-00

Figure 90

Clock Component Location

______________________________________________________________________________________________________________________________________________________________________________________________



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Training Manual A 319/320/321 ATA 31 Indicating/Recording 31-32 Centralized Fault Display System

Line & Base Maintenance


Lufthansa Issue: JAN. 1998 Technical Training GmbH For Training Purposes Only Book No: All A319/320/321 31-32 LEVEL 3 E Lufthansa Base Lufthansa 1995 ______________________________________________________________________________________________________________________________________________________________________________________________



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ATA 31

INDICATING AND RECORDING

31-32

CENTRALIZED FAULT DISPLAY SYSTEM (CFDS)

______________________________________________________________________________________________________________________________________________________________________________________________



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PRINTER The PRINTER is used for printing information automatically or on request.

CFDS PRESENTATION CFDIU The Centralized Fault Display Interface Unit (CFDIU) receives failure messages from the aircraft systems. It memorizes and manages them. Information is available in various reports. The CFDIU consists of two distinct channels: a NORMAL CHANNEL which ensures all the functions. a STANDBY CHANNEL (or BACKUP CHANNEL) which permits restricted operation when the normal channel is faulty.

ACARS The ACARS (Aircraft Communication Addressing and Reporting System) is used to exchange data between the aircraft and a ground station via a radio VHF link.

BITE The BITE is a function incorporated in the computers which detects, localizes and memorizes failures. All systems including a Built in Test Equipment (BITE) are connected to the CFDIU. ECAM The ECAM monitors the aircraft systems. The warning information is delivered to the Centralized Fault Display System. FWC: Flight Warning Computer Only the primary and the independent failure information is sent to the CFDS. MCDU The Multipurpose Control and Display Unit (MCDU) is a display unit and a keyboard used by the CFDS to display and interrogate BITE‘s and to initiate system tests. The 2nd MCDU is also connected to the CFDS. You can only use the CFDS on one MCDU at a time.

FRA TS84, wd 7.12.93 Seite: Page: 2 ______________________________________________________________________________________________________________________________________________________________________________________________ Revision No : 02 Issue Date : 21/05/2013


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AIRCRAFT SYSTEM COMPUTERS

ECAM ACARS D.U.

CFDIU

FWC 1 2

MCDU 1

MCDU 2

PRINTER Figure 1

CFDS Presentation

______________________________________________________________________________________________________________________________________________________________________________________________



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SYSTEM BITE PHILOSOPHY BITE Most aircraft systems are equipped with a Built-In Test Equipment (BITE). The BITE monitors permanently the system operation. It can also store and transmit the detected failure. Each system computer includes a BITE circuit which detects failures. When a failure is detected, it is stored in the BITE memory and is transmitted to the centralized fault display system. Memorization of the 64 previous legs report is done by most of the BITE‘s. CFDS The Centralized Fault Display System centralizes all information concerning aircraft system failures. Reading or printing of all the failure information is done in the cockpit. The CFDS functions are accessed through the MCDU.



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AIRCRAFT SYSTEM COMPUTER ECAM BITE

memory

ECAM WARNING MESSAGES

BITE INFORMATION

CFDIU Centralized Fault Display Interface Unit

MCDU =====-

CFDS

=-=-=-=-==-=-=-=-=-==-==-=..... ..... ..... ..... ..... .....

=====-

Multipurpose Control and Display Unit Figure 2

PRINTER System BITE Philosophy

______________________________________________________________________________________________________________________________________________________________________________________________



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SYSTEM BITE PHILOSOPHY (CONT.) LAST LEG REPORT A CURRENT LEG REPORT is elaborated during the flight. After the flight, its title becomes LAST LEG REPORT. All failures are reported in the same form and also indicate ATA reference and time of failure occurence. Example of failure message: GMT ATA 0920 28-21-00 FUEL L TK PUMP 1QM Time ATA Failure Message FIN (Functional Reference Item Number) LAST LEG ECAM REPORT The primary failure and the independent failure messages are delivered by the ECAM to the CFDIU. Warning messages coming from the ECAM are stored in the CURRENT LEG ECAM REPORT during the flight. After the flight, the title of this report becomes LAST LEG ECAM REPORT. Example of warning message: GMT: 1125 PH: 06 ATA: 3155 HYD BLUE RSVR OVHT POST FLIGHT REPORT (PFR) The POST FLIGHT REPORT is the sum of the LAST LEG REPORT and of the LAST LEG ECAM REPORT. The POST FLIGHT REPORT can only be printed on ground. The list of ECAM WARNINGS and FAULT MESSAGES with the associated time and ATA reference allow the maintenance crew to make a correlation for easier trouble-shooting.



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LAST LEG REPORT

A/C IDENT F-WWAI

DATE FEB23

GMT 2350

FLTN 280

LAST LEG ECAM REPORT

CITY PAIR LFBO/LFPO

MAINTENANCE POST FLIGHT REPORT ECAM WARNINGS GMT PH ATA 1125 06 31-55 HYD BLUE RSVR OVHT 1100 06 27-00 SFCC 1 FAULT 0920 05 28-21 FUEL L TK PUMP 1 LO PR 0904 04 36-22 BLEED LOOP 0854 04 22-00 LAND 3 INOP

FAULT MESSAGES GMT ATA 1125 31-55-00 HYD BLUE TEMP SENSOR 1105 26-17-00 SDCU CHANNEL 1 10WQ 1100 27-00-00 NO SFCC 1 DATA 0920 28-21-00 FUEL L TK PUMP (1QM) 0915 26-12-00 CHECK EIU 1 0904 36-22-00 CHECK R WING LOOP A 0854 22-00-00 FMGC 1

Standard for Example

DLH Version Figure 3

Post Flight Report

______________________________________________________________________________________________________________________________________________________________________________________________



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SYSTEM BITE PHILOSOPHY CONT.) Internal/External Failures Each Built-In Test Equipment (BITE) can make the difference between an internal and an external failure. Let us suppose that an angle of attack sensor failure has been detected and that systems A, B and C are affected by this failure. The AIR DATA system will transmit an INTERNAL FAILURE (= SOURCE on the POST FLIGHT REPORT). Systems A, B and C will transmit EXTERNAL FAILURE (= IDENTIFIERS on the POST FLIGHT REPORT).



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AIR DATA SYSTEM

C ANGLE OF ATTACK SENSOR

B AIR DATA COMPUTER

A

INTERNAL EXTERNAL FAILURE

FAILURE

CFDIU

Figure 4

Internal/External Failures

______________________________________________________________________________________________________________________________________________________________________________________________



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SYSTEM BITE PHILOSOPHY (CONT.)

CFDS MODES OF OPERATION

Memorization Memorization of failures is different when the aircraft is on ground or in flight. The full BITE functions and memorization operate in flight. On ground, the memorization is done only in the BITE‘s. The BITE‘s are provided with flight and ground memory zones.

Two CFDS modes are available. NORMAL MODE is always active except on ground when MENU MODE is selected. NORMAL MODE In this mode, the CFDIU scans all the connected system outputs and memorizes the failure messages in order to generate the current (last) leg report and the current (last) leg ECAM report. In flight the CFDS always operates in normal mode. MENU MODE In this mode, the CFDIU dialogues with one computer at a time in order to read the contents of its BITE memory and initiate various tests (SYSTEM REPORT/TEST). This mode can only be selected on ground and interrupts the normal mode of operation.

Failure Gravity The failures are classified according to their importance. Class 1 failures are the most serious ones and require an immediate maintenance action subject to the minimum equipment list. Class 2 failures may have consequences if a second failure occurs. A maintenance action is necessary at the next adequate opportunity. Class 3 failures can be left uncorrected until the next scheduled maintenance check.



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MEMORIZATION

FAILURE GRAVITY

CLASS 1 BITE GROUND MEMORY

FLIGHT MEMORY

CLASS 2

OPERATIONAL CONSEQUENCE

CLASS 3 NO IMMEDIATE OPERATIONAL CONSEQUENCE

NO CONSEQUENCE ON AIRCRAFT SAFETY

Figure 5

Memorization/Failure Gravity

______________________________________________________________________________________________________________________________________________________________________________________________



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SYSTEM TYPES Type 1 Systems Most Systems are type 1 systems. These systems can memorize failures occured in the last 64 flight legs. Type 1 systems are connected to the Centralized Fault Display Interface Unit (CFDIU) via an ARINC 429 input bus and an ARINC 429 output bus. SINGLE COMPUTER The first configuration in TYPE 1 is a single computer. Example: VHF 1 Transceiver MULTI COMPUTER The second configuration in TYPE 1 includes several computers in the same aircraft system. One computer concentrates the maintenance data of the other computers. Example: Flight Management and Guidance Computers (FMGC) and Flight Augmentation Computer (FAC) with FMGC1 as A FMGC2 as B FAC as C DUPLICATED SYSTEM A duplicated system includes two different subsystems in the same computer. Example: Air Data and Inertial Reference Unit (ADIRU) with ADR as subsystem 1 IR as subsystem 2

Type 3 Systems Type 3 Systems are simple systems linked to the CFDS by only two discrete signals. Type 3 systems cannot memorize failure messages. The discrete input permits to initiate the test or reset. The discrete output indicates if the system is OK or not. Example: Transformer Rectifier Unit (TRU)

Type 2 Systems Type 2 Systems memorize only failures from the last flight leg. The discrete signal is provided to initiate the test of the system. Example: Avionic Electronic Ventilation Computer (AEVC)



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TYPE 1 SYSTEMS B I T E

SINGLE COMPUTER

A

TYPE 2 SYSTEM

B I T E

MULTI COMPUTER SYSTEM

C B

B I T E

C F

S Y S T B E I M T E

D I

B I T E

U

TYPE 3 SYSTEM B I T E

SUB SYSTEM

DUPLICATED SYSTEM

1 SUB SYSTEM 2

B I T E B I T E

Figure 6

System Types

______________________________________________________________________________________________________________________________________________________________________________________________



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FAILURES CLASSIFICATION Class 1 Failure Class 1 failures have an operational consequence on the flight. You can display the class 1 failures on the Multipurpose Control and Display Unit (MCDU): In the LAST (or CURRENT) LEG REPORT In the LAST (or CURRENT) LEG ECAM REPORT. Refer to the Minimum Equipment List (MMEL): ”GO, ”GO IF” or ”NO GO”.



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ELEC

MASTER

PUMP

FAULT

CAUT OFF

HYD

B RSVR

A U T O

OVHT

CURRENT LEG REPORT date:feb 23

-BLUE ELEC PUMP..OFF

STS

GMT:1125 ATA:31-55-00 HYD BLUE TEMP SENSOR

3000 PSI

3000

PSI

3000

< RETURN

PRINT *

OVHT

Figure 7

Class 1 Failure

______________________________________________________________________________________________________________________________________________________________________________________________



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FAILURES CLASSIFICATION (CONT.) Class 2 Failure Class 2 failures have no immediate operational consequence and can be displayed on request on the ECAM STATUS page. You can display the class 2 failures on the Multipurpose Control and Display Unit (MCDU): In the LAST (or CURRENT) LEG REPORT In the LAST (or CURRENT) LEG ECAM REPORT. Refer to the MMEL: ”GO” without condition. Example: Single smoke detector fault in Smoke Detection Control Unit (SDCU).



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CURRENT LEG REPORT date:feb 23 GMT:1105 ATA:26-17-00 SDCU CHANNEL 1(10WQ)

STATUS

< RETURN

MAINTENANCE

PRINT *

SDCU

Figure 8

Class 2 Failure

______________________________________________________________________________________________________________________________________________________________________________________________



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FAILURES CLASSIFICATION (CONT.) Class 3 Failure Class 3 failures have no operational consequence. All aircraft systems remain available. You can display the name of the systems affected by at least a class 3 failure in the AVIONICS STATUS. The Class 3 failures can be left uncorrected until the next scheduled maintenance check. (At least before 400 hours or a A check). Do not refer to the MMEL.



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AVIONICS STATUS

DMC 3 (CLASS 3) STATUS NORMAL

< RETURN

Figure 9

PRINT *

Class 3 Failure

______________________________________________________________________________________________________________________________________________________________________________________________



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FAILURES CLASSIFICATION (CONT.) Synthesis NOTE: AVIONICS STATUS displays on ground the title of the systems currently affected by any failure class. Class 1 and 2 failures are displayed in the LAST LEG REPORT and in the LAST LEG ECAM REPORT.



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CLASS 1 FAILURE Operational consequence on the current flight Indicated to the pilots

Dispatch consequences

Maintenance information

CLASS 2 FAILURE

CLASS 3 FAILURE

NO

NO

YES YES Warnings/ flags System pages

YES On the system Display ”STATUS” page.

REFER TO MMEL FUNCTIONS LOST may be: INDICATED IN MMEL ”GO” ”GO IF” ”GO” without ”NO GO” condition Have to be reported by the pilots in the log book. Are indicated at the end of each flight leg. MMEL entry is required.

Figure 10

NO

NO REFERENCE IN MMEL Available on request. Can be left uncorrected until the next scheduled maintenance check.

Failure Classification Synthesis

______________________________________________________________________________________________________________________________________________________________________________________________



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MCDU DESCRIPTION The brightness knob enables the brightness of the display to be adjusted. By pressing the MCDU menu key, the MCDU menu page is displayed, and any one of the systems connected to the MCDU can be selected. The green colour indicates the system already in dialogue with the MCDU. The other systems are displayed in white. A multiple page display is indicated by an arrow in the right upper corner of the screen. In this case the NEXT PAGE key must be used to give access to the various pages of the display. The NEXT PAGE key can be used as long as the arrow is displayed. Some displays contain too many data for a single page. In this instances, vertical scroll keys can be used to scroll display, up or down. The scroll keys can be used as long as these arrows are displayed. Twelve line select keys, six on the left and six on the right, give access to a page or a function. The line select keys permit access to a page or a function when these symbols appear (>, <, *). They are identified as 1L to 6L on the left, and 1R to 6R on the right.

The MCDU menu annunciator illuminates white when a system connected to the MCDU request the display.

The CRT contains 14 lines, each having 24 characters. The top line is used as title line and the bottom one, is the scratchpad. Two character size can be used, as well as various colours: white, cyan, green, amber. Various symbols <, >, *, , , can be displayed to indicate special functions.



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Figure 11

MCDU

______________________________________________________________________________________________________________________________________________________________________________________________



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CFDS REPORTS General On ground, all the functions are available. In flight, only CURRENT LEG REPORT and CURRENT LEG ECAM REPORT are available. Note that the CFDS menu comprises two pages on ground.

______________________________________________________________________________________________________________________________________________________________________________________________



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< PFR FILTER PROGRAM

Figure 12

CFDS Menu on Ground

______________________________________________________________________________________________________________________________________________________________________________________________



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CFDS REPORTS (CONT.) LAST LEG REPORT The LAST LEG REPORT displays failure information delivered by the BITE‘s of the aircraft systems. It can store up to 20 failures occured during the last leg. Pressing the SFCC channel 1 (4L) line key allows access to the corresponding IDENTIFIERS page. The Last Leg Report displays the internal failures (class 1 and 2) only. On the ground, the title of this item is ”LAST LEG REPORT”. In the flight, it is ”CURRENT LEG REPORT”. When the report is displayed on several pages, an arrow appears on the top right-hand corner. The NEXT PAGE key permits to see the following pages. If you select the NEXT PAGE key on the last page, you come back to the first page. When you select the PRINT line key, all the LAST LEG REPORT is printed, even if it contains several pages.



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Figure 13

Last Leg Report

______________________________________________________________________________________________________________________________________________________________________________________________



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CFDS REPORTS (CONT.) LAST LEG ECAM REPORT The LAST LEG ECAM REPORT displays the list of ECAM warning messages sent to the CFDIU by the flight warning computers. It can store up to 20 warnings occured during the last leg. On ground, the title of this item is ”LAST LEG ECAM REPORT”, in flight it is ”CURRENT LEG ECAM REPORT”. DOCUMENTARY DATA appears on the print report: the A/C identification the city pair the flight number date and GMT (UTC). All the report is printed.



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Figure 14

Last Leg ECAM Report

______________________________________________________________________________________________________________________________________________________________________________________________



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CFDS REPORTS (CONT.) PREVIOUS LEGS REPORT At each new flight leg, the content of the LAST LEG REPORT is transferred into the PREVIOUS LEGS REPORT. This report can store up to 200 failures over the last 63 flight legs. Each failure message contains the same data as the LAST LEG REPORT: i.e.: NO FAC 1 DATA FEB 22 13 12 22-00-00 It also contains a flight leg counter relative to the previous flight. -XX is the number of flight legs before the last flight leg: i.e.: -01 (previous leg). The PREVIOUS LEGS REPORT is displayed only on ground. (INTM) means that the failure has occurred intermittently. When you make a print of the PREVIOUS LEGS REPORT, only the displayed page is printed.



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Figure 15

Previous Legs Report

______________________________________________________________________________________________________________________________________________________________________________________________



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CFDS REPORTS (CONT.) AVIONICS STATUS The AVIONICS STATUS presents the list of systems which are currently affected by a failure. This function is only available on ground. The information presented is permanently updated. The message contains the name of the systems presently affected by a failure, i.e. VHF 3, or a NO X DATA message when the related system X bus is not active, i.e. NO ILS 2 DATA. The AVIONICS STATUS also indicates the class 3 failures. (Class 3 ) means that the system is affected by at least one class 3 failure. Note that there could also be class 1 or 2 failures. When you make a print, all the AVIONICS STATUS report is printed even if it contains several pages.



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Figure 16

Avionics Status

______________________________________________________________________________________________________________________________________________________________________________________________



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CFDS REPORTS (CONT.) SYSTEM REPORT/TEST General The SYSTEM REPORT/TEST function is available on the ground only. It enables a dialogue between the CFDS and one system computer. The SYSTEM REPORT/TEST menu presents the list of all the systems connected to the Centralized Fault Display Interface Unit, in ATA chapter order. An example for each system type is available: in INST for type 1 systems in AIR COND for type 2 systems in ELEC for type 3 systems ECAM 1 (Type 1 system) Type 1 systems are the most common systems. The menu they present depends on the system itself. Now, you are in MENU mode. The menu is transmitted by the system itself. You talk directly with the system. The menu includes three basic functions: the LAST LEG REPORT the PREVIOUS LEGS REPORT the LRU IDENTIFICATION and optional functions, depending on the system, for example here TROUBLE SHOOTING DATA CLASS 3 FAULTS TEST GROUND SCANNING

AEVC, Avionics Equipment Ventilation Computer (Type 2 system) Type 2 systems present a menu with one basic function, the LAST LEG REPORT and optional functions depending on the system. You are in PSEUDO-MENU mode. The menu is transmitted by the CFDIU. You don‘t talk directly to the system. The system permanently transmits its data on the system bus, and the CFDIU reads them, except for the test. The menu includes one basic function: the LAST LEG REPORT and optional functions depending on the system, for example here TEST CLASS 3 FAULTS GCU EMER, Generator Control Unit Emergency (Type 3 system) Type 3 systems present only one function on their menu. Type 3 systems have no MENU mode. The available functions are displayed by the CFDIU. The only possible functions are TEST or RESET. When you make a test or a reset, the CFDIU initiates the test or reset, and reads the result on the CFDIU discrete for MCDU display..



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SYSTEM REPORT/TEST MENU

Figure 17

System Report/ Test

______________________________________________________________________________________________________________________________________________________________________________________________



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CFDS REPORTS (CONT.) POST FLIGHT REPORT (PFR) The POST FLIGHT REPORT is the sum of the LAST LEG REPORT and of the LAST LEG ECAM REPORT. It is only available on the printer. ECAM WARNINGS (or ECAM WARNING MESSAGES) display the LAST LEG ECAM REPORT. FAULT MESSAGES (or FAILURE MESSAGES) display the LAST LEG REPORT. You can send this report to the ACARS.

______________________________________________________________________________________________________________________________________________________________________________________________



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Standard for Example

DLH Version Figure 18

Post Flight Report

______________________________________________________________________________________________________________________________________________________________________________________________



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GMT/DATE INITIALIZATION The GMT/DATE INITialization function is available only in case of clock failure plus CFDIU power interrupt. The Centralized Fault Display System permits to reinitialize the time and the date on the multipupose control and display unit. GMT (UTC) and date are entered using the scratchpad.



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TO ALL OTHER A/C SYSTEMS

Figure 19

GMT/ Date Initialization

______________________________________________________________________________________________________________________________________________________________________________________________



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ACARS/ PRINT PROGRAM An ACARS/PRINT PROGRAM function is available. It enables programming of the CFDS reports transmission to the ACARS and to the printer The functions written in green are delivered by the ACARS or the CFDIU. They cannot be modified by the flight crew. The functions written in blue can be changed manually. If you select one of these functions, you will switch the YES message to NO and vice versa. The REAL TIME FAILURES provide, in real time, all the internal failure messages delivered by the systems or created by the CFDIU (CURRENT LEG REPORT DATA). When the associated SEND is on YES, this data is automatically transmitted to the ACARS, in real time. The REAL TIME WARNINGS function provides, in real time, warning messages, sent by the Flight Warning Computers (CURRENT LEG ECAM REPORT DATA). When the associated SEND is on YES, the report is automatically transmitted in real time, to the ACARS. The POST FLIGHT REPORT is the sum of the LAST LEG REPORT and of the LAST LEG ECAM REPORT. When the associated SEND is on YES, the POST FLIGHT REPORT will be automatically transmitted to the ACARS at the end of the flight (transition from flight phase 9 to 10). The PRINT function associated to the POST FLIGHT REPORT the REAL TIME FAILURES the REAL TIME WARNINGS permits an automatic print of the report. The POST FLIGHT REPORT will be printed automatically at transition from flight phase 9 to 10 (Second engine shutdown). Upon power on, the last selected programmed functions are still present. At the initialization, the manual programming functions present the last configuration in vigor.

The AVIONICS DATA function enables to send and/or print system pages available in the SYSTEM REPORT/TEST item. The printing or/and sending is not automatic; you must select the print line key displayed in the system page. NOTE: In the system pages, the PRINT message cannot be modified. But when you print and send the system pages, the ”PRINT ALSO SEND” message appears in the scratchpad. Functions delivered by the ACARS: When the CFDIU has not received any programming from the ACARS, the YES or NO message is replaced by a blank.



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10

Figure 20

ACARS/Print Programm

______________________________________________________________________________________________________________________________________________________________________________________________



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BACKUP MODE The BACKUP MODE function is only displayed in case of CFDIU main channel failure. It enables access to the functions of the backup channel. In flight, no function is available. On ground, the only function possible is SYSTEM REPORT/ TEST. This function is available for the main systems only.



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Figure 21

Backup Mode

______________________________________________________________________________________________________________________________________________________________________________________________



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CFDS FLIGHT PHASES A- In flight and before touchdown In flight, the full BITE functions are operative. Transmission and storage of internal and external failures for type 1/2/3 systems in their flight memory. B- Between touchdown and 80 kts + 30 seconds This phase differs from the previous one because type 2 systems are now ”on ground” and store only internal failures in their ground memory while type 1 and 3 systems are still considered ”in flight”. C- Between 80 kts + 30 seconds and 5 minutes after engine shutdown In this phase, the storage of failures is done in the ground memory for all the systems. Storage of internal failures of all systems in their ground memory. All CFDS functions are available. D- Between 5 minutes after engine shutdown and first engine start Note that when the aircraft power supply is turned on, the CFDS starts in this phase. Information of the last flight is always there because stored in non-volatile memory. Storage of internal failures of all systems in their ground memory. All CFDS functions are available on request. Transmission of last flight failures for type 2 systems, even after electrical power restart.

E- After fist engine start (+ 3 minutes) and before 80 kts At engine start (after 3 minutes), the contents of the last leg report is stored in the previous legs report. The leg number is then incremented. Type 1 systems store internal failures in their flight memory (External failures not stored). Type 2 systems store internal failures in their ground memory (still on ground). NOTE: Transfer of the LAST LEG REPORT into the PREVIOUS LEGS REPORT is done at engine start, both inthe CFDS and in type 1 systems. F- After 80 kts and before lift-off After 80 knots, type 1 systems receive a signal from the CFDIU and then store all internal and external failures in the flight memory. Type 1 systems store internal and external failures in their flight memory. Type 2 systems store internal failures in their ground memory (still on ground). G- After lift-off and in flight The fault memories in type 2 systems are erased at each ground/flight transition. Transmission and storage for internal and external failure for type 1/2/3 systems in their flight memory.



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CFDS

FLIGHT

GROUND

FLIGHT

LEG X

LEG X+1

5 MIN AFTER 2 ND ENG SHUTDOWN

FLIGHT TOUCH DOWN

A

80KTS +30S.

B

C

FLIGHT

1 ST ENG START

D

TAKEOFF

80 KTS

E

FLIGHT MENU

F

G

GROUND MENU

Figure 22

Flight Phases



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CFDIU FUNCTIONS

INTERFACES

MAIN FUNCTIONS Memory The CFDIU stores the failure messages and the ECAM warning messages in a non volatile memory. Management The CFDIU manages failure information and adds data such as GMT, DATE, LEG, FLIGHT PHASE to elaborate reports. Correlation If a computer internal failure is detected, the CFDIU achieves a correlation function that means it isolates or ignores the malfunctions of systems relating to this failure. Example: ”ADC FAILURE” causes ”NO DATA FROM ADC” in other computers. The CFDS will present only the initial failure in the last leg report. The function ”IDENT” will then present the systems affected by this failure. Monitoring The CFDIU scans permanently all input buses in order to detect a transmitted failure message. The CFDIU detects intermittent operation of the systems and adds (INTM) to the failure message. Detection The CFDIU can detect the nature of the failure by reading the ARINC words. Nature of failures: Internal Example: ”SDAC FAULT” External Example: ”FWC1: NO DATA FROM ADIRU1” Intermittent (INTM) added Class 3 (CLASS 3) added Messages requiring more investigation with the help of the trouble-shooting manual Example: ”CHECK EIU 1”

Clock The CFDIU permanently receives the GMT (UTC) and the date from the aircraft clock and then sends these two parameters to all type 1 systems. The GMT and date are used by the system BITE’s as well as the CFDIU for the various maintenance reports. FAC (Flight Augmentation Computer) The CFDIU receives the flight number and city pair from the FAC. The city pair (FROM/TO airports) is sent to the Management Unit (MU) of the Aircraft Communication Addressing and Reporting System (ACARS) and to the Data Management Unit (DMU) of the AIDS (Aircraft Integrated Data System). FDIU (Flight Data Interface Unit, part of Flight Recorder System) The CFDIU receives the aircraft identification from the Flight Data Interface Unit and sends this parameter to all type 1 systems. The CFDIU is used as an interface between the FDIU and the FWC (Flight Warning Computer) to send some FDIU class 2 failures to the FWC in order to constitute the maintenance status. FWC (Flight Warning Computer, part of ECAM) The CFDIU receives the flight phases and ECAM warnings from the FWC. The ECAM warnings are used by the CFDIU to generate the LAST or CURRENT LEG ECAM REPORT. Only PRIMARY failures, INDEPENDENT failures and CLASS 2 failure messages (Maintenance status) are received. DMU (Data Management Unit, part of AIDS) The CFDIU is used as an interface between the DMU and the FWC to send some DMU class 2 failures. DMU class 2 failures are used for the maintenance status on the ECAM.



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FWC

FAC FLIGHT NUMBER

FLIGHT PHASES CLASS 2 FAILURES

ECAM WARNINGS

C

F

D I

U

INTERNAL CLOCK FLIGHT PHASE

CLOCK

GMT

A/C IDENT

FDIU

DMC

A/C IDENT

ENGINE SERIAL NUMBER

GMT

CITY PAIR

CLASS 2 FAILURES

DATE

MAIN CHANNEL

BACKUP CHANNEL

* MEMORY * DETECTION * MANAGEMENT * CORRELATION * MONITORING

CITY PAIR

DATE

TYPE 1 SYSTEMS

MU

Figure 23

CLASS 2 FAILURES

CITY PAIR

DMU


EVMU

CFDIU Functions

______________________________________________________________________________________________________________________________________________________________________________________________



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CFDIU FUNCTIONS (CONT.) INTERFACES (CONT.) DMC (Display Management Computer, part of EIS) The CFDIU receives the Engine Serial Number from the DMC and sends this parameter to the EVMU (Engine Vibration Monitoring Unit). MU (Management Unit of ACARS) The ACARS Management Unit receives the city pair from the FAC through the CFDIU. EVMU (Engine Vibration Monitoring Unit) The EVMU receives the Engine serial number from the DMC through the CFDIU. The DMC receives it from the ECU (Engine Control Unit). ABNORMAL OPERATION Clock Back Up If the aircraft clock fails, the CFDIU takes over and its internal clock sends GMT (UTC) and DATE on the output bus to all type 1 systems. Upon power-on after A/C clock failure, the item ”GMT/DATE INIT” is added to the CFDS Menu. This option enables GMT and date initialization. BACKUP Mode In BACKUP Mode, only the main computers are available and only the ”SYSTEM REPORT/TEST ” function is available. In the event of main channel failure: If this failure is serious (Power Supply or Microprocessor) the backup channel takes over. Only ”BACKUP MODE” is displayed on the CFDS menu. No function is available in flight. If this failure is minor, the item ”BACKUP MODE” is added to the CFDS menu. This enables the access to the backup channel. The main channel remains available.



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FWC

FAC FLIGHT NUMBER

FLIGHT PHASES CLASS 2 FAILURES

ECAM WARNINGS

C

F

D I

U

INTERNAL CLOCK FLIGHT PHASE

CLOCK

GMT

A/C IDENT

FDIU

DMC

A/C IDENT


GMT

CITY PAIR

CLASS 2 FAILURES

DATE

MAIN CHANNEL

BACKUP CHANNEL

* MEMORY * DETECTION * MANAGEMENT * CORRELATION * MONITORING

CITY PAIR

DATE

TYPE 1 SYSTEMS

MU

Figure 24

CLASS 2 FAILURES

CITY PAIR

DMU


EVMU

CFDIU Functions

______________________________________________________________________________________________________________________________________________________________________________________________



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COMPONENTS Student Notes:



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Figure 25

Components Location

______________________________________________________________________________________________________________________________________________________________________________________________



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A319/A320/A321

______________________________________________________________________________________________________________________________________________________________________________________________



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Figure 26

PFR Filter-Loading

______________________________________________________________________________________________________________________________________________________________________________________________

Revision No : 02US 84 wd FRA Issue Date : 21/05/2013

01.02.95


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Figure 27 FDS Menu (new)

______________________________________________________________________________________________________________________________________________________________________________________________



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Figure 28

PFR Messages Filter Activation

______________________________________________________________________________________________________________________________________________________________________________________________

Revision No : 02US 84 wd 01.02.95 FRA Issue Date : 21/05/2013


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Figure 29 Data Base Identification

______________________________________________________________________________________________________________________________________________________________________________________________



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Figure 30

Non filtered PFR

______________________________________________________________________________________________________________________________________________________________________________________________



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Figure 31

Print of a Filter Data Base

______________________________________________________________________________________________________________________________________________________________________________________________



Page 769

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Figure 32

Filtered PFR

______________________________________________________________________________________________________________________________________________________________________________________________



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DATA LOADING SYSTEM



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Figure 33

Data Loading System Locations

______________________________________________________________________________________________________________________________________________________________________________________________



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Figure 34

Data Loading System Schematic

______________________________________________________________________________________________________________________________________________________________________________________________



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Training Manual A 319/320/321 ATA 31 Indicating/Recording 31-33 31-36 31-35

DFDR System AIDS Printer


Lufthansa Issue: October 1997 Technical Training GmbH For Training Purposes Only Book No: A 319/320/321 31-30 Level 3 e Lufthansa Base Lufthansa 1995 ______________________________________________________________________________________________________________________________________________________________________________________________



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ATA31 INDICATING/RECORDING SYSTEMS

______________________________________________________________________________________________________________________________________________________________________________________________



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31-30

CENTRALIZED FAULT DISPLAY SYSTEM (CFDS) AND DATA RECORDING SYSTEM

GENERAL The subchapter 31-30 Centralized Fault Display System (CFDS) and Data Recording System is divided into the parts listet below: 31-32 Centralized Fault Display Interface (CFDIU) 31-33 Digital Flight Data Recording System Interconnection (FDIU, DFDR, LA, QAR) 31-34 DFDRS Input Interface 31-35 Multifunction Printing (Printer) 31-36 AIDS Interconnection (DMU, DAR) 31-37 AIDS Input Interface. The CFDS is a system on its own. All the other parts can be summarized as ”Data Recording System”.

CFDS The CFDS is a centralized maintenance aid system which gives the maintenance technicians a means to read the maintenance information related to most of the aircraft systems and to initiate the tests of these systems from the cockpit.

Data Recording System This system is a centralized system to record the parameters from most of the aircraft systems. These recordings serve to determine the causes of incidents or accidents and are also an aircraft maintenance and monitoring aid. Two systems are isnstalled: DFDRS as a basic system and AIDS as an option.

HAM US/E Bo 5.3.97 Page: 2 ______________________________________________________________________________________________________________________________________________________________________________________________ Revision No : 02 Issue Date : 21/05/2013


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AIDS

DFDRS

DFDR DMU

FDIU

QAR

Figure 1

Aircraft Data System - General

______________________________________________________________________________________________________________________________________________________________________________________________



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31-33

DFDRS

GENERAL The main function of the DFDRS (Digital Flight Data Recorder System) is to convert various critical flight parameters into a recordable Form and to record them on a magnetic tape or a solid state memory. The stored data is also applicable to monitor the condition of the connected aircraft Systems. The system design covers the basic DFDRS. This includes the units and parameters which are necessary for the mandatory requirements and an additional part to standardize the installation for different customers. The electrical characteristic is in compliance with ARINC 717.



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Ground Control Pushbutton

MCDU

MCDU

Event Marker Button

Figure 2

Component Location (Cockpit)

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ARCHITECTURE FDIU The FDIU (Flight Data Interface Unit) is connected to different aircraft systems. DATA (parameters) are received in discrete and digital form. The FDIU collects these parameters and converts them for internal processing. A standardized set of flight critical parameters are transmitted in serialized digital form to the DFDR. The FDIU is also connected with the CFDIU via ARINC 429 for fault-transmission and test-activation.

Advice: The minimum equipment of a basic DFDRS (FDIU, LA, CP and EVENT) must be installed on each aircraft. This is to meet the requirement of the authorities for recording of mandatory parameters.

DFDR The FDIU-processed parameters are stored on the Recorder in data frame cycles. The Digital Flight Data Recorder has the capability to store the last 25 hours of data. Two different technologies are used for the data storage. The older version is the Tape Recorder, the newer the Solid State Flight Recorder. LA The Linear Accelerometer is installed to provide the FDIU with acceleration data appearing in the center of gravity. The SDAC digitizes the analog signal and sends it to the FDIU via ARINC 429 output bus. CP With GND CTL (Ground Control) button pushed ON, the Systems DFDR, CVR and QAR (if available) will be activated on GROUND (that means no Engine running). EVENT A push of the EVENT button sets a marker on the data frame of the activated Recorders DFDR and QAR. QAR An optional QAR (Quick Access Recorder) stores the same data as the DFDR. The installed QAR-Cassette has the maximum storage capacity of 50 houers. Before reaching the end of the tape track, the Casette has to be changed by the maintenance.



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ARINC 429 SYSTEM BUSSES

DFDR

LA

FDIU Control Panel

FDIU

Event Marker PB

QAR Cassette

Figure 3

System Block Diagram

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FDIU The function of the FDIU and the electrical interface complies with ARINC 717. If more than one data bus with the same content, e.g. SDAC 1 and SDAC 2, is connected to the FDIU, the data from system 1 is recorded on the DFDR. On A/C 033-099, 101-199: Test Connector To enable the connection of a Portable MCDU or a Portable Data Loader (PDL), a test connector is installed on the front panel of the FDIU. DFDR-CVR Synchronisation The full 32 data bit word received from the GMT clock bus is used to generate a frequency modulated output. This time code word is send to the CVR via audio output.

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FDIU (Hamilton)

FDIU (SFIM)

AMM 31-33-00

Figure 4

FDIU

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FDIU INTERFACE The FDIU prefers indication data and changes it to recordable formats. Inputs The FDIU receives data from FWCs SDACs DMCs Clock BSCUs FCDCs Additionally, it can be interrogated by the CFDIU. Only the SFIM FDIU receives data from the BSCUs (Parking and Steering Control Unit) and FCDCs (Flight Control Data Concentrator).

Outputs The FDIU sends data for recording to DFDR QAR. The CFDIU receives BITE data in usual ARINC 429 format.



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connected to SFIM FDIU only

connected to SFIM FDIU only

FDIU ( SFIM )

Figure 5

FDIU Interface

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UNDERWATER LOCATOR BEACON

DIGITAL FLIGHT DATA RECORDER General The DFDR has the capability to store the last 25 hours of data collected by the FDIU. The data is stored on a magnetic tape which is in a fire and shock protected box which is painted red. The data is stored sequentially on 8 tracks. The reverse tape motion and the tracks are switched automatically.

General An ULB is directly attached to the front panel of the DFDR. The beacon actuates on immersion in water down to a depth of 6000 meters. It has a detection range of 1800 to 3600 meters. You can service the ULB without disassembly of the DFDR. Maintenance has to be performed at determined time intervals to replace the battery of the ULB.

Interface The data input andoutput is connected to the FDIU. The format on this line is coded in Harvard Biphase. The output port for playback data is provided with the same data speed. Status signals indicate the condition of the DFDR.

Playback Data The playback data is generated directly from the received data. The transmitted playback data stream is interrupted in case of internal failures detected by BITE functions. Correct recording of data is ensured by comparing the memory stored data with the data read blockwise from the tape.

BITE The BITE functions include,beside the read/write data verification, control of tape motion, track switching and BI-Phase decoding. The DFDR BITE discrete is connected to the FDIU.

Status Discrete A status discrete is sent to the SDACs when the DFDR does not work. This happens, when the DFDR is defective the DFDR receives no data the FDFR is not supplied with power.



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ULB

Figure 6

DFDR with magnetic tape

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UNDERWATER LOCATOR BEACON

DIGITAL FLIGHT DATA RECORDER General The DFDR is a solid state flight data recorder in compliance with ARINC 717. The DFDR stores all aircraft information in erasable EEPROM devices. The DFDR has no moving parts requiring replacement and therefore requires only a minimum of maintenance. The recorder has the capability to store all data which the FDIU has collected over the last 25 hours. The DFDR is painted in red.

General An ULB is directly attached to the front panel of the DFDR. The beacon actuates on immersion in water down to a depth of 6000 meters. It has a detection range of 1800 to 3600 meters. You can service the ULB without disassembly of the DFDR. Maintenance has to be performed at determined time intervals to replace the battery of the ULB.

Inputs The DFDR receives all data via the FDIU.

Outputs The DFDR sends playback daten to the FDIU for control purposes. A BITE discrete is sent to the FDIU, when the DFDR detects an internal fault.

Status Discrete A status discrete will be sent to the two SDACs, when the DFDR does not operate. This happens in the following cases: DFDR faulty (same reasons like for BITE discrete) DFDR receives no input data DFDR is not supplied with power.



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ULB

Read out Connector

Figure 7

DFDR with solid state memory

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LINEAR ACCELEROMETER The task of the LA is to measure the acceleration of the aircraft in all three axes. It is installed in the center of gravity installiert. The power supply is 28 V DC.

Outputs The LA sends the analog outputs for the three acceleration values to the SDACs. The SDACs digitalize these inputs and send them to the FDIU.



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28 V DC

to SDACs

Figure 8

Linear Accelerometer

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QAR Function The purpose of the QAR is to store serial data (equivalent to the DFDR data) on a tape cassette for on ground performance, maintenance or condition monitoring task. A door, incorporated on the front panel of the QAR, givis quick access to the cassette. Three windows are installed on the front panel for checking the indicator lamps The data is recorded on 12 tracks on the cassette. The detection BOT / EOT (begin of tape / end of tape) automatically switches the tracks. At the beginning of track No. 11 (that means 8.33 hours before end of tape) the TAPE LOW flag is activated. At the end of the tape or in case of QAR-Failure the FLAG Light is activated. The light FIRST TRACK comes on automatically if a new cassette is installed. If the light does not comes on automatically after cassette change, first track can be set manually by pushing the FIRST TRACK push button on the QAR.

Change of the Cassette The operational procedure to change the cassette is as follows: open the recorder door, lift the cassette extraction/locking lever, extract the partially ejected cassette, insert a new cassette with the shutter face first and the housing button at indicator side, push the cassette fully home, until it is properly engaged in its chamber, lower the extraction/locking lever, depress the FIRST TRACK pushbutton if necessary

Flag The red FLAG lamp indicates internal faults, but not for door open. The recorder operation is correct, when: a cassette is installed, data is available at recorder input, the recorder head power is supplied, phase-loop locked from the drive-motor servo, the track storage power battery level is correct.



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Figure 9

QAR

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POWER INTERLOCK The FDIU and QAR (optional) are supplied directly via the CB 8TU. The power source to the DFDR are controlled by the Relays 8RK in case of one ENG running and by the Relay 6RK in case of FLIGHT condition. Parallel to this the QAR gets a RUN CONTROL Signal. The purpose of the TD-Relay 10RK is to keep the units DFDR, CVR and QAR activated 5 Minutes after second ENG SHUT DOWN. A push of the GND CTL Button activates the 6RK-Relay only if no Engines are running. An electric latch holds the override function. The blue ON-Light comes on. If the GND CTL button is pushed again or in case of an ENG-Start the override function returns to normal and the ON-Light goes off.



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Figure 10

Power Interlock

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DFDRS STATUS Indication The GND/CTL button has to be pressed for FAULT/STATUS-Advices. The DFDR Status line (FAIL) and the FDIU FAIL line (BITE) are connected to the SDAC’s for WARNING Messages on the ECAM E/W-Display. The FDIU is able to send some messages to the CFDIU by collecting the Maint Flag (BITE) discrete and the Play Back Data from the DFDR and the TAPE LOW andSTATUS discretes from the QAR. In case of a Class II Fault, the FDIU transmits a failure message to the CFDS. These failures are not indicated to the crew in flight, but are the subject of an ECAM report on the ground after shut down of the engines. If a Class III Fault occurs, the related flag is set in the Fault Memory of the FDIU (up to 30 faults). This fault information is sent to CFDIU and can be displayed on the MCDU screen. The following FAULT/STATUS-Advices are shown on the E/W-Display for DFDR- and FDIU-Faults (Class I), POST FLIGHT REPORT (Class I and II), STATUS-Page under MAINTENANCE-Advices (Class II), MCDU’s under the CFDS-INSTRUMENTS-Menu (Class III). Note Some DFDRs contain a FAULT / BITE-Indicator on the front panel. The QAR ’’TAPE LOW’’ condition is written on the POST FLIGHT REPORT, but on the STATUS-Page it points out only the MAINTENANCE-Message ’’QAR’’. After cassette change the MAINTENANCE-Advice ’’QAR’’ disappears only after next ENG-Start.



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NO INPUT DATA NO PWR

STATUS

BITE 2

DFDR FAULT FDIU FAULT

RECORDER

Flight Warning System

Figure 11

2

DFDR FAULT (BITE) FDIU FAULT (BITE) ACC FAULT

DFDR FDIU ACC

QAR FAULT QAR TAPE LOW

QAR

1

DFDRS Status

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LOCATION Be carefulwhen you change the following LRUs: DFDR: Some System Reports must be activated before you remove theFlight Recorder. ULB: The Underwater Locator Beacon is a part of the A/C. After the FRchange the ULB has to be installed on the new DFDR. LA : After the Accelerometer-change the capability of the X-, Y-, Z-detection has to be checked by using the AIDS-’’Label Call Up’‘-method (see chapter 31-36 AIDS). Sometimes it is difficult to find the Passenger Compartment Floor Panel under the carpet, where the LA is located. QAR: A Cassette-Change has to be made with powersourced QAR. Attention: Never reinstall a used Cassette !

______________________________________________________________________________________________________________________________________________________________________________________________ P



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Linear Accelerometer

ULB

DFDR QAR

Cassette

Figure 12

FDIU

DFDRS Component Location

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BITE TEST CFDS access to the Digital Flight Data Recording System is done via selection of FDIU in the INST-Menu.

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Figure 13

DFDRS BITE Access

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FDIU-Menu On the schematic, you see some functions selected.

Tasks Following tests are described on the AMM: Operational Test of the Power Interlocks and Status Monitoring: TASK 31-33-00-710-001. QAR Self Test: TASK 31-33-00-710-005. FDIU Fault Monitoring Test via MCDU: TASK 31-33-00-710-006. DFDR Self Test: TASK 31-33-00-710-007. Functional Test of the ULB: TASK 31-33-00-720-001. Operational Test of the Recorder Control with CFDS: TASK 31-33-00-710-004

After Removal / Installation of the LRUs the following Tests must be done: LA Do an accelerometer test on the MCDU (PARAM LAB Call Up) Subtask 31-33-16-700-051-A. FDIU Do the operational test of the FDIU (Ref. TASK 31-33-00-710-006). QAR NOTE: If the sticker of the cassette is in the incorrect position, the QAR does not operate.Thus we recommand the replacement of the cassette with the QAR energized. - SERVICING, Replacement of the Cassette with QAR energized TASK 31-33-52-600-001. with QAR not energized TASK 31-33-52-600-002. - Do the operational test of the QAR (Ref. TASK 31-33-00-710-005). DFDR Do the operational test of the DFDR (Ref. TASK 31-33-00-710-007). ULB - Discard ULB TASK 31--33-55-920-001. - Replacement of the ULB Battery and Functional Test of the ULB is not a Maintenance Task by DLH.

a ______________________________________________________________________________________________________________________________________________________________________________________________ Revision No : 02 Issue Date : 21/05/2013


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Figure 14

FDIU Menu

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31-36

AIDS

GENERAL The main function of the AIDS (Aircraft Integrated Data System) is to process continuously condition monitoring for various A/C systems to do an Engine Condition Monitoring (EGM), Aircraft Performance Monitoring (EPM), APU Condition Monitoring (ACM). A part of the process is used for creating various A/C condition reports. These reports are available for maintenance purposes : as a hard-copy from the PRINTER, on ground via ACARS (if installed). A Remote Print Button is installed on the cockpit pedestal. An activation of this button causes a flight phase dependent Print Report.



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Figure 15

AIDS Components (Cockpit)

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DESCRIPTION The MCDU are connected to the AIDS and used for programming and controlling the system. The printer is connected to the AIDS and used for hard copies of the results and reports. The DMU is able to identify A/C type information (ID code) on the FDIU bus. SFIM DMU only: An integral part of the DMU is the optional Smart AIDS Recorder SAR. This function is based on a nonvolatile ’’Solid State Mass Memory’’ module. SAR data are retrievable with floppy disks by using the MDDU (Multi Disk Drive Unit). Because of the use of SAR, no QAR is installed. The various communication interfaces for operator dialogue are mostly programmable. For example, reports can be either printed out, transmitted to the ground via ACARS, or retrieved by the use of a floppy disk via the MDDU. The available communication channels are as listed below: MCDU - Manual requests of report and SAR/DAR recording. - Display of list of stored reports and SAR files. - Online display of selected A/C parameter. - Various control and reprogramming menus. PRINTER - Manually initiated (by MCDU) print out of reports. - Automatic print out of reports. - Print out of MCDU screens. - Print out of software load messages. MDDU - Manually initiated (by MCDU) retrieval of reports and SAR files. - Automatic retrieval of reports and SAR files. - Load of DMU software.

ACARS - Manually initiated (by MCDU) download of reports. - Automatic download of reports. - Upload of request for report generation. - Upload of programming messages. DAR (optional) - Manually initiated (by MCDU) recording of AIDS data. - Automatic recording of AIDS data. To initiate manually some specific reports a Remote Print Button is located on the pedestal in the cockpit. Also SAR recording is triggered through the print button. The report/SAR channel assignment of the Remote Print Button is GSE programmable (Ground Support Equipment).



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Figure 16

SYSTEM BLOCK DIAGRAM

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LOCATION The DMU (Data Management Unit) and DAR (Digital AIDS Recorder), if installed, are located in the electronics rack. The DAR records mainly free programmable parameters from the DMU. The DAR is physically identical to the QAR, but the tape motion is 4 times higher. Two different DMUs are in use: Hamilton or SFIM.



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DMU DMU (Hamilton)

DAR

Cassette

DMU (SFIM)

Figure 17

Location

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POWER SUPPLY The DMU and the DAR (when installed) are connected to the 115 V AC bus via the circuit breaker 4TV (AIDS). For status indication the DAR is also connected to the 28 V DC bus via the circuit breaker 9TU (ACCLRM).



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*

*On older aircraft 103 XP AMM 31-36-00

Figure 18

Power Supply

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MAIN MENU (HAMILTON DMU) By pushing the following button on the MCDU you get access to the AIDS menu : MCDU MENU button, LSK 4L (AIDS). The AIDS menu now provides the sub’ menu by pushing the following LSK’s : 1L < CALL UP PARAM LAB (B, C, D). Parameters which do not have an alpha code can be selected by their Parameter-Number in the following sequence: - EQ, Equipment Identifier (HEX), - SYS, System No. (1, 2, 3), - LAB, Label (OCTAL), - SDI, Source and Destination Identifier (BINARY), - DATA BITS, counts of used data bits. 2L < DOCUMENTAT/ PROGRAMMING (F, G). This menu gives information of the software version and enables programming of - the DAR and - the Triggering for the Print Reports. Prior to this, the ’’ENTER CODE’’ has to be inserted. Only the appearance of the ’’Asterix’’ (*) enables programming.

5L, 5R *RUN, STOP* of DAR Both LSK‘s make it possible to START/STOP the DAR manually, if installed. 6L < MAN REQ REP (H) By activating this menu it is possible to do a manual Report-Downlink (SEND) via ACARS or a manual Report-PRINT (hard copy). 6R > STOR REP (J) This menu makes it possible to create one of the last 10 stored reports (PRINT or SEND downlink via ACARS).

3L < START MENU FOR SPECIAL REPORT (E). This menu enables a selection of 3 free programmable Reports (16, 17, 18), with the possibility of RUN / STOP / PRINT. 1R > CALL-UP PARAM ALPHA (L, M, N). For about 200 parameters the selection is possible by keying in the appropriate alpha code. The alpha codes are described in the parameter list. 2R > LIST OF PREV REP (K). This menu shows the last 20 created reports in a list.



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Figure 19

Main Menu (Hamilton DMU)

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AIDS BASIC MCDU UTILIZATION



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Figure 20

Alpha Call Up

______________________________________________________________________________________________________________________________________________________________________________________________



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ALPHA CALL UP



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Figure 21

Alpha Call Up (cont.)

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ALPHA CALL UP LIST The alpha call up list is in the AMM. The alpha codes are sorted alphabetically. Detailed information is given to each parameter as follows : EQ (Equipment Identifier of the transmitting system), System (1, 2 or 3), Label (of the parameter), SDI (Source and Destination Identifier), Name of the parameter, Unit (engineering unit, if available).



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Figure 22

Alpha Call Up List

______________________________________________________________________________________________________________________________________________________________________________________________



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LABEL CALL UP



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Figure 23

Label Call Up (cont.)

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LABEL CALL UP (CONT.)



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Figure 24

Label Call Up (cont.)

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DMU Input Parameter List In the Column named ”SYSTEM” the DMU Input Parameter List mentions the computers which send data to the DMU. The column named ”EQ” lists their Equipment Identifier. So the first data to be inserted into the MCDU scratchpad are shown. The column ”CHAPTER” shows where to find the other data. The Parameter List can be found in the AMM 31-37-00.



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Figure 25

Parameter List

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REPORTS General The following AIDS standard reports are defined: Engine Cruise Report <01> Cruise Performance Report <02> Engine Take-Of f Report <04> Engine Report O/R <05> Engine Gas Path Adsvisory Report <06> Engine Mechanical Advisory Report <07> Engine Divergence Report <09> Engine Start Report <10> Engine Run Up Report <11> APU MES/IDLE Report <13> APU Shutdown Report <14> Load Report <15> Programmable Report <16>, <17> and <18> ECS Report <19>

Manual Request Reports Each AIDS report may be printed and/or transmitted to ACARS by manual request via the MCDU. In this case the report will be generated at once. The procedure is as follows: 1. Select ’AIDS’ on MCDU MENU. 2. Select ’MAN REQ REP’ mode on the AIDS initial menu. A list of all available AIDS reports is presented on the screen. If the printer is available and not busy, a star appears besides each report on the right side. If ACARS is available, a star appears besides each report on the left side. Activating the scroll key will cause all 10 lines to be rotated. In the direction indicated in order to present the next five reports. 3. Start the generation and the print out of the desired report by activating the appropriate LSK which is on the right of the screen (below ’PRINT’). If the report is sent to ACARS within 2 seconds the appropriate LSK on the left side has to be activated (below the ’SEND’). The stars then disappear from the screen as long as the printer and/or ACARS are occupied.

Reports 16, 17 and 18 can be programmed by the airline engineering.



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Figure 26

Manual Request Report

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Previous Reports The menu Previous Reports displays a list of the latest 20 generated Print Reports, with information about the reason for the generation the time of the generation the flight when the report was generated.



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Figure 27

Previous Reports

______________________________________________________________________________________________________________________________________________________________________________________________



Page: 55

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Stored Reports The menu Stored Reports allows access to the latest 10 generated Print Reports. Each report which is listet can be selected and printed (PRINT) with the left LSKs or selected and sent (SEND) via ACARS with the right LSKs. The message ’DNLKD’ (downlinked) confirms the datalink-Xmission via ACARS-MU to the ground station. ’IN ACARS’ means that data have been transferred to the ACARS-MU, but the ground station did not confirm. ’PRINTED’ says that the report has been printed automatically or by manual request.



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Figure 28

Stored Reports

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PRINT REPORT STANDARD HEADER DESCRIPTION



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Figure 29

Print Report Standart Header

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GENERAL DATA 1



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Figure 30

Print Report Standart Header (cont.)

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GENERAL DATA 2



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Figure 31

Print Report Standart Header (cont.)

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SOFTWARE LOAD STATUS The procedure to display the actual software on a MCDU is as follows: 1. Select ”AIDS” on MCDU Menu 2. Select ”DMU PROG/DOC” mode on the AIDS initial menu (LSK 2L). The software version number of the DMU and OBRMs are displayed. Also in addition a seperate customer software version number is presented which describes the state of the actual customer modifications.



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Figure 32

Software Load Status

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ADJUSTMENT / TEST (HAMILTON DMU) Maintenance Practices Update the APU hours and cycles into the DMU after APU or DMU replacement. TASK 31-36-00-740-004 Print out the APU hours and cycles TASK 31-36-00-740-006 Print out the Engine hours and cycles. TASK 31-36-00-740-007 Adjustment / Test Tasks Test of the DMU Test of the DAR

TASK 31-36-00-710-001 TASK 31-36-00-740-003

When you do the test of the Hamilton DMU, you have to select: LAST LEG REPORT PREVIOUS LEGS REPORT LRU IDENTIFICATION CLASS 3 FAULTS Look if the selected menus are displayed. Then you must select TEST. The next step is RESULT POWER UP TEST, then you have to select CREATE TEST.



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Figure 33

AIDS BITE Test (Hamilton DMU)

______________________________________________________________________________________________________________________________________________________________________________________________



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MAIN MENU (SFIM DMU) By pushing the following button on the MCDU you get access to the AIDS menu : MCDU MENU button, LSK 4L (AIDS). The AIDS menu now provides the sub’ menu by pushing the following LSK’s : 1L < CALL UP PARAM This menu allows a access to the following sub-menus - < PARAM LABEL CALL-UP Parameters which do not have an alpha code can be selected by their Parameter-Number. - < PARAM ALPHA CALL-UP For about 200 parameters the selection is possible by keying in the appropriate alpha code. The alpha codes are described in the parameter list. 2L
1R LOAD STATUS > The ’Software Loading‘ page displays the actual status of the software loading (TRANSFER IN PREGRESS or NO TRANSFER IN PROGRESS). If the AIDS is available, the LOAD STATUS is automatically displayed on the MCDU screen. 3R LIST OF PREV REP > This menu shows the last 20 created reports in a list. 4R STORED REPORTS > This menu makes it possible to create one of the last stored reports (PRINT or SEND downlink via ACARS). 5R MAN REQST REPORTS > By activating this menu it is possible to do a manual Report-Downlink (SEND) via ACARS or a manual Report-PRINT (hard copy) or store the data. 6L, 6R *RUN, STOP* of DAR Both LSK‘s make it possible to START/STOP the DAR manually, if installed.

3L < SAR The SAR stores data in a 2 MByte Solid State Mass Memory (SSMM). The data are compressed in the storage and organized in one SAR File. 4L < MICRO 3 Micro 3 programming is not applicable. 5L < REMOTE PRINT The Remote Print page displays the Report Nunber, Engine Number, DMU Internal Flight Phase and incrementation of report counter (Y= incremented, N= not incremented). By pushing the REMOTE PRINT BUTTON in the cockpit the flight phase related Print Report will be activated.



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Figure 34

Main Menu (SFIM DMU)

______________________________________________________________________________________________________________________________________________________________________________________________



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PARAMETER CALL UP Alpha Call Up For all the parameters, which are listed in the alpha call-up table the selection is possible by their aplha codes. If parameters from two systems are available, both parameters are displayed on the MCDU upon a single alpha call-up code entry. The displayed parameter values are refreshed once per second. All numeric type alpha call-up parameters are displayed on the MCDU in engineering units. The display format is a floating point representation of max. 6 characters including the decimal point and the ,-, sign if applicable. In case of positive numbers without decimal point all 6 characters are available for digits. Display of loading zeros are suppressed. The applicable units are diplayed below the alpha call-up code on the MCDU screen. Alpha call-up parameters, which consist of a combination of sereval discrete bits or packed discretes are displayed in hexadezimal representation. In this case the indication of the applicable units of the parameter are ,HEX‘.



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Figure 35

Alpha Call Up

______________________________________________________________________________________________________________________________________________________________________________________________



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Label Call Up Parameters which do not have an alpha code can be selected by their Parameter-Number in the following sequence: EQ, Equipment Identifier (HEX), SYS, System No. (1, 2, 3), LAB, Label (OCTAL), SDI, Source and Destination Identifier (BINARY), DATA BITS, counts of used data bits. Additionally the number of data bits to be used for decimal representationare selectable. The parts of the parameter number are separated by a slash ,/,. Up to 2 parameters are displayed on one page. 8 pages car be seleceted via slew up/slew down button, which leads to a maximum number of 16 parameters to be monitored simultaneously. Example Parameter Call-Up with EQ and System Number: EQ/SYS/LAB/SDI = tA/2/156/Ol



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Figure 36

Label Call Up / Menus

______________________________________________________________________________________________________________________________________________________________________________________________



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PROGRAMMING MENU The AIDS DMU Programming Menu displays the A/C Type, DMU Partnumber, Operational Software Partnumber, Database Version and Revision Levels. Before selecting the ,


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Figure 37

Programming

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Figure 38

SAR

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Figure 39

Remote Print

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REPORTS Previous Reports The ’PREVIOUS REPORT’ page displays the Report Number, Trigger Code, Date, UTC and the Flight Leg (OO=current flight leg, XX=privious leg).

Stored Reports The ’STORED REPORTS’ page displays a list of all AI reports, which are generated and stored in the DMU report buffer. To display each report select the adjacent line key. To print out them push the print line key. All Stored Reports are deleted as soon as a new A/C ident (A/C tailnumber) is recognized by the DMU.

Manual Request Report The ’MAN REQST REPRT’ page displays a list of all AI reports, which can generated by selecting the adjacent line key. The 1L key is used to roll options ’PRINT, SEND (ACARS) or STORE’. An asterisk (*) is displayed at the beginning of each report, if only a single data set or format is defined for the related report.If multiple formats or data sets are defined for a report, the (<) character is displayed intead of the asterisk. If the report storage is full and a generated report is stored in the report buffer, the oldest free declared report is deleted to enable the storage on an actual report. In case that the report storage is full with no free declared reports the oldest report is printed and deleted if stored capacity is required. When a report is selected and complete generated it will be printed out or sent via ACARS.



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Figure 40

Reports

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PRINT REPORT STANDART HEADER DESCRIPTION General A standard header to be printed on each report is installed. The header data are taken at the time when the respective report is generated. Following symbols for the value fields are used: ’A’ = any character in the range from A..Z, ’1’ = digits possible (0, 1), ’9’ = any digit in the range from 0..9, ’X’ = any character or digit in the range from A..Z and 0 .. 9, ’.’ = checksum. Header Lines 1 to 3 The content is frely programmable to enable airline specific messages.

Header Lines 8 to 10 A/C ID (XXXXXXX) Aircraft Identification (Tail Number) DATE (AAA99) DATE (Month/Day) 01=JAN, 02=FEB... UTC (999999) Universal Time Coordinated (Hours/Minutes/Seconds) FROM TO (AAAA AAAA) Identification of City Pair FLT (XXXX) Flight Number (actual)



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Figure 41

Standart Header

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TEST BITE (SFIM DMU) General For Adjustment / Test you can find the following tasks: Test of the DMU TASK 31-36-00-710-001 Test of the DAR TASK 31-36-00-740-003 Test BITE The DMU contains an adequate Built-in Test Equipment (BITE> according to ARINC 604. The BITE is able to detect the failure occuring in the DMU. The board, the functional block or the integrated circuit in which the failure appears is described by the failure message. General Rules for the BITE All facilities of the already existing hardware and software that can reasonably be used to detect faults of Systems or system componets are made available for the fault isolation and detection function. In order to recognize transmission faults the ARINC 429 inputs of the DMU are monitored continuously for update, sign status matrix and if necassary parity. The BITE of the DMU is able to distinguish between system internal faults (DAR and DMU> and external faults (connected systems>. The equipment supplier is propose a fault isolation and detection concept that finally accepted by the purchaser.

From the AIDS Menu, you can Set: < LAST LEG REPORT, < PREVIOUS LEGS REPORTS, < LRU IDENT, < GND SCANNING < TROUBLE SHOOT DATA < CLASS 3 FAULTS, < TEST, < GROUND REPORT, < SOFTWARE LOAD STATUS. When you do the test of the SFIM DMU, you have to do the following steps: 1. Select TEST . 2. Select DMU POWER UP TEST. The message TEST OK must come on. 3. Press RETURN. 4. Select DMU BATTERY TEST. The message BATTERY OK must come on.

CFDS Menu Function You can set the MCDU to show the failures that occurred during the Last llight. The data come from the CFDS (Rel. 45-10-00). You can also start system tests from the MCDU. To show the AIDS control data on the MCDU or to Start a System test, you must: select the line key adjacent to ’CFDS‘ on MCDU Menu, select the line key adjacent to ’SYSTEM REPORT/TEST’ CFDS Main Menu, select the line key adjacent to ’RECORDER‘ on System Report/Test Menu, select the line key adjacent to ’


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Figure 42

AIDS BITE Test (SFIM DMU)

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TEST When you push the ,



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Figure 43

TEST / Software Load Status (SFIM DMU)

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SOFTWARE LOAD STATUS Software Loading Menu The ’Software Loading‘ page displays the actual status of the software loading (TRANSFER IN PREGRESS or NO TRANSFER IN PROGRESS). If the AIDS is available, the LOAD STATUS is automatically displayed on the MCDU screen. Possible Scratchpad Messages:

Database Loading During the OBPM loading, the System displays automatically the OBPM loading menu, if AIDS is seiected. It is unpossible to select the return key to select another menu At the end of the loading process the message ,LOAD COMPLETED‘ is displayed. If any error is occured during the load process the message ,LOAD FAILED‘ is displayed.



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Figure 44

Loading Menu

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REMOVAL/INSTALLATION DMU Before you remove the DMU, you have to do two tasks: print out the APU hours and cycles (TASK 31-36-00-740-006). print out the ENG hours and cycles /TASK 31-36-00-740-007). You need the printouts to reprogram the new DMU. When you have removed a SFIM DMU, you have to set the OPERATING /STORAGE switch on the front face to ”STORAGE”. On a new SFIM DMU, you have to set the OPERATING /STORAGE switch on the front face to ”OPERATING” before installation. To install a new DMU, do the following steps: 1. Close CB. 2. Do the updating of the APU hours and cycles data ( TASK 31-36-00-740-004). 3. Do the updating of the engine hours and cycles data ( TASK 31-36-00-740-005). 4. Do the DMU Software Load Procedure ( TASK 31-36-00-710-004). 5. Do the test of the DMU ( TASK 31-36-00-710-001).

DAR NOTE: If the sticker of the cassette is in the incorrect position, the DAR does not operate. Thus we recommend the replacement of the cassette with the DAR energized. Do the test of the DAR (Ref. TASK 31-36-00-740-003): 1. Close various CBs according to the AMM. 2. insert a new cassette into the recorder. During DAR tape loading, the READY and the BUSY indicators come on. After DAR tape loading, the BUSY indicator flashes and the READY indicator stays on. 3. On the MCDU, push MENU mode key. The MCDU MENU page comes on. 4. Push the line key adjacent to the AIDS indication. The AIDS menu page comes on. 5. For the Hamilton DMU installed: push the line key adjacent to the *RUN indication. The indication *RUN changed to RUN. The indication STOP changed to STOP*. 6. For the SFIM DMU installed: push the line key adjacent to the START* indication. The indication DAR=STOPPED changed to DAR=RUNNING. The indication START* changed to STOP*. 7. On the DAR, make sure that the cassette is running. 8. On the MCDU: For the Hamilton DMU, push the line key adjacent to the STOP* indication. The indication RUN changed to *RUN. The indication STOP* changed to STOP. For the SFIM DMU, the indication DAR=RUNNING changed to START*. 9. On the DAR, make sure that the cassette is not running and close the recorder door.



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DMU (Hamilton)

DMU (SFIM) DAR

Figure 45

DMU and DAR

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31-35

MULTIFUNCTION PRINTER

GENERAL The Printer (PRTR) is designed to print out on ’’high contrast low abrasive’’ paper reports coming from various systems such as AIDS, ACARS, FMGC, CFDIU, EVMU either on ground or in flight. Simple ’’one hand’’ in flight or on ground paper roll loading allows 90 feet printing, 3 rolls being stowed on the left rear cockpit wall.

TEST Functional Test The Functional Test is not available on the A/C but in the workshop.

Operational Test The operational function of the Printer is checked by creating a CFDS LAST LEG REPORT, TASK 31-35-00-710-001.

Servicing of the Printer TASK 31-35-22-600-001



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Figure 46

Printer

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SYSTEM DESCRIPTION The thermal line printer provides ’’on board print outs’’ for various aircraft systems, one at a time. When power is applied, the printer detemines which inputs are active and which specific system is connected to each active port. Also after a sequence of active port polling, a single inactive port is monitored so that a system which became active after initialization can be added to the active system list.



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Figure 47

Printer Interconnection Block Diagram

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Training Manual A319/320/321 ATA 34 Navigation 34-20 34-10

Standby Navigation ADIRS

34-58

Satellite Navigation


Lufthansa Issue: May 1998 Technical Training GmbH For Training Purposes Only Book No: A319/320/321 34-10 LEVEL 3 E Lufthansa Base Lufthansa 1995 ______________________________________________________________________________________________________________________________________________________________________________________________



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ATA 34

NAVIGATION

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34-20

STANDBY NAVIGATION SYSTEMS

STANDBY INSTRUMENTS PRESENTATION The standby navigation system enables the flight crew to check the navigation data provided by the Air Data Inertial Reference System (ADIRS). The standby navigation system comprises four instruments. Each providing different indications : Standby Compass for magnetic heading, Standby Horizon Indicator for attitude, Standby Airspeed Indicator (IAS) for airspeed, Standby Altimeter for altitude.



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Figure 1

Standby Instruments

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STANDBY DATA: ALTITUDE AND AIRSPEED General One standby airspeed indicator, one standby altimeter and one metric altimeter are directly connected to the standby pitot and static sources. The standby circuit can be drained by means of a water drain.

Standby Altimeter The standby altimeter is supplied with static pressure by the standby air data system to indicate the barometric altitude of the aircraft in feet. When the altitude is below 10,000 feet, the figure zero of the left drum is replaced by black and white stripes. The figure nine is replaced by an orange fire stripped zone. - The baro correction is displayed on a counter graduated in hecto Pascal. - A knob, located at the L corner of the indicator, enables the display of the reference baro correction in the range of 750 to 1050 h Pa. - Four manually adjustable white bugs are provided for manual altitude setting. The internal vibrator is supplied with 28VDC through a landing gear relay.

Metric Altimeter The metric altimeter is supplied with static pressure by the standby air data system to indicate the barometric altitude of the aircraft in meters. The barometric altitude is displayed by means of: - a pointer performing one revolution of the dial for 1000 meters. - a display counter made up of two drums displaying respectively the tens of thousands, and the thousands of meters. The altitude dial is calibrated from 0 to 1000 meters with 50 meters graduations - The baro correction is displayed on a counter graduated in hecto Pascal - A knob located at the L corner of the indicator enables the display of the reference baro correction in the range of 870 to 1050 h Pa.

Standby Airspeed Indicator The standby airspeed indicator contains a capsule-operated mechanism which measures the pitot /static pressure differential from the standby air data system and provides airspeed indication in terms of knots. The airspeed indication is displayed by means of : - A pointer which moves on a dial graduated between 60 kts and 450 kts. The scale is linear from 60 kts to 250 kts with 5 kts graduations and from 250 kts to 450 kts with 10 kts graduations. - Four manually adjustable white bugs provided for manual speed setting.



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Metric Altimeter

Standby Altimeter

Standby Airspeed Indicator

Figure 2

Standby Altitude and Airspeed

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STANDBY DATA: ATTITUDE AND HEADING Standby Heading The standby heading is performed by a magnetic compass that is an independent instrument which provides the flight crew with the A/C magnetic heading. It is installed on the top of the windshield center post and allows a check of the heading provided by the main sources of the heading system. It acts in standby when these systems are inoperative. The standby compass consists of a magnetic element rotating inside a compass bowl, immersed in a damping liquid. The magnetic element is linked to a graduated compass card which moves against a lubber line and gives the magnetic heading. Below the viewing window are two apertures marked N.S and E.W, allowing to achieve compensation by positioning the two small magnetized bars (compensator). Above the viewing window is a non-magnetic lamp assembly which provides illumination of the compass card.

Standby Attitude The standby attitude is performed by a gyroscopic horizon that is an independent instrument which provides the flight crew, with a constant indication of the aircraft attitude. It allows a check of the attitude provided by the main sources of attitude system. It acts in standby when these systems are inoperative. The standby horizon indicator is supplied with 28VDC from essential bus 401PP. A static inverter in the instrument converts this 28VDC into three phase alternate current to supply the gyroscopic motor. The gyro rotor rotates at high speed (> 23,000 RPM) around its vertical axis and provides the vertical provides the vertical reference.The fast resetting of the gyroscopic horizon can be activated by pulling the knob located in the lower R corner of the indicator. Failure Warning :The flag comes into view if a failure is detected in the electrical power supply or if the gyro rotor speed drops below 18,000 RPM.



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Figure 3

Standby Attitude and Heading

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34-10

AIR DATA / INERTIAL REFERENCE SYSTEM

GENERAL The main air data and heading/attitude data are provided by a air data inertial reference system (ADIRS). This configuration provides for triple redundant information for all inertial and air data functions. Each channel is isolated from the others and provides independent information. The Air Data/Inertial Reference System (ADIRS) provides the main air data and heading/attitude/navigation data to the aircraft systems. The main computers of the ADIRS are the three Air Data/Inertial Reference Units (ADIRU) which are controlled by the ADIRS Control and Display Unit (CDU). Attitude, heading and navigation data are displayed on the Electronic Flight Instrument System (EFIS) displays (Primary Flight Display (PFD); Navigation Display (ND)) and on the VOR/DME RMI which recopies the heading data.



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Figure 4

ADIRS Schematic

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DESCRIPTION Each ADIRU contains: - an Air Data Reference (ADR) portion. - an Inertial Reference (IR) portion. Power supply is common for ADR and IR.

ADR The Air Data Reference (ADR) portion of the Air Data/Inertial Reference Unit (ADIRU) provides main data sources which are air data references for the aircraft avionics systems. The ADR receives and processes the outputs of the. ADM, Air Data Module, TAT Probe, Total Air Temperatur Probe, AOA Sensor, Angle Of Attack sensor, It computes the aerodynamic parameters in the form of ARINC 429 low speed buses.

IR The Inertial Reference (IR) portion of the Air Data/Inertial Reference Unit (ADIRU) provides main data sources which are precision attitude, magnetic heading references and navigation data to the aircraft avionics systems. The ADIRU is interfaced with the ADIRS control and display unit (ADIRS CDU) for mode control and status annunciation.



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Figure 5

ADIRU Block Diagramm

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PROBES AND SENSORS General The aircraft is equipped with 3 Air Data/Inertial Reference Units (ADIRUs). Each ADIRU receives data from the four types of sensors after: - 3 pitot probes which provide total pressure data, - 6 static probes which provide static pressure data, - 2 Total Air Temperature (TAT) sensors which provide air temperature data, - 3 Angle Of Attack (AOA) sensors which provide angle of attack data of the aircraft. The TAT sensors and the angle of attack sensors are directly connected to the ADIRUs. The pitot probes and the static probes are connected to 8 Air Data Modules (ADM) which convert pressure data before they send them to the ADIRUs.



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Figure 6

Probes and Sensors

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Static Probes Each of the three systems (CAPT (1), F/O (2), STBY (3)) comprises two static probes which are linked to each ADR portion of the ADIRUs through five ADMs. The probe is protected from icing with a 28VDC heater circuit. The static probes linked to ADIRU 1 and ADIRU 2 are set at 48.64 below the fuselage datum line (Z=0). The static probes linked to ADIRU 3 are set at 29.5 below the fuselage datum line. Pitot Probes Each system comprises one pitot probe (CAPT (1), F/O (2), STBY (3)) which is linked to each ADR portion of the ADIRUs through one ADM. The probe is protected from icing with a 115VAC - 400 Hz heater circuit. The pitot probes 1 and 2 are set at 40.08 below the fuselage datum line (Z=0). The pitot probe 3 is set at 59.56 below the fuselage datum line (Z=0).

HAM US E Bo 3.11.95 Page: 14 ______________________________________________________________________________________________________________________________________________________________________________________________ Revision No : 02 Issue Date : 21/05/2013


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Figure 7

Pitot / Static Probes

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AIR DATA MODULE The term Air Data Module (ADM) refers to any remotely located LRU which senses pressure information and transmits it to the ADIRU in ARINC 429 format. The ADM’s are identical. Each ADM has one pressure input and several discrete inputs. The discrete inputs determine the ADM location and the type of pressure . On the data bus it provides digital pressure information type of pressure, ADM identification, BITE status. The ADMs are remotely mounted near and above the level of the pitot and static probes, this in order to make the ADM pneumatic plumbing self draining when the aircraft is stationary on the ground.



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Figure 8

Air Data Module

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TAT SENSOR The aircraft is equipped with two total air temperature sensors, with two sensing elements each. The sensing elements of the sensor have variable resistances. The TAT sensor 1 is linked to the ADR portion of ADIRUs 1 and 3, the TAT sensor 2 is linked to the ADR portion of ADIRU 2. The TAT sensors are set at 2.33 m from the nose and at 0.60 m of the aircraft axis below the fuselage. The TAT sensor 1 is located on the left side and the TAT sensor 2 on the right side. The air flow enters the scoop of the sensor, goes through a calibrated choke and flows over the hermetically sealed platinum resistance sensing element where the temperature is measured. The speed of the flow over the element is controlled by the choke in the element tube. Sensor The ADR portion is designed to operate with 500 ohms (at 0C) temperature sensor unit corresponding to the basic Callender - Van Dusen equation. To improve the accuracy of the sensor, a network of precision resistors is used. This technique is identified by the term Precision Calibration Interchangeability (PCI). Heat These sensors are heated with 115VAC through the probe heating system. The heating element must not be energized on the ground. The heating element is implanted in the scoop and strut and keeps the probe free of ice under the most severe icing conditions.



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Figure 9

TAT Sensor

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AOA SENSOR The aircraft is equipped with three AOA sensors. Two are located on the right side and one on the left side of the fuselage. Each of these AOA sensors is respectively linked to each ADR portion of the ADIRUs. The AOA sensors 1 and 3 are set at 6.08 deg. and 31 deg. below the fuselage datum line (Z = 0) on the left side. The AOA sensor 2 is set at 6.08 deg. below the fuselage datum line (Z = 0) on the right side.

Vane Type The angle of attack sensor is of the wind vane type. Its sensing element is a small wing which is positioned in the direction of airflow. The small wing is mechanically linked to a free turn-shaft which drives the devices transmitting the local angle of attack signal. These transmitting devices are made up of resolver transformers which convert the angular information into proportional electrical information (angle sine and cosine). The resolvers are supplied with a 26VAC signal. The same signal is also received by the ADIRU as a reference for the decoding of AOA values.Each sensor has 3 resolver outputs but only two are wired to the ADIRU. The characteristics of the resolvers are as follows : scale factor : 1 /Degree of AOA index reference : 0 resolver input = 25 AOA The whole mechanism is stabilized around the rotation axis. In addition, a damping device enables a satisfactory dynamic response to be obtained (filtering of mechanical oscillation).

Heating A self regulated heating element (CTP resistances: positive coefficient of temperature) inserted into the vane eliminates or avoids icing. It is supplied with 115VAC through the PHC.

Test The AOA sensor is equipped with a self test device which is activated by a 28VDC signal, from the ADR (through the relay 21FP1, 21FP2 or 21FP3) when the test is entered via the maintenance system (CFDIU and MCDU). The self test positions the vane at a resolver angle of +15 deg. (left side test) or -15 deg. (right side test). The mounting and wing of AOA resolvers determine the relationship between the measured resolver angle and indicated angle of attack. The ADRs receive the same 26VAC, 400 Hz reference as the AOA resolvers. This reference is common to both AOA resolver inputs 1 and 2.



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Figure 10

AOA Sensor

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SWITCHING PANEL The two selector switches AIR DATA and ATT HDG are rotary selector switches with 3 positions: CAPT/3, NORM and F/O/3. These selector switches are used for the functions listed below : AIR DATA SEL SW 15FP = 34-14-00 Selection of the ADR used by IR3 34-52-00 ATC mode S 31-68-00 DMC 22-85-00 FMGC ATT HDG SEL SW 13FP = 34-11-00 Power Supply 34-14-00 Selection of the ADR used by IR3 34-41-00 Weather Radar 34-57-00 VOR/DME RMI 31-68-00 DMC 22-85-00 FMGC.



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Figure 11

ADIRS Switching

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COMPONENT LOCATION The 3 ADIRU’s are installed in the Avionic Compartment. The ADIRS CDU is installed in the Cockpit. Two different types of ADIRU’s may be istalled in the aircraft, one is the Honeywell System, the other the Litton System.



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ADIRS CDU

HONEYWELL ADIRU

Figure 12

LITTON ADIRU

Location

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COMPONENT DESCRIPTION ADIRS CDU The CDU is a three-channel unit. Each CDU channel is dedicated to one separate ADIRU and includes the following features: a three-position mode selector switch. The modes are: - power off (OFF) - navigation (NAV) - reversionary attitude (ATT) an indicator announcing when the IR is aligning (ALIGN legend of IR annunciator) an IR fault indicator (FAULT legend of IR annunciator) a pushbutton switch to disable ADR output buses. It is a momentary action pushbutton switch an indicator announcing when the ADR output buses are turned off (OFF legend of ADR pushbutton switch) an ADR fault indicator (FAULT legend of ADR pushbutton switch). The following items of equipment are shaped between the three channels : a keyboard to enter the initial position in degrees, minutes and tenth of minutes or magnetic heading in the attitude mode two data pushbutton switches (ENT and CLR) with cue lights a liquid crystal display for selected parameters. the LCD has 16 digits and each digit has 14 segments

a DATA DISPLAY selector switch to select parameters for display on the LCD: - wind (WIND) - present position (PPOS) - true heading (HDG) - status of selected system (STS) - track and ground speed (TK/GS) - test values (TEST) a SYS DISPLAY selector switch with four positions: OFF, 1, 2, 3. The OFF Position disables the display of the CDU but the mode control of the ADIRUs remains active an ON BAT annunciator. The CDU contains three identical connectors. No cooling air is provided to the CDU. The CDU receives 28VDC power from the selected ADIRU to drive internal circuits and the data display. The 28VDC inputs are isolated from each other. The aircraft supplies 5VAC power for panel lighting/LCD backlighting and for annunciator lighting.

HAM US / E Bo 7.11.95 Page: 26 ______________________________________________________________________________________________________________________________________________________________________________________________ Revision No : 02 Issue Date : 21/05/2013


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Figure 13

ADIRS CDU

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ADIRU The Honeywell ADIRU is contained in a 10 MCU, the Litton ADIRU is contained in a 4 MCU case as defined in ARINC 600. The ADIRU has to be aligned on a special shelf in the avionics compartment in accordance with the installation design described in ARINC 738. This installation involves modification of the ARINC 600 standard to include three alignment pins and a floating connector. The ADIRU contains an ADR and an IR portion supplied by a common power (115VAC, 28VDC). ADR: Five resolvers can be used for the analog baro-correction and the AOA inputs. The ADR provides 8 ARINC 429 low-speed output buses (buses 5-8 are reserved for engine control). Each bus can drive 20 ARINC bus loads. IR: The gyros/accel sensors block contains three accels and three gyros mounted along each axis. This sensor block is supplied by a high voltage power supply provided by the IR portion. The IR provides 4 ARINC 429 high-speed output buses. Each bus can drive 20 loads.



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ADR

IR

Figure 14

ADIRU

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ELECTRICAL INTERFACE The ADIRS CDU contains three identical connectors. Each connector is linked to one ADIRU. The four annunciator discretes ADR OFF, ADR FAULT, IR ALIGN, IR FAULT are linked to the ADIRS CDU from the ADIRU, through the annunciator light test and interface boards. The CDU panel lighting and LCD backlighting are provided by bulbs supplied with 5VAC (from the A/C generation). The CDU exchanges data with the ADIRU. The data sent by the CDU can be used for the initialization of the IR portion. The data received by the CDU are displayed on the Liquid Crystal Display (LCD).



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Figure 15

Electrical Interface

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POWER SUPPLY The sensors, the probes, the ADMs and the ADIRUs are power supplied as follows: EQUIPMENT

28 VDC

26 VAC

X

X

AOA Sensor

X

X

Pitot Probe

X

ADIRU

Static Probe TAT Sensor ADM

X

115 VDC

X X 13.5 VAC from ADIRU

The ADIRU is normally supplied with 115VAC, 400 Hz power for the ADR and IR functions. However its AOA resolver converter module is supplied with 26VAC, 400 Hz. The 28VDC back-up generation is provided by batteries and is automatically used when the main power exceeds its normal limits. At the beginning of each power cycle the ADIRU switches from the main to the back-up power to test the electrical generation.

ADIRS Power Supply Distribution after the Loss of Main Electrical Generation Loss of the main generation and ATT HDG selector switch in NORM position Captain side :The ADIRU 1 is supplied as in normal configuration. First Officer side :The ADIRU 2 is no more supplied with 115VAC and 26VAC. When the 26VAC is lost, the ADR detects a fault and flags the output parameters. The ADIRU is still powered with 28VDC from the 28VDC HOT BUS 702PP but the Time Delay Opening (TDO) relay 17FP will cut this supply after 5 minutes in emergency configuration. The ADR 2 function is lost immediately. The IR 2 function is lost after 5 minutes. Standby side : The ADIRU 3 is no more supplied with 115VAC and 26VAC. When the 26VAC is lost, the ADR detects a fault and flags the output parameters. The ADIRU is still powered with 28VDC from the 28VDC HOT BUS 701PP but the Time Delay Opening (TDO) relay 14FP will cut this supply after 5 minutes in emergency configuration. The ADR3 function is lost immediately. The IR 3 function is lost after 5 minutes. Loss of the main generation and ATT HDG selector switch in CAPT/3position The CAPT/3 position of the ATT HDG selector switch corresponds to the selection of the ADR 3 in place of the ADR 1. The power supplydistribution must then be modified to keep the ADR 3 in emergency configuration. Captain side : The ADIRU 1 is supplied as in normal configuration. First Officer side : ADIRU 2 supply: Ref. Para. (NORM pos., FO side) Standby side : ADIRU 3 is no more supplied with 115VAC. The ADIRU 3 is still powered with 28VDC from the 28VDC HOT BUS 701PP. The ADR 3 function is lost immediately. The IR 3 function is available.

Ground Warning If one of the 3 ADIRU’s looses 115 VAC power (the ON BAT light on the ADIRS illuminates illuminates), the ’HORN MECH CALL’ sounds with a time delay of 15 seconds, if the A/C is on GROUND. Parallel to this the ADIRU & AVIONICS VENT light in the nose well illuminates.



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Figure 16

ADIRS Power Supply

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ADIRS START PROCEDURE Modes of operation Operation interface with the IR is performed through the MCDU 1(2) or the CDU. The MCDU 1(2) is used for entering initialization data and for displaying IR data. The CDU is used for mode selection, IR annunciation (FAULT, ALIGN), for entering initialization data and displaying IR data. The IR has three selectable modes: OFF, NAV and ATT. The relation between the mode selection and system response is described later in the IR-part .

OFF mode When the OFF/NAV/ATT selector switch on the CDU is in the OFF position, all circuitry in ADIRU is de-energized, except for any logic associated with the power off function. When the ADIRU has turned off, it consumes less than 10 milliamps (needed for power supply turn-on control). The power supply of the ADMs is switched off. A period is required between switching to OFF and actual power off (Honeywell ADIRU : 15 seconds). During this sequence, the last position computed is stored .

MAINTENANCE PRACTICES TASK 34-10-00-860-002 For this procedure the electrical aircraft circuit has to be energized and several Probes and ADIRS Circuit Brakers must be closed first.

Procedure On the ADIRS CDU, set the 3 OFF/NAV/ATT selector switch to NAV. Make shure that the ON BAT light comes on for 5 seconds and the related ALIGN legend comes on. Make shure that the ADR FAULT/OFF legends are off. On the CPT and F/O PFD’s : - Make shure that the CAS, ALT, V/S data are shown. - Make shure that the attitude data is shown 40 seconds after start up. Set the SYS DISPLAY selector switch to 1 and the DATA DISPLAY selector switch to PPOS, and make shure that dashes are shown on the CDU DISPLAY.

NAV mode After selection of the NAV mode on ground, the IR automatically enters the NAV mode, if a self-determined satisfactory alignment has been completed. If alignment is not completed, the IR remains in the Align submode. No updating of the IR present position latitude and longitude is allowed once the IR has completed the Align submode. The IR latitude and longitude entered during alignment is the starting point for its computation. The following logical processes are mechanized: OFF to NAV provides automatic alignment in 10 minutes for latitudes between 73 N and 60 S, with automatic entry to NAV mode. Requires initial position data to be entered. NOTE : The automatic alignment requires 15 mn delay for latitudes between 73 N and 82 N. For high latitudes the alignment (high latitude alignment) is provided by an operational procedure which delays the entering of the initial position by the crew. Accuracies of the system are slightly degraded.



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_ _ _ _ _ _ _ _ _ _ _ _ _

afte r 5 seconds

Figure 17

ADIRS Start Procedure

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IR ALIGNMENT The Inertial Reference part of the ADIRU needs on ground a 10 minute Alignment periode. During this time the aircraft must stay unmoved and the initial Present Position inserted. The Present Position entry may be done via the ADIRS CDU or the MCDU.

Check of the IR Alignment Procedure TASK 34-10-00-710-003 Prior to this procedure several CB’s of the sytems Autflight, Landing Gear, Probes, ADIRS and Hydraulic has to be closed. Do the EIS start procedure. On the overhead panel : - on the FLT CTL panels 23VU and 24VU, make sure that the FAC, ELAC and SEC pushbutton switches are not pushed (in). The OFF legends are on. - on the ADIRS CDU, on the panel 20VU, make shure that the 3 OFF/ NAV/ATT selector switches are at OFF. On the panel 13VU, on the EFIS control section of the FCU, on cpt and F/O side, set the ROSE_NAV mode. NOTE : During the alignment phase until the NAV mode is got, the aircraft must not move. NOTE : On the ADIRS CDU, if the ALIGN legend flashes (maintenance indication : ENTER P POS), make sure that the airport latitude and/or longitude are correct. Enter again the latitude and/or longitude. Do the procedure again until the ALIGN legend stops flashing. NOTE : During the alignment phase necessary to get the NAV mode, the time for alignment is shown : - in the right part of the display window of the ADIRS CDU if the DATA DISPLAY selector switch is at HDG. It is shown in this form: TTN 5 - on the upper ECAM DU, minute after minute (from 6 MN to 1 MN) in this form : IRS IN ALIGN 6 MN.



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Figure 18

Prior Procedure

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IR alignment from the ADIRS CDU SUBTASK 34-10-00-710-056 On the ADIRS CDU : - set the 3 OFF/NAV/ATT switch to NAV - make shure that the OFF legend of the ADR1, ADR2, ADR3 pushbutton switches are off - set the DATA DISPLAY selector switch to PPOS - set the SYS DISPLAY selector to1. Result : - on the ADIRS CDU : * the OFF legend of the ADR pushbutton switches stays off * the ON BAT light stays on for 5 seconds * the ALIGN legends come on and after approximately 30 seconds: - on the CAPT and F/O PFD : * the ATT warning flag are no more shown * the ATT indications are available - on upper ECAM display unit, the MEMO page shows this indication : * IRS IN ALIGN 7 - 10 MN. - on the CDU keyboard, enter the present position (example for HAM) : N 53 37,7’ , then push the ENT key, E 9 59,6’ ,then push the ENT key. NOTE : If you entered incorrect coordinates, push the CLR key and enter new data as above. Result : The ADIRS CDU shows the coordinates N 53 37,7’ (in the left part of the display window) and E 9 59,6’ (in the right part of the display window).

- set the SYS DISPLAY selector switch successively to 2 and 3 to make sure that the coordinates come into view in these 2 positions. Go back to 1. Result : - after atime delay of approximately 5 minutes : * on the CPT and F/O ND , the HDG warning flags are no more shown. The rose is available. - after atime delay of approximately 10 minutes : * on the ADIRS CDU, the 3 ALIGN legends go off * on the upper ECAM display unit, the IRS ALIGN 1 MN indication is no more shown. The IRS ALIGNED indication is shown (the ADIRU is in the NAV mode).



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Figure 19

ADIRS CDU PPOS Entry

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IR alignment from the MCDU SUBTASK 34-10-00-710-057 On the ADIRS CDU : - set the 3 OFF/NAV/ATT selector switches to NAV - make sure that the OFF legend of the ADR1, ADR2, ADR3 pushbutton switches are off Result : - the ON BAT light comes on for 5 seconds - the ALIGN legend comes on - on the upper ECAM display unit, the MEMO page shows this indication : IRS IN ALIGN > 7 MN - on the CPT and F/O PFD : * CAS, V/S and ALT data are shown * after approximately 40 seconds, the ATT warning flags are no more shown and the attitude data comes into view. - set the SYSTEM DISPLAY selector switch to 1 (2 or 3) and the DATA DISPLAY selector to PPOS On the MCDU 1 or 2 : - make sure that the MCDU MENU page is in view and check brightness. - push the line key adjacent to FMGC indication Result : -the A/C STATUS page comes into view. - push the INIT mode key Result : - the INIT page comes into view. - enter either FROM/TO, or a COMPANY ROUTE, or the LAT and the LONG position Result : -the chosen FROM/TO, or COMPANY ROUTE or LAT and LONG is shown on the scratchpad line (lower part of the MCDU).

- push the LSK adjacent to either FROM/TO, or CO RTE, or LAT and LONG indication Result : - the place latitude and longitude are shown below the LAT and LONG indication. The slew promps (arrow up and arrow down) adjacent to the LAT indication are shown - the ALIGN IRS indication is shown on the line above the LONG coordinates. NOTE : The LAT or /and LONG magnitude (regardless of N, S, E or W) can be incremented or decremented as follows For the LAT /LONG change : - on the MCDU keyboard, push one of the two slew keys. Result : -the LAT /LONG coordinates increment or decrement by 1 minute per key press or 1 minute per second if the slew key is pushed (in) and held in this position. On the MCDU, push the line key adjacent to the ALIGN IRS indication. Result : - on the MCDU, the ALIGN IRS indication is no more shown. On the ADIRS CDU, successively set the SYS DISPLAY selector switch to 2 and 3 to make sure that the coordinates are shown for the 3 ADIRU. Result : - the coordinates which are in view on the ADIRS CDU are the same as the coordinates shown on the MCDU - after a time delay of approximately 5 minutes : * on CPT and F/O ND, the HDG warning flags are no more shown. The rose is available. - after a time delay of approximately 10 minutes : * on the ADIRS CDU, the 3 ALIGN legends go off * on the upper ECAM display unit, the IRS IN ALIGN 1 MN indication is no more shown. The IRS ALIGNED indication is shown (the ADIRU is in the NAV mode).



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Figure 20

MCDU PPOS Entry

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ADIRS INTERFACE (AIR DATA) The ADR portion of the ADIRU provides main data sources which are air data references for the aircraft avionic systems. The ADR receives and processes the outputs of the Air Data Modules (ADM) and other sensors. It computes the aerodynamic parameters in the form of digital outputs. The ADR software performes five basic computational elements which are under the air data calculations as follows : pressure altitude functions (ALT / ALT-rate) Mach calculation (M) airspeed calculation (CAS / TAS) temperature calculation (SAT / TAT) output signal processing. Aircraft-dependent calculations are also included in the operational software : static source error correction angel of attack (AOA) maximum operating speed (VMO / MMO). The ADR data outputs are transmitted in two forms : digital by 6 ARINC 429 low speed busses and discrete. The system tests includes continuous in-flight monitoring and manually-activated test modes. The continuous monitoring detects and annunciates faults in the ADR during normal operation. Faults are stored in Non Volatile Memory (NVM) BIT and sent to the Centralized Fault Display System (CFDS) via digital words.



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Figure 21

ADR Architecture

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INPUTS

DIGITAL INPUTS Air Data Module (ADM) inputs Three of five possible inputs are used to receive air mass data from remotely mounted ADM. For the ADIRU 3, only two ADM input busses are used : one for the total pressure data and the other for the everaged static pressure data. Flight Control Unit (FCU) inputs The ADR receives two input buses from the FCU, for digital baro corrections, but uses only one at a time. Centralized Fault Display System (CFDS) input For maintenance puposes, the ADR receives one input bus from the CFDIU. Air Data Reference (ADR) input Each ADR receives two intercommunication buses from the other ADR’s for cross channel comparison purpose.

ANALOG INPUTS Total Air Temperatur (TAT) input The ADR measures the resistance of the sensing element of the TAT sensor. Angle of Attack (AOA) inputs The ADR receives two resolver inputs for angle of attack computation



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Figure 22

Sensor Inputs

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DISCRETE INPUTS The ADR is provided with the following input discretes :

Heat discrete come from the associated Probe Heat Computer (PHC). AOA averadge / uniqui : The ADR portion receives the AOA on two resolvers. This discrete indicates wether the computation must use an average value from the two resolvers , or or the value of resolver 1 in priority, with the second as a back-up in case of failure. This last solution is chosen on the A/C. VMO /MMO dicrete provides the position of the L/G DOWN VMO /MMO SELECTION Switch (22 FP). Static Source Error Correction (SSEC) and AOA correction and selection discretes come from the Slat and Flap Control Computers (SFCC) and are linked to the flap position. A/C identification : 7 discretes provides the ADIRU with the identification of the aircraft. They are used to select the appropriate SSEC and AOA correction laws. ADR OFF indicates to the ADIRU, that the crew has pushed the ADR pushbutton switch on the ADIRS CDU. This commands the ADIRU to stop the transmission of the ADR output busses. Baro-correction source selection provides the ADIRU with the following : - the form (digital or analog) in which the baro-correction transmission is made - the number of sources (2 or 3) - the type of transmission used by the digital sources (single bus or various buses). On the A/C the FCU transmits the CPT and F/O baro-correction in digital form on separate buses.



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Figure 23

Discrete Inputs

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OUTPUTS The ADR data outputs are transmitted in to forms : digital (ARINC 429 LS bus) and discrete.

Digital form Calculated ADR parameter are transmitted on 6 data buses. The parameters on each data bus are coded in different form : BNR : binary data word BCD : binary coded decimal data word ISO : data word coded in ISO5 code DIS : discrete data word HEX: hexadecimal code HYB: mixed code. Discrete form The ADR provides seven standard OPEN /GROUND output dicretes : ADR OFF light ADR FAULT Low Speed Warning Discretes 1, 2, 3 and 4 AOA Special Test. The AOA self test is commanded via the CFDIU interface bus.When the AOA test is active, the AOA sensor is offset to +15 C

IR PROCESSING



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Figure 24

ADR Data Users

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CONTROL The ADIRS CDU provides the control and warning of the three ADRs by means of three ADR illuminated pushbutton switches : the pushbutton switch is used to disable the ADR output buses. It is a momentary action pushbutton switch when the ADR output buses are disabled, the ADR controls the activation of the ADR OFF legend by its output discrete : ADR OFF status when an ADR failure is detected, the ADR controls the activation of the ADR FAULT legend by its output discrete : ADR FAULT each ADR is de-energized when the associated OFF/NAV/ATT selector switch is set to OFF when the associated OFF/NAV/ATT selector switch is set to NAV or ATT, each ADR is switched on independently of the previous selection on the ADR pushbutton switch.



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Figure 25

Control

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AIR DATA INDICATING Altitude (ALT), Computed Airspeed (CAS), Mach number (M) and Vertical Speed (V/S) are computed by the ADIRU (ADR portion), processed by the associated DMC and displayed on the PFDs. True Airspeed (TAS) is supplied in the same way but is displayed on the NDs. In normal configuration, with the AIR DATA selector switch in NORM position, the ADR 1 displays information on CAPT PFD and ND. The ADR 2 displays information on F/O PFD and ND. Static Air Temperature (SAT) and Total Air Temperature (TAT) are also supplied in the same way but are permanently displayed on the lower part of the lower ECAM DU. These items of information are displayed by the ADR 2.

PFD display CAS The CAS indication is displayed in analog form by means of a white tape with graduations every 10 kts and digital values every 20 kts. This tape moves up and down so as to indicate the A/C actual speed value in front of a fixed yellow reference line. The displayed part of the scale represents an 84 kts range. The scale is graduated from 30 kts to 520 kts and the digital values from 40 to 520 kts (item A). In no case can the displayed CAS be lower than 30 kts. In case of computed airspeed failure, the speed scale goes out of view and is replaced by a red SPD flag (item B). Mach When the Mach number is above 0.5, it is displayed just below the speed scale. In case of failure, a red MACH flag is presented.

ALT The baro altitude indication is provided by means of a tape which moves up and down behind a window within which the A/C actual altitude is displayed. The tape of the scale is graduated every 100 ft and digital values are displayed every 500 ft in hundreds. The A/C actual altitude is provided by a counter located at the middle of the scale in which the actual value is displayed in green digits. The hundreds of feet are written in a large size whereas the tens and units are displayed by a drum operating as a classical mechanical altimeter. Small white marks are positioned in front of each number on the tape (item A). If the altitude is negative, a NEG white indication is added at the left of the digital value. The digital value is limited to minus 1500 ft (item B). Different displays are presented depending on the baro setting reference (standard or baro corrected). In case of baro altitude failure, the scale goes out of view and a red ALT flag flashes for a few seconds in the altitude window then remains steady (item C). In case of discrepancy between the altitude given by the CAPT air data source and the altitude given by the F/O air data source, a CHECK ALT amber flag is presented on the right side of the altitude scale (item D).



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Figure 26

PFD Display

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V/S The baro vertical speed is automatically displayed in the right side of the PFD when the inertial vertical speed is not available (item A). It is a degraded mode. The vertical speed scale consists of: - a trapezoidal grey background colored surface - a fixed white scale with 500 ft/mn spaced marks from -2000 ft/mn to +2000 ft/mn - a needle giving in analog form the actual vertical speed value - a number in a moving amber window. This window accompanies the needle (above the needle if V/S > 0, below if V/S < 0). The number gives the V/S value in hundreds of ft/mn. Between -200 ft/mn and +200 ft/mn, both the window and the number disappear. - above +6000 ft/mn (or below -6000 ft/mn), the needle remains stopped where it is. When the vertical speed exceeds +6000 ft/mn or -6000 ft/mn, the digital indication and the analog needle change from green to amber. In addition, those indications change to amber in approach, in the following cases: - V/S less than -2000 ft/mn below 2500 ft RA - V/S less than -1200 ft/mn below 1000 ft RA. In case of a failure warning, the vertical speed scale is removed and replaced by a red V/S flag which flashes for a few seconds then remains steady (item B).

ND display The true airspeed (TAS) is displayed on the ND in ROSE, ARC and PLANmode (item A). The TAS information is displayed by a numerical indication of 3 digits preceded by TAS indication. This information is displayed in the left upper corner of the ND for speed higher than 100 kts. Below this value TAS indication remains visible but is followed by three dashes (item B).



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Figure 27

ND Display

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ECAM SD display The Static Air Temperature (SAT) and the Total Air Temperature (TAT) are permanently displayed on the lower part of the lower ECAM DU by a numerical indication of two digits preceded by the plus or minus sign (item A). These data are delivered by the ADR 1. In case of failure or when NCD information is received from the ADR 1, these data are replaced by crosses (item B).

Reconfiguration display In case of loss of AIR DATA parameters on CAPT or F/O PFD and ND the ADR 3 can be used as a back up source by placing the AIR DATA selector switch in CAPT/3 position for EFIS 1 and F/O/3 position for EFIS 2. In case of loss of TAT/SAT parameters on the lower ECAM DU the ADR 3 can be used as a back up source by placing the AIR DATA selector switch in CAPT/3 position.



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Figure 28

ECAM Display

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AIR DATA WARNINGS In addition to the AIR DATA flags displayed on the PFDs and NDs and on the DU, warning messages are displayed on the lower part of the upper ECAM DU. Two kinds of warning messages can be displayed: failure warning messages in case of loss of AIR DATA parameters configuration warning messages in case of dangerous configuration of the aircraft. When the CLR key is pushed on the ECAM control panel, a STATUS page isdisplayed on the lower ECAM DU and indicates the STATUS and INOP SYS (systems).

Failure Warning messages NAV ADR 1 (2) (3) FAULT NAV ADR 1 (2) + 2 (3) FAULT When these messages are displayed : - the MASTER CAUT light on the glareshield come on - the Single Chime (SC) sounds - the FAULT legend of the ADR push button switch on the CDU comes on.



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Figure 29

NAV ADR

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NAV ALTI DISCREPANCY This message is displayed when a difference higher than plus or minus 250 ft is detected by the external comparison inside the FWCs between the baro-corrected altitude (or plus or minus 500 ft for the standard altitude) provided by two ADRs. When it is displayed: - the MASTER CAUT lights on the glareshield come on - the Single Chime (SC) sounds - the CHECK ALT message appears on the PFD.



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Figure 30

NAV ALTI DISCREPANCY

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Configuration warning messages OVERSPEED VMO/MMO OVERSPEED VFE/VLE When these messages are displayed : - the MASTER WARN lights on the glareshield flash - the Continuous Repetitive Chime (CRC) sounds. OVERSPEED VMO/MMO warning processed by the FWC is triggered when the CAS/Mach calculated by the ADR exceed the VMO/MMO threshold by more than 4 kts/0.006 Mach. OVERSPEED VFE/VLE processed by the FWC is a function of Vc and depends on slat/flap position for the VFE and landing gear position for the VLE. For overspeed VLE, the warning is triggered at 284 knots.

Stall Warning When this wrning is activated : - The MASTER WARN lights on the glare shield flash - the cricket and the voice STALL sound. This warning is processed by the FWC and is a function of angle of attack value and slat position following three conditions: - normal law: if corrected angle-of-attack exceeds 23 or if corrected angle-of-attack exceeds 15 and slat < 15 . - alternate law: if corrected angle-of-attack exceeds 13 or if corrected angle-of-attack esceeds 8 and slat < 15 .



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OVERSPEED

STALL

Figure 31

Configuration Warning

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ADIRS INTERFACE (INERTIAL REFERENCE) The Inertial Reference (IR) portion of the Air Data /Inertial Reference Unit (ADIRU) provides main data sources which are precision attitude, magnetic heading references and navigation data to the aircraft avionics systems. Attitude, heading and navigation data are displayed on the Electronic Flight Instrument System (EFIS) displays (Primary Flight Display (PFD); NavigationDisplay (ND)) and on the VOR/DME RMI which recopies the heading data. On A/C’s with GPS : The IR portion also provides selected GPS data and accurate GPIR hybrid position. The IR portion is a strapdown inertial system which provides a quality reference for attitude, heading (true and magnetic), angular rates and accelerations. The IR software also computes: the inertial position the ground velocities the baro inertial vertical speed the drift angle the wind data the flight path data. The IR processing unit are linked to its own ADR portion of the ADIRU via an internal data bus and additionally it provides two digital data input ports for receiving data from the other ADIRUs in case of an internal ADR-Failure. The incoming data are in ARINC 429 LS format and incude altitude and true airspeed. With this inputs the IR processor is able to calculate Inertial Vertical Speed (IVS) and the WIND-Components (speed and direction).

The IR software operates in one of three basic modes: alignment, navigation, or a reversionary attitude mode. These modes include various portions of the major functions. The real-time executive and built-in test functions interface with each function in each mode. On A/C’s with GPS : The GPIR function computes a hybrid GPS/IRS solution utilizing inputs received from the IR function and GPSSUs. The GPIR function has two operating modes: GPIR NAV mode and GPIR ATT mode, as indicated by IR mode command. In GPIR NAV mode, all the system state transitions are slaved to the IR function (Align, NAV). The system tests includes continuous in-flight monitoring and manually-activated test modes. The continuous monitoring detects and annunciates faults in the IR during normal operation.

The IR data outputs are transmitted in two forms : digital by 3 ARINC 429 high speed busses (4 buses, if GPS is installed) and discrete.



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Figure 32

IR Architecture

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INPUTS Digital Inputs FMGC inputs The IR portion is provided with two ARINC 429 LS buses from the two FMGCs. These buses transmits the following data : - Set Latitude - Set Longitude - Set Magnetic Heading - FMGC Discretes. CDU inputs The IR portion is provided with one ARINC 429 LS bus from the ADIRS CDU. This bus transmits the following data : - Set Latitude - Set Longitude - Set Magnetic Heading - CDU Test. ADR inputs The IR portion is provided with two ARINC 429 LS buses from the two other ADIRUs (ADR portion) and with one bus from its associated ADR. These buses transmit the following data : - Altitude - True Airspeed. CFDS inputs The IR portion is provided with one ARINC 429 LS bus from the CFDS. This bus transmits the following data : - UTC - Flight Phase - A/C Config - CFDS Command - Date - Flight Number - A/C Tail Number

Discrete Inputs The IR portion is provided with the followin discrete inputs :



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Figure 33

IR Inputs

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OUTPUTS The IR data outputs are transmitted in to forms : digital (ARINC 429 HS bus) and discrete. Digital form Calculated IR parameter are transmitted on 3 data buses (4 buses, if GPS is installed). The parameters on each data bus are coded in different form : BNR : binary data word BCD : binary coded decimal data word ISO : data word coded in ISO5 code DIS : discrete data word HEX: hexadecimal code HYB: mixed code ALPHA CODE :indicates the parameter mnemonic code Only for GPS-ADIRUs : When the GPS PRESENT programming pins input discretes are grounded (indicating GPS present), both the GPSSU outputs and the GPIRS integrated navigation solution outputs are transmitted on the IR output buses with the IR output data. Discrete form The IR provides 3 discrete outputs : ON BAT When the IR is powered with batteries, this discrete delivers a 28 VDC state and sets the ON BAT light to on. IR FAULT when a failure is detected by the IR, this discrete delivers a 28 VDC state and sets the FAULT legend to on. IR ALIGN When the IR is aligning, this discrete delivers a ground and sets the ALIGN legend to on.



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IR Data Users

IR Dig Output Characteristics

Figure 34

IR Outputs

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ALIGNMENT Initialization Data IR alignment is carried out on ground before takeoff and after the entry of the current aircraft coordinates on the INIT page of the MCDU 1(2) or on the CDU (DATA DISPLAY selector switch in PPOS position). Valid initial position data must be received and verified by the IR during the 10-minute alignment or automatic sequencing to the NAV mode will be delayed after position data is received. Initial position data are verified by the IR. If a miscompare exists then: a message is displayed on the upper ECAM DU : NAV IR 1(2)(3) NOT ALIGN POSITION MISMATCH PRESENT POS-----INSERT a message ENTER PPOS is displayed on the ADIRS CDU (DATA DISPLAY selector switch in STS position) The miscompare is removed and the position data verified by the IR when: the last two Set Latitudes received by the IR are identical and the last two Set Longitudes received by the IR are identical, or the last Set Latitude and Set Longitude received by the IR compare within one degree of the latitude and longitude from the previous flight.



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Figure 35

Alignment Data initialization

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Latitude Comparison The IR compares Set Latitude with a self-computed gyro-compass latitude after 10.0 minutes into alignment or any subsequent time when a valid Set Latitude is available. In case of discrepancy the following message is displayed on the upper ECAM DU: NAV IR 1(2)(3) NOT ALIGN POSITION MISMATCH -PRESENT POS-----INSERT The message ENTER PPOS is displayed on ADIRS CDU (DATA DISPLAY selector switch in STS position). The discrepancy exists when: the entered latitude differs from the computed latitude by greater than 0.5 The miscompare is removed if a subsequent entry of Set Latitude passes the test. If latitude test fails two times with identical set latitude inputs then: the IR FAULT legend flashes on the CDU the message IR FAULT appears on the CDU liquid crystal display (DATADISPLAY selector switch in STS position) A warning message appears on the upper ECAM DU: NAV IR 1(2)(3) FAULT.

Excessive Motion The IR performs an excessive motion test during the Align submode. If taxiing or towing causes a step input which exceeds 0.2 ft/s, in the X or Y velocity then: the EXCESS MOTION message is displayed on the ADIRS CDU (DATA DISPLAY selector switch in STS position) the following message is displayed on the upper ECAM DU: NAV IR 1(2)(3) NOT ALIGN EXCESS MOTION IR 1(2)(3) IN ALIGN the attitude information is flagged on the PFD. Thirty seconds after motion detection, the system reverts to a full alignment (time to the end of alignment will revert to 9 min 30 s). It is not necessary to re-enter the position.



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Figure 36

Align Procedure

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RAPID REALIGNMENT The IR also offers the possibility to enter into a variant of the alignment mode called ”rapid realign” or ”30-second realign”. Thismode is selected by moving the CDU selector switch from NAV to OFF then to NAV within five seconds, when the aircraft is on ground (ground speed less than 20 knots). Valid position data must be received. During the ”rapid realign” mode all computed velocities are set to zero and a fine tuning of the alignment is performed using theattitude reference vertical and the heading data available from the last NAV phase as initial conditions.



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Time to NAV in Minutes

Pull Switch from NAV Detent before turning to OFF and back to NAV

Figure 37

Rapid Realignment

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ATT MODE The IR has a reversionary mode which can be activated only by manual selection of ATT mode on the CDU. The mode can be activated on the ground or in the air and is intended to provide a rapid attitude/heading restart capability in the event that the IR has experienced a total power shutdown or a failure has occurred resulting in the following : IR FAULT legend flashing on the CDU IR 1(2)(3) FAULT message displayed on the upper ECAM DU: IR x MODE SEL...ATT SELECT ATT message displayed on the CDU (DATA DISPLAY selector switch in STS position). The IR is designed so that the ATT mode can be used after BITE has detected failures which will cause excessive NAV mode data errors but does not disable the ATT mode mechanization. However, it is recommended to stay in NAV mode even with excessive navigation errors because of higher accuracy of attitude signals and a more complete signal processing. ATT mode must always be used after loss of power or a similar situation in the air where a new alignment /leveling is required. The ENTER HEADING message is displayed on the CDU (DATA DISPLAY selector switch in STS position) when ATT mode is selected until valid heading initialization is received from the MCDU or the CDU. The ATT mode is normally engaged with the aircraft in level flight. A 30-second period (Honeywell ADIRU : 20 second) is needed with the aircraft in level flight to perform an attitude erection to initialize a ”level” attitude. During this period, the data normally computed in ATT mode, have SSMs set to NCD. NOTE : In ATT Mode the ADIRU is a ’’Free Azimuth System. That means, the HDG-value drifts and must be updated time after time.



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Figure 38

Attitude Mode

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MODE SELECTION The relation between the mode selection and system respons is shown in the following ’Mode State Diagram’.



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Figure 39

Mode State Diagram

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INERTIAL REFERENCE INDICATING Attitude and heading information is computed by the ADIRU (IR portion) and processed by the associated DMC. The attitude data are displayed on the PFD and the heading data are displayed on the PFD, the ND and the VOR/DME RMI. In addition, vertical speed (V/S) is displayed on the PFD and Ground Speed and wind indications are displayed on the ND. In normal configuration, with the ATT/HDG selector switch in NORM position, the IR1 data are displayed on the CAPT PFD, ND and the VOR/DME RMI. The IR2 data are displayed on the F/O PFD and ND. The following parameters can be displayed on the CDU liquid crystal display according to the position of the DATA DISPLAY selector switch on the CDU: wind (WIND) present position (PPOS) true heading (HDG) status of selected system conditions (STS) track and ground speed (TK/GS) test values (TEST). The sources of the data displayed are controlled by the SYS DISPLAYselector switch on the CDU.

ATTITUDE INFORMATION The aircraft roll and pitch attitude is indicated in the center part of the PFD by a sphere representing a conventional ADI drum.

(1) Fixed Aircraft Symbol (black, yellow boxed) The airplane symbol represents the longitudinal- and lateral axis of the A/C.

(2) Roll Scale (white) This fixed roll scale comprises white marks for the 10 degrees, 20 degrees, 30 degrees and 45 degrees significant values, on either side of the zero position (horizontal wings) which is indicated by a small fixed triangle.

(3) Roll Index (yellow) A yellow triangle which remains on the line going through the center of the A/C reference and which is perpendicular to the horizon line, moves against the fixed roll scale on the upper contour of the attitude sphere.

(4) Pitch Scale (white) The scale moves behind the cut-sphere shaped window, limited by the lines of an upper and a lower sector. The scale rotates around the center of the A/C reference in accordance with the A/C present roll angle. The lines are given every 2.5 degrees from 0 to 30 degrees, then for the 50 degrees and 80 degrees values for positive pitch angles.

(5) Side Slip Index (yellow) Represents on GROUND the A/C-accelloration in the latitude axis and during FLIGHT the side slip.

(6) Heading Reference Line (yellow) and Heading Scale (white) (see next page)

(7) Actual Track Symbol (green) The actual Track Symbol represents the A/C-movement in relation to True North. Pitch Angle Information (item A) The A/C present pitch angle is given by the vertical displacements of the pitch attitude scale with respect to the center of the A/C reference. Beyond 30 degrees, red large arrow heads (V-shaped) indicate an excessive attitude and the direction to follow in order to reduce it (item B). Roll angle information (item A) A yellow triangle which remains on the line going through the center of the A/C reference and which is perpendicular to the horizon line, moves against the fixed roll scale on the upper contour of the attitude sphere. Attitude failure In case of attitude failure concerning the pitch and/or roll information the attitude sphere goes out of view and is replaced by a red ATT flag which flashes for a few seconds then remains steady (item C). In case of discrepancy detected by the FWC between the pitch or roll attitude information presented on the CAPT and F/O PFDs, a CHECK ATT amber message flashes for a few seconds on both PFDs, then remains steady (item D).

Pag0 ______________________________________________________________________________________________________________________________________________________________________________________________ Revision No : 02 Issue Date : 21/05/2013


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(3)

(2)

(5)

(1) (4)

(6)

(7)

Figure 40

ATT Information

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HEADING INFORMATION The aircraft magnetic or true heading is displayed on the PFD, the ND and the VOR/DME RMI. The true heading can be displayed on the CDU.

On the PFD When true heading is displayed, a TRUE white message appears above the heading scale (item A). A blank heading scale (with 10 deg. spaced marks without any indicated value) is provided on the horizon line. The marks are just under this line. This scale moves as the aircraft heading varies. For important nose up or nose down the heading graduations remain at the lower or upper sector limit. Below the sphere, a heading scale provides the pilot with the aircraft actual track and relative selection. This heading scale is graduated every 5 deg. (item A). In case of failure, the heading graduation disappears on the two scales and a red HDG flag appears on the lower heading scale (item B). It flashes for a few seconds then remains steady. Furthermore, in case of discrepancy detected by the FWC between CAPT and F/O heading indications, with the heading signal valid, a CHECK HDG amber message is displayed at the center of the heading scale (item C).

On the ND The heading data is displayed on the ND in the three following operating modes: ROSE, ARC and PLAN. The ROSE mode and the ARC mode are oriented with respect to the aircraft heading, while the PLAN mode is oriented with respect to the true north. True heading display In ROSE or ARC mode, when true heading is displayed, a cyan TRUE message appears at the top of the ND. ROSE mode (item A) In this mode each pilot has 3 different sub-modes of presentation of his ND: ROSE-ILS/ROSE-VOR/ROSE-NA V. In the three ROSE sub-modes, the ND provides a display which is similar to that of a conventional HSI, i.e. a rotating heading dial orientated to the North and giving to the pilot the aircraft actual magnetic or true heading with as reference the fixed yellow lubber line at the top of the dial. ARC mode (item D) In this mode the ND displays a 90 deg. heading sector ahead of the aircraft giving the aircraft actual magnetic or true heading with respect to the fixed yellow lubber line at the top of the scale. PLAN mode The ND displays a static map orientated with respect to the true North. Heading failure (items C and D) In case of heading failure, the scale and all symbols positioned on the ROSE and ARC scales go out of view; a red HDG flag comes into view below the scale after flashing for a few seconds, when the DMC has detected an anomaly concerning the heading parameter. In addition, if a discrepancy between CAPT and F/O sides is detected by the comparison inside the FWCs, the CHECK HDG message is displayed in amber on both NDs (item E).



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PFD

ND

Figure 41

HDG Information

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On the VOR /DME RMI The heading indication is given by a dial which rotates in front of a fixed index (item A). In case of heading failure, the fire orange warning flag with the black HDG inscription comes into view at the top of the compass card (item B). On the CDU The time heading can be displayed on the CDU if the DATA DISPLAY selector switch is placed in the HDG position.

GROUND SPEED The ground speed is displayed in the left upper corner of the ND for ROSE, ARC or PLAN mode (item A). The GS title is displayed in white color and the ground speed value in green. In case of failure or NCD, the ground speed value is replaced by three dashed lines (item B). The ground speed can also be displayed on the CDU if the DATA DISPLAY selector switch is placed in the TK/GS position.

WIND INDICATIONS The wind origin, force and direction is displayed in the left upper corner of the ND, for ROSE, ARC and PLAN mode (item A): the wind origin is displayed in green color in degrees with respect to the true North the wind force is displayed in green color in knots the wind direction, in analog form, is represented by means of a green arrow orientated with respect to the north reference in use. This arrow is displayed only if the wind force is greater than 2 knots. In case of failure or NCD, the digital data are replaced by three dashed lines and the wind direction arrow disappears (item B). The wind indications can also be displayed on the CDU if the DATA DISPLAY selector switch is placed in the WIND position.

VERTICAL SPEED The inertial vertical speed is displayed in the right side of the PFD (item C). The vertical speed scale consists of: a trapezoidal grey background colored surface a fixed white scale with 500 ft/mn spaced marks from -2000 ft/mn to +2000 ft/mn a needle giving in analog form the actual vertical speed value a number in a moving blanking window. This window accompanies the needle (above the needle if V/S > 0, below if V/S < 0). The number gives the V/S value in hundreds of ft/mn. Between -200 ft/mn and +200 ft/mn, both the window and the number disappear. above +6000 ft/mn (or below -6000 ft/mn), the needle remains stopped where it is. When the vertical speed exceeds +6000 ft/mn or -6000 ft/mn, the digital indication and the analog needle change from green to amber. In addition, those indications change to amber in approach, in the following cases: V/S less than -2000 ft/mn below 2500 ft RA. V/S less than -1200 ft/mn below 1000 ft RA. In case of failure, the inertial vertical speed display is automatically replaced by the baro vertical speed display

RECONFIGURATION DISPLAY In case of loss of inertial parameters on the CAPT or F/O PFD and ND, the IR3 can be used as back up source by placing the ATT HDG selector switch in CAPT/3 position for the CAPT PFD and ND or in F/O/3 position for the F/O PFD and ND.



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Figure 42

IR Information

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INERTIAL REFERENCE WARNINGS In addition to the ATT or HDG flags displayed on the PFDs, NDs and on the CDU, warning messages are displayed on the lower part of the upper ECAM DU. When the CLR key is pusshed on the ECAM control panel, a STATUS page is displayed on the lower ECAM DU and indicates the STATUS and INOP SYS (systems).

Failure Warning Messages NAV IR 1 (2) (3) FAULT NAV IR 1 (2) + 2 (3) FAULT When this messages are displayed : - the MASTER CAUTION lights on the glareshield come on - the Single Chime (SC) sounds - the IR FAULT legend flashes on the CDU.



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Figure 43

IR Fault

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NAV ATT DESCREPANCY This message is displayed when a difference higher than 5 deg. is detected by comparison inside the FWCs between the roll angle or the pitch angle provided by two IRs. When it is displayed: - the MASTER CAUT lights on the glareshield come on - the Single Chime (SC) sounds - the CHECK ATT message appears on the PFD. NAV HDG DISCREPANCY This message is displayed when a difference higher than 7 deg. (or 5 deg. in true heading) is detected by comparison inside the FWCs between the heading value provided by two IRs. When it is displayed : - the MASTER CAUT lights on the glareshield come on - the Single Chime (SC) sounds - the CHECK HDG message appears on the PFDs and NDs.



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Figure 44

ATT/HDG Discrepancy

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ADIRU PERFORMANCE CRITERIA History The inertial parameters to be considered to evaluate the level of performance of an inertial system after flight completion are: the radial position error (in NM) the residual ground speed error (in kts). Depending on their magnitude noticed at the end of the flight, the concerned inertial system(s) shall or shall not be removed from the aircraft.

Radial Position error In order to address the statistical term of this requirement with the most relevant approximation, the removal criteria use a limit based on the recording of the radial position error on two consecutive flights. The use of a two-strike method presents the advantage to decrease the removal rate of healthy units that have shown, by chance, or by an inaccurate position entry at alignment, a radial position error beyond the specified criterion.The removal boundaries described on the figure present three different areas: Area 1 - ADIRU OK all the time Area 2 - ADIRU to be checked after second flight Area 3 - ADIRU to be replaced.

Residual Ground Speed error The residual ground speed for each IR is determined at the end of the flight when the aircraft has come to a complete stop. Check of the residual ground speed can be made: - On the CAPT (IR1) and F/O (IR2) Navigation Displays (ND): The residual ground speed of the IR3 can be read on the CAPT ND by setting the ATT HDG selector switch to CAPT/3. - On the ADIRS CDU: set the DATA DISPLAY selector switch to TK/GS set the SYS DISPLAY selector switch to 1, 2, 3 read the respective ground speed in the CDU display. Compare the recorded ground speed values with the following limits: - if the residual ground speed error is 15 kts or greater after each of two consecutive flights, replace the ADIRU - if the residual ground speed error is 21 kts or greater at the end of any one flight, replace the ADIRU.



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Figure 45

ADIRU Performance Criteria

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ADIRS BITE (LITTON ADIRU) The BITE facilitates maintenance on in-service aircraft. It detects and identifies faults related to the Air Data/Inertial Reference System (ADIRS) and reports them to the Centralized Fault Display Interface Unit (CFDIU). The BITE is included in the following LRUs: Air Data/Inertial Reference Unit (ADIRU) Air Data Module (ADM). Control and Display Unit (CDU).

AIR DATA BITE Air Data Module The ADM performs various tests to detect its own faults and failed input signals (check of programming pins). Faults are annunciated to the ADR by omission or labeling of a faulty output word (pressure label) and through the use of a discrete fault-code word output on the ARINC bus. Fault reports are also stored in a non-volatile memory inside the ADM. It is linked to the CFDIU through the ADIRU (ADR and IR portions) which summarizes the BITE results for its own channel.

Air Data Part of ADIRU The ADR BITE monitors certain internal functions, the functionality of other ADR internal hardware, the status of analog, digital and discrete inputs and cross-channel comparisons with the other ADRs. These BITE tests are performed either at power up or continuously with the exception of cross-channel comparisons which are run once at takeoff.

Last Leg Report This item describes the in-flight fault status of the selected ADR and related LRUs during the last leg (flight leg 00).

Previous Legs Reports This item describes the in-flight fault status of the selected ADR and related LRUs during the previous legs (flight legs 01-62).

LRU Ident On page 1/2, the Part Numbers and the Serial Numbers of the ADIRU and the total pressure ADM are displayed. On page 2/2, the associated data to the remaining ADM are displayed.

Trouble Shooting Data A maximum of sixteen 16 bit words can be recorded for the Trouble Shooting Data.

Class 3 Faults

Output Tests The output tests are divided in three parts : Slew tests Interface test AOA test Ground Report

FRA US/E se/rw 2.12.95 ______________________________________________________________________________________________________________________________________________________________________________________________ Revision No : 02 Issue Date : 21/05/2013


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ADIRU 1

ADIRU 3

ADIRU 2

CFDIU

Figure 46

ADR/CFDS Block Diagramm

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Ground Scanning The Ground Scanning function performs most of the continuous tests. All tests with an important time delay (temperature ...) are not performed within this function.

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Figure 47

ADR Ground Scanning

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System test The system test function performs power-up tests and various continuous tests to provide a complete status of the ADR part ofthe ADIRS. It is necessary to re-initialize the system for IR part because the navigation data are erased by the test of the RAM performed at the power-up.

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Figure 48

ADR System Test

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Output Tests 1.: Slew Tests The Slew tests are divided in two parts : Altitude dynamic slew This function slews the altitude between the lower limit and the upper limit (up or down). These data are entered by the operator. The altitude limit values are tested to be within -2000 and +50000 feet. The altitude lower limit is tested to be less than the altitude upper limit. The altitude slew rate is tested to be within 1 to +20000 ft/mn. CAS dynamic slew This function slews the Computed Air Speed between the lower limit and the upper limit (up or down) at the CAS slew rate. These data are entered by the operator. The CAS limit values are tested to be within 0 and +450 knots. The CAS lower limit is tested to be less than the CAS upper limit. The CAS slew rate is tested to be within 1 to +100 Kts/Min. Note: Pull the baro reference selector knob. The Slew Test outputs are only displayed when STD is selected.

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Figure 49

ADR Slew Tests

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2.: Interface Test Initiation of the ADR Interface Test causes the following sequence to occur : 0-5 seconds (Failure Warning Test) - For 0-5 seconds after initiation of the test mode the ADR Outputs are transmitted. Since this is the failure warning test period, the BCD output parameters are not transmitted and the SSM of the BNR parameters are set to FW. Timing tolerance is +/- 0,5 second. 5-10 seconds (Altitude Ramp Test) - For 5-10 seconds after initiation of the test mode the ADR Outputs are transmitted. Since this is the altitude ramp test period, the altitude outputs are slewed in a positive direction for the entire 5-second period at a rate of 600 ft/min, starting at the ambient computed altitude. Timing tolerance is +/- 0,5 second. 10 second until test completion (Fixed Output Test) - From 10 seconds after initialization of the test mode until the test completion is commanded, ADR fixed outputs are transmitted. Timing tolerance is +/- 0,5 second.

Note The required outputs are included in the Task ”BITE Test of the ADR System” ( AMM 34-10-00 ).



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Figure 50

ADR Interface Test

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3.: AOA Test The AOA test function set the AOA test output discrete to command the AOA sensor to fixed position greater than the stall warning threshold. So, the Flight Warning Computer activates the aural stall warning.

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Figure 51

ADR AOA Test

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Current Status The Current Status function displays on 8 pages the state (as read by the computer) of the discrete inputs, digital inputs (ADM and baro-correction), analog inputs and power condition.

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Figure 52

ADR Current Status

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INERTIAL REFERENCE BITE The IR BITE monitors certain functions. Some of them enable to monitor operation errors. These tests are: Align in air Excessive Motion Latitude comparison test. They result in IR warnings but without fault message sent to the CFDIU.

Last Leg Report

Previous Legs Report LRU Ident The Part Numbers and the Serial Numbers of the ADIRU are displayed.

Ground Scanning The Ground Scanning function performs most of the continuous tests. All tests with an important time delay (temperature ...) are not performed within this function.

Trouble Shooting Data A maximum of sixteen 16 bit words can be recorded for the Trouble Shooting Data.

Class 3 Fault System Test The system test function performs power-up tests and various continuous tests to provide a complete status of the IR part of the ADIRS. The layout of the System test pages of the ADR and IR parts are equivalent except of the title ADR which is replaced by IR. It is necessary to re-initialize the system because the navigation data are erased by the test of the RAM performed at the power-up.

Ground Report

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ADIRU 1

ADIRU 3

ADIRU 2

CFDIU

Figure 53

IR/CFDS Block Diagramm

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Interface Test Initiation of the IR Interface Test shall cause the following sequence to occur : 0-2 seconds - BCD, BNR and Discrete Words - output with SSM set to Functional Test and data set. Annunciator discretes shall be energized. Over 2 seconds - BCD, BNR and Discrete Words - output with SSM set to Functionnal Test and data set. Annunciator discretes released to indicated status.

Note The required outputs are included in the Task ”BITE Test of the IR System” ( AMM 34-10-00 ).



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Figure 54

IR Interface Test

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Current Status The Current Status function displays on 8 pages the state (as read by the computer) of the discrete inputs, digital inputs (Set Lat and Long from FMGEC), power condition.



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AMM 34-18-00 Config 2 8/94

Figure 55

IR Current Status

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Specific Data On some aircraft, you don’t have direct access to the Current Status menu. You must select SPECIFIC DATA. Then you can select Current Status GPIRS Report GPIRS Report The GPIRS Report function displays GPS primary (AIME) parameters which are stored either in case of GPS primary failure or routinely at the end of each flight (historic on several flights).



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AMM 34-18-00

Figure 56

GPIRS Report

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ADIRS BITE (HONEYWELL ADIRU) The BITE facilitates maintenance on in-service aircraft. It detects and identifies faults related to the Air Data/Inertial Reference System (ADIRS) and reports them to the Centralized Fault Display Interface Unit (CFDIU). The BITE is included in the following LRUs: Air Data/Inertial Reference Unit (ADIRU) Air Data Module (ADM). Control and Display Unit (CDU).

AIR DATA BITE Air Data Module The ADM performs various tests to detect its own faults and failed input signals (check of programming pins). Faults are annunciated to the ADR by omission or labeling of a faulty output word (pressure label) and through the use of a discrete fault-code word output on the ARINC bus. Fault reports are also stored in a non-volatile memory inside the ADM. It is linked to the CFDIU through the ADIRU (ADR and IR portions) which summarizes the BITE results for its own channel.

Last Leg Report This item describes the in-flight fault status of the selected ADR and related LRUs during the last leg (flight leg 00).

Previous Legs Report This item describes the in-flight fault status of the selected ADR and related LRUs during the previous legs (flight legs 01-62).

LRU Identificaton This item describes, for the ADR portion, the current part numbers and revision status of the selected ADR hardware and software, including ADM identification.

Current Status This item describes the current on-the ground fault status of the selected ADR.

Air Data Part of ADIRU The ADR BITE monitors certain internal functions, the functionality of other ADR internal hardware, the status of analog, digital and discrete inputs and cross-channel comparisons with the other ADRs. These BITE tests are performed either at power up or continuously with the exception of cross-channel comparisons which are run once at takeoff.



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ADIRU 1

ADIRU 3

ADIRU 2

CFDIU

Figure 57

ADR/CFDS Block Diagramm

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ADR INTERACTIVE FUNCTION Interface Test This item controls the functional test modes of the selected ADR. The text which is displayed on the MCDU discripes the various steps of the Interface Test. When you push the LSK adjacent to the TEST START indication, the following test sequence is run: 0-5 seconds: Failure Warning Test On the ADIRS CDU, the FAULT and the OFF legend of the pushbutton switch come on. On the PFD, the speed and altitude scale are no more shown and the SPD and ALT flags are shown. The Master Coution Lights come on with a single chime. 5 to 10 seconds: Altitude Ramp Test The warnings are no more shown. On the PFD, the altitude increases. After 10 seconds: Fixed Output Test Fixed ADR values are shown. The required outputs are listet in the AMM 34-10-00, Task ”BITE Test of the ADR System”.

Altitude Dynamic Slew Test The activation of this test causes the ADR to output a simulated ramp of altitude between low and high altitude limits specified by the operator. These limits, as well as the slew rate, are entered by the operator by means of the MCDU keyboard. The maximum slew rate is 20 000 ft/min. Note: Pull the baro reference selector knob. The Slew Test outputs are only displayed when STD is selected.

CAS Dynamic Slew Test The activation of this test causes the ADR to output a simulated ramp of computed airspeed between low and high speed limits specified by the operator. These limits, as well as the slew rate, are entered by the operator by means of the MCDU keyboard. The maximum slew rate is 100 kts/min.

AOA Sensor Test The activation of this test causes the ADR to output the actual measured values of AOA and to send the test discrete which causes the AOA sensor to be commanded to a fixed position greater than the stall warning threshold. Thus, the Flight Warning Computer activates the aural stall warning.



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Figure 58

ADR Interactive Function

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INERTIAL REFERENCE BITE The IR BITE monitors certain functions. Some of them enable to monitor operation errors. These tests are: Align in air Excessive Motion Latitude comparison test. They result in IR warnings but without fault message sent to the CFDIU.

Last Leg Report This item describes the in-flight fault status of the selected IR during the last leg (flight leg 00).

Previous Legs Report This item describes the in-flight fault status of the selected IR during the previous legs (flight legs 01-62).

LRU Identificaton This item describes the ADIRU part number and serial number. The part number and the serial number of the CDU are also displayed.

Current Status This item describes the current on-the ground fault status of the selected IR.



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ADIRU 1

ADIRU 3

ADIRU 2

CFDIU

Figure 59

IR/CFDS Block Diagramm

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Interface Test This item controls the functional test modes of the selected IR. The Test Mode is inhibited whenever any of the following conditions exists: Ground Speed greater than 20 kts IR in the ATT Mode. The text which is displayed on the MCDU discripes the various steps of the Interface Test. When you push the LSK adjacent to the TEST START indication, the following test sequence is run: 0-2 seconds: On the ADIRS CDU, the ON BAT light, the FAULT light and the ALIGN lights come on. The ENT and CLR keys come on. After 2 seconds: Fixed IR values are shown. The required outputs are listet in the AMM 34-10-00, Task ”BITE Test of the IR System”.



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Figure 60

IR Interactive Function

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34-58

SATELLITE NAVIGATION

GLOBAL POSITIONING SYSTEM GENERAL GPS is a satellite supported navigation system to determine the Present Position. It supplements the ADIRS, because inertial refence systems lose accuracy depending on the time in the navigation mode. 24 satellites are in an orbit and send signals using the frequency 1575.42 Mhz. In order to determinate the present position, GPS uses up to eight satellites.

Indications In the cockpit there are no GPS indications on the instruments.

BITE There is no specific CMS BITE for the GPS.

GPS INTERFACE In the A 321, two GPS sensor units ( GPSSU ) with one antenna each are installed. For a quick alignment the GPSSUs receive the present position from the ADIRUs. The UTC is also used. Each GPSSU supplies the ADIRUs with those data: Present Position Ground Speed Track.

Failures The GPSSU status and GPSSU data are monitored by the ADIRUs. The ADIRUs send the GPS fault message to the FWCs which generate an ECAM message. If GPS 1 fails the following level 2 message is displayed: NAV GPS1 FAULT.

Normally GPSSU 1 supplies ADIRU 1 and ADIRU 3. GPSSU 2 supplies ADIRU 2. If a GPSSU fails the ADIRUs select the remaining GPSSU automatically. Each ADIRU computes hybrid GPIR data, using the data computed by its IR-part and the input from the supplying GPSSU. The IR-part of each ADIRU supplies the FMGCs with those data: Hybrid GPIR data IR data GPS data. In the last case, the signals are routed through the ADIRU.



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ADIRU 1

Att/Hdg Sel Switch

GPSSU 1

FMGC 1 FWC 1

ADIRU 3

GPSSU 2

FWC 2

ADIRU 2

FMGC 2

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GPSSU MODES Initialisation Mode At power on, a self test is initiated within the GPSSU. If an internal failure is detected, the GPSSU sends ”invalid”. To reduce initialization time, the GPSSUs receive position data from the ADIRUs and UTC / Date from the FMGEC through the ADIRUs.

Acquisition Mode In the acquisition mode the GPSSU tries to receive satellite signals. When the selectes satellites have been acquired, the GPSSU transfers into the navigation mode.

Navigation Mode In the navigation mode the GPSSU sends data to the ADIRUs. For computation four to eight satellites are used at a time.



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Figure 62

GPSSU Modes

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GPS MONITOR PAGE General On the MCDU, the GPS monitor page displays GPS data.

Access Procedure The procedure to call the GPS monitor page is: 1. Push the key for DATA on the MCDU. The DATA INDEX appears. 2. Push LSK 3L for GPS MONITOR. The GPS monitor page appears.

GPS Data On the GPS monitor page the following data are displayed: Present Position True Track ( Track refered to geographc north) Figure of Merit ( Accuracy in meters ) Ground Speed Mode ( Acquisition Mode ACQ or nav mode NAV ).

Note: During poor receiving conditions ( aircraft in the hangar ) no data are displayed.



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Figure 63

GPS Monitor Page

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MANUAL GPS DESELECTION When the FMGCs are supplied with hybrid GPIR data, they use this data for navigation and no DME/DME or VOR/DME radio updating is made. This FMGC mode is called GPS / inertial mode The procedure to deselect the GPS / inertial mode manually is: 1. Push the key for DATA on the MCDU. The DATA INDEX appears. 2. Push the LSK 5R for POSITION MONITOR. The POSITION MONITOR page appears. 3. Push the LSK 6R for SEL NAVAIDS. The SELECTED NAVAIDS page appears. 4. Push the LSK 5L for DESELECT GPS. GPS will not be used by the FMGCs any longer. Instead of DESELECT GPS now SELECT GPS is offered. The procedure to re-select GPS is the same. Note: The procedure to select the position monitor page is shown in the AMM 22-72-00.



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AMM 22-72-00

Figure 64

Manual GPS Deselection

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LOCATION Antenna Location Both GPS antennas are mounted in 12 o’clock position.

GPSSU Location The GPSSUs are mounted close to the GPS antennas behind the ceiling in 1 o’clock position.



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Figure 65

Component Location

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AMM 34-58-00

Figure 66

AMM 34-58-11

GPS Antenna

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GPSSU

AMM 34-58-00

AMM 34-58-31

Figure 67

GPSSU

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Training Manual A 319/320/321 ATA 34 Navigation Radio Navigation

Level 3

Lufthansa Issue: August 2001 Technical Training GmbH For Training Purposes Only Book No: A320 34-36 L3 Lufthansa Base Lufthansa 1995 ______________________________________________________________________________________________________________________________________________________________________________________________



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ATA 34

NAVIGATION Radio Navigation

______________________________________________________________________________________________________________________________________________________________________________________________

Revision No : 02US/E MK FRA Issue Date : 21/05/2013

11.7.96


Page: 1

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34-36

Antenna The GlideSlope (G/S) and Localizer (LOC) antennas are common to both receivers. Each antenna has two independent connectors, used to feed the two ILS receivers.

ILS-SYSTEM

DESCRIPTION General The Instrument Landing system allows the aircraft to follow an optimum descent axis in order to perform safe landing with poor visibility conditions. The A320 uses two independent ILS systems. The localizer operates in a frequency band which ranges from 108 MHz to 111.95 MHz. The glide operates in a frequency band which ranges from 328.6 MHz to 335.4 MHz. Tuning Auto Tuning In normal operation the ILS receiver 1 (2) is automatically tuned by the onside FMGC 1 (2) through the associated RMP 1 (2). In this case, the RMP is only used to transmit the frequency and course information from the FMGCs to the frequency input port A of the receiver. Manual Tuning Frequency and course data can by manually entered on the RAD/NAV page of the MCDUs. The FMGCs sent this information to the receivers in the same way like the auto-tuning mode. FM Switching If a FMGC fails, a discrete is sent to the receiver (via the RMP) to activate the frequency input port B. This port receives information direct from the opposite FMGC. In this case, one FMGC tunes both ILS receivers. NAV Back Up Tuning If both FMGC fail, each ILS receiver must be tuned directly from the onside RMP. To do so, press the NAV and the ILS pushbuttons on both RMPs. The RMP now uses manually entered data and not the data coming from the FMGC. A discrete selects the frequency input port A, which is directly supplied from the associated RMP. A second discrete inhibits the data display on the RAD/NAV Page of the MCDUs to indicate that no FMGC tuning is possible. To avoid different ILS data on the two receivers, the ILS data is exchanged between the RMPs if both RMP are in NAV back up mode.

Inputs The ILS TUNE/TEST INHIBIT discrete from the associated FMGC is used to lock the last used frequency and to inhibit the test of the receiver during approach phase below 700 ft. Each LGCIU sends discrete signals to the ILS receiver for internal BITE purposes. Indication All DMCs receive ILS data from both receivers such as LOC/GS deviation, ILS frequency, course (Runway Heading) and ILS identifier. ILS 1 data is shown on Capts PFD and F/Os ND, ILS 2 data on F/Os PFD and Capts ND. Audio The ILS audio signal is processed by the receiver and sent to the AMU and can be heard by the crew on headphones or cockpit loudspeaker. Users The FWCs receive ILS data in order to create ILS warnings on the ECAM in case of ILS failure and to create the ILS deviation warning. The GPWC uses die GS deviation of ILS 1 to create the Below GS-Warning (Mode 5). The FMGCs get ILS data for navigation purpose during various flight phases. The CFDIU is used to communicate with the internal BITE functions of the ILS receivers (tests only available on ground). Warnings and Flags A faulty ILS system results in the following cockpit effects: Flags on PFD and ND Master Caution Lights on the glareshield Aural Warning (Single Chime) NAV ILS 1 (2,1+2) FAULT on the upper ECAM display.

FRA US/E MK 20.12.95 Page: 2 ______________________________________________________________________________________________________________________________________________________________________________________________ Revision No : 02 Issue Date : 21/05/2013


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ILS 1 DATA

ILS 2 DATA

ILS 1 DATA

Figure 1

ILS 2 DATA

ILS System Schematic

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INDICATION Normal Indication on PFD If the ILS pushbutton is pressed and a ILS frequency is sent to the receiver (flight plane insertion on auto tuning, MCDU or RMP insertion on manual tuning), the white ILS deviation scales appear. The magenta deviation indexes appear, when the Localizer or Glide Slope signals are valid. When the deviation is out of range, the index is against one stop and only its outer half remains in view. The scale and the index flash, when the deviation is excessive (ILS deviation warning). The magenta course cursor or dagger shows the ILS course against the heading scale. When the course is out of range, the numeric value is shown on the left or right corner of the heading scale. The magenta ILS Information shows: ILS identifier, if decoded by the ILS receiver. ILS frequency ILS DME distance, if there is a ILS/DME.

Flags or NCD Indication on PFD If the ILS pushbutton is pressed and the ILS receiver fails (LOC or GS) a red ILS message is displayed instead of ILS information in the left bottom corner. Frequency and identifier disappear. With LOC failure, a red LOC flag (flashing 9s, then steady) comes into view in the middle of the LOC scale and the LOC deviation bar goes out of view. With LOC data not available (NCD), the LOC deviation index goes out of view. With G/S failure, a red G/S flag (flashing 9s, then steady) comes into view in the middle of the G/S scale and the G/S deviation bar goes out of view. With G/S data not available (NCD), the G/S deviation index goes out of view. If course input is not available (fail or NCD), the course cursor disappears. The last use frequency will be locked, if the frequency information becomes NCD or fail.

FRA US/E Mk 20.12.95 Page: 4 ______________________________________________________________________________________________________________________________________________________________________________________________ Revision No : 02 Issue Date : 21/05/2013


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G/S Scale G/S Dev. Index

G/S Flag

LOC Scale LOC Dev. Index

ILS Characteristics

Runway Heading

ILS Flag

Figure 2

LOC Flag

ILS Indication on PFD

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Normal Indication on ND ND in Rose ILS Mode The white ILS deviation scales appear. The magenta course cursor or dagger shows the ILS course (runway heading) against the heading scale. The magenta LOC deviation bar appears, when the Localizer signal is valid. It moves perpendicular to the course cursor. When the deviation is out of range, the bar moves against one stop. The scale and the bar flash, when the deviation is excessive (ILS deviation warning). The magenta G/S deviation index appear, when the glide slope signal is valid. When the deviation is out of range, the index moves against one stop and only its outer half remains in view.. The scale and the index flash, when the deviation is excessive (ILS deviation warning). The magenta ILS information shows: - ILS system and frequency - ILS course - ILS identifier, if decoded by the ILS receiver. ND in Rose NAV or ARC Mode If the ILS pushbutton on the EFIS control panel is pressed, the magenta course cursor or dagger shows the ILS course (runway heading) against the heading scale.

Flags and NCD Indication on ND If the ILS receiver faILS (LOC or GS) a red ILS message is displayed instead of ILS Information in the right top corner. Frequency and identifier disappear. With LOC failure, a red LOC flag (flashing 9s, then steady) comes into view in the middle of the LOC scale and the LOC deviation bar goes out of view. With LOC data not available (NCD), the LOC deviation bar goes out of view. With G/S failure, a red G/S flag (flashing 9s, then steady) comes into view in the middle of the G/S scale and the G/S deviation bar goes out of view. With G/S data not available (NCD), the G/S deviation index goes out of view. If the course input fails, a vertical red dagger and a red course flags (CRS XXX) is shown. If the course information is NCD, a course of 0is displayed and the LOC deviation bar goes out of view. The last use frequency will be locked, if the frequency information becomes NCD or fail.



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Dagger

LOC Scale and Dev.-Bar

ILS Information

Dagger

ILS Flag

LOC Flag

G/S Scale and Dev.-Index

Figure 3

Course Flag

G/S Flag

ILS Indication on ND

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MANUAL TUNING OF ILS, VOR/DME, ADF MCDU-T uning A frequency selection is done at the RAD/NAV page via the Alpha-Numeric Keys on the MCDU. To clear a selection, press the CLEAR key (CLR is written in the scratchpad) and press the LSK associated to the system data you want to clear. ILS-TUNING On the RAD/NAV page it is possible to enter a ILS identifier or a frequency and a course. The course will be automatically cleared, if a new ILS station is entered. - After insertion of a new identifier, the FMGEC uses the NAV DATA BASE to search for the new frequency and sent it to the receiver. If the identifier is not in DATA BASE, the NEW NAV AID page is shown. - After insertion of a new frequency, the FMGEC uses the NAV DATA BASE to search for the new identifier to display the data on the MCDU screen. If the identifier is not in DATA BASE, the message NOT IN DATA BASE is shown in the scratchpad and the identifier field is empty. If a flight plan is entered, the system compares this frequency : - on preflight: with the ILS frequency of the Origin-Airport - after preflight: with the ILS frequency of the Destination-Airport If there is no difference, the identifier and the frequency are displayed in cyan (identifier in small fonts, frequency in large fonts). If there is a difference and the frequency is found in the DATA BASE, the frequency is displayed in cyan and the message RWY/ILS MISMATCH is shown in the scratchpad. VOR-TUNING On the RAD/NAV page it is possible to enter a Station identifier or a frequency and a course. The course will be automatically cleared, if a new station is entered.

ADF-TUNING same as VOR-Tuning. After selection of a new ADF frequency, in the lower left or right corner (LSK 6L or 6R ) a ADF 1 or 2 BFO prompt appears. Press the LSK to activate this function. To cancel this function, enter a new frequency or use the CLEAR function key on the MCDU. Manual tuned station will be displayed: on MCDU screen in large fonts on PFD with mode of tuning field shows M. RMP Tuning (Radio Navigation Back Up Mode) The RMPs can be used to tune the radio navigation systems: - RMP 1 for VOR 1, DME 1, ILS 1 and ADF 1 - RMP 2 for VOR 2, DME 2, ILS 2 and ADF 2 To do so, the guarded NAV pushbutton must be pressed to switch the RMP in the radio navigation back up mode (green NAV LED on). All navigation systems associated to that RMP now uses the last stored RMP NAV frequencies. After selection of the NAV system via the pushbuttons, a new frequency and a new course can be entered by using rotary knob and the transfer switch. If both RMPs are in navigation back up mode, the ILS frequency and course are exchanged between both RMPs to avoid different ILS settings to the ILS receivers.



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Figure 4

ILS-, VOR/DME-, ADF- Manuell Tuning by MCDU, RMP

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FAULT ISOLATION AND BITE

ECAM WARNING In case of ILS 1 (2) failure, the ILS waring message ”NAV ILS 1 (2) FAULT” is shown on the upper ECAM display, the MASTER CAUTION comes on and the single chime sounds.

The different BITE menu selections are: LAST LEG REPORT PREVIOUS LEGS REPORT LRU IDENTIFICATION TEST Faults detected by the System and transfered to the CFDS causes the following messages displayed on the MCDU during BITE. ILS 1(2) : NO DATA FROM CONTROL SOURCE There is no correct frequency data input on the active input port of the ILS receiver. RECEIVER The ILS receiver is faulty. NO DATA FROM CFDIU No connection to the CFDS.



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Capt PFD

Capt ND

ECAM

F/O ND

F/O PFD

Warning

EFIS System LOC Antenna

other Systems FMGC 1,2 FWC 1,2 GPWS AMU (Audio)

ÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂ

GS Antenna

ILS 1

ÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂ ILS 2

CFDS

RMP 1

other Systems FMGC 1,2 FWC 1,2 AMU (Audio)

RMP 2

FMGC 1

FMGC 2

ÂÂÂ ÂÂÂ

CFDS monitored

Figure 5

ILS BITE Schematic

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1

1

Figure 6

ILS CFDS BITE Menu

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Figure 7

ILS CDFS BITE Test

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BITE TEST INDICATION



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1

GS scale and index (white and magenta)

2 LOC scale and index (white and magenta) 3

ILS data (magenta)

4

LOC flag (red)

5

G/S flag (red)

6

ILS flag (red)

1 5

6

4

3

2

6

3 1

4

5

2

EFIS mode selector ILS pushbutton

Figure 8

ILS BITE Indication on PFD and ND

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ACTIVATION OF THE FRONT PANEL TEST The front panel test can be activated by pushing the TEST pushbutton switch on the face of the receiver, if the test inhibit discrete is not active. During the first 3 seconds, all LEDs on the face of the receiver are on. During the next 3 seconds, all LEDs go off. During the last 6 seconds (or until the TEST pushbutton switch is released) the green ILS LED is on (except if a fault has been detected during the test). The name, color and function of the three LEDs are as follows: ILS (red) indicates that an internal fault is detected ILS (green) indicates that no internal fault is detected DATA IN (red) indicates that no control input is available (frequency). However, FMGC sends a discrete to inhibit this test during approach phase below 700 feet.



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ILS

green red

DATA IN

red

TEST

ILS RECEIVER

Figure 9

ILS Front Panel Test

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LOCATION

lLS Receiver

GLIDE/SLOPE ANTENNA

LOCALIZER ANTENNA


Figure 10A319/320/321 ILS Location Receiver and Antenna / TRAINING MANUAL

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Figure 11

ILS Location Control and Indication

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34-36

ILS (MULTI MODE RECEIVER)

DESCRIPTION General The primary function of the Multi-Mode Receiver (MMR) is to receive and process Instrument Landing System (ILS) and Global Positioning System (GPS) signals. The A/C comprises two independent MMRs, linked to: a common glide/slope antenna a GPS active antenna, linked to MMR1 a GPS active antenna, linked to MMR2. The MMR is a navigation sensor with two internal receivers: ILS Receiver The function of the ILS is to provide the crew and airborne system users with lateral (LOC) and vertical (G/S) deviation signals, with respect to the approach ILS radio beam transmitted by a ground station. The localizer operates in a frequency band which ranges from 108.1 MHz to 111.95 MHz and the glide uses the band from 329.15 MHz to 335 MHz. GPS Receiver The GPS is a radio aid to worldwide navigation which provides: - the crew with a readout of accurate navigation information, e.g. position, track and speed. - the Flight Management and Guidance Computer (FMGC) with position information, after hybridization in the Air Data/Inertial Reference Unit (ADIRU) with inertial parameters, for accurate position fixing. ILS Operation The equipment given below can control the ILS operation: the Multipurpose Control and Display Units (MCDU) and the Flight Management and Guidance Computers (FMGC) for frequency/ course selection in normal operating mode. the Radio Management Panels (RMPs) for frequency/course selection in back-up mode.

The ILS data are shown on the EFIS displays: the CAPT PFD and F/O ND show the deviations from the ILS1. the F/O PFD and CAPT ND show the deviations from the ILS2. The Morse-coded audio identification signals are sent to the Audio Management Unit (AMU). GPS Operation In normal operation, the GPS 1 data are used by the ADIRUs 1 and 3; the GPS 2 data by the ADIRU 2. NOTE:

IN ORDER TO REDUCE GPS INITIALIZATION TIME, THE GPS 1(2) RECEIVES DATA FROM THE ADIRU 1(2). The IR portion of the ADIRU 1(2) provides the FMGC 1(2) with: pure IR data pure GPS data (in this case the ADIRU operates as a relay) The pure GPS data are used for display on the MCDU 1 and 2. hybrid GPIR data. The hybrid GPIR 1(2) data are used by the FMGC 1(2) for position fixing purposes. In case of one GPS failure, the three ADIRUs automatically select the only operative GPS to compute hybrid GPIR data. In case of ADIRU 1 failure, the FMGC 1 uses ADIRU 3 / GPS 1 data. In case of ADIRU 2 failure, the FMGC 2 uses ADIRU 3 / GPS 2 data. NOTE:

THE PRIMARY SOURCE OF THE ADIRU 3 BEING THE GPS 1, IT IS NECESSARY TO SELECT THE SECONDARY INPUT PORT OF THE ADIRU 3 (GPS 2) BY MEANS OF THE ATT HDG SELECTOR SWITCH TO PRESERVE SIDE 1 / SIDE 2 SEGREGATION (GPS 1 / ADIRU 1 / FMGC 1 AND GPS 2 / ADIRU 3 / FMGC 2 ARCHITECTURE). In case of failure of two ADIRUs, the two FMGCs use only the operative ADIRU. This ADIRU receives data from its own side GPS (e.g. ADIRU 1 - GPS 1).

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Figure 12

MMR Data Acquisition

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Maintenance Operation The MMR system provides the Centralized Fault Display Interface Unit (CFDIU) with an interface for onboard testing and fault reporting purposes. The MCDUs show the maintenance data. Power Supply Each system is energized through 115VAC busbars as follows: 401XP for system 1 204XP for system 2. The system is supplied through these circuit breakers: 49VU COM NAV/MMR/1 COM NAV/MMR/2 121VU

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Figure 13

MMR Monitoring and Display

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ILS FUNCTION Normal operation Each MMR is connected to one Radio Management Panel (RMP). The MMR 1 is connected to the RMP 1 (the MMR 2 to the RMP 2). The MMR 1 receives management bus from the FMGC 1 through the RMP 1 (the MMR 2 from the FMGC 2 through the RMP 2). In normal operation, the FMGC 1(2) tunes the MMR 1(2) either automatically or manually by means of the MCDU. In this case the RMP 1(2) operates as a relay which sends the frequency information from the FMGC 1(2) to the receiver 1(2). Via a second port, the MMR 1(2) receives a second management bus (ILS FREQ + RWY HDG) directly from the FMGC 2(1). The receiver selects one of the two input ports according to the FREQ / FUNCT DATA SOURCE SEL discrete signal, which is received from the FMGC 1(2) through the RMP 1(2). Operation in case of failure With failure of one FMGC, the second FMGC, can control the two MMRs, the off side directly, the on side through its RMP. With failure of the RMP 1(2) or two RMPs, the RMP concerned is transparent to data and discrete from FMGC. Manual operation In manual operation (at any time, or with failure of two FMGCs) the RMP 1 can control the MMR 1 after ON NAV mode selection. Same possibility for the RMP 2 (MMR 2). In this mode the RMP 1 can control the MMR 2 through the RMP 2 after ON NAV mode selection on the RMP 2. Same possibility for RMP 2 through RMP 1. After any frequency selection it is always necessary to select the associated course.

Reconfiguration switching In normal utilization, the ILS 1 data are shown on the CAPT PFD and the F/O ND; the ILS 2 data on the F/O PFD and the CAPT ND. The DMC 1 supplies data to the CAPT PFD and ND; the DMC 2 to the F/O PFD and ND. With failure of the DMC 1(2) it is possible to switch over to the DMC 3 with the EIS DMC selector switch located on the center pedestal. In this case, the DMC 3 totally replaces the DMC 1(2) through the stage of the output switching relay of the failed DMC. With failure of the PFD, there is an automatic transfer of the PFD image onto the ND. With failure of the CAPT (F/O) ND, you obtain the transfer of the ND image onto the CAPT (F/O) PFD when you push the PFD/ND XFR pushbutton switch. When you set the PFD potentiometer to OFF this causes: deactivation of the CAPT (F/O) PFD transfer of the PFD image onto the CAPT (F/O) ND. Audio control The MMR applies its audio output to the audio integrating system. This system controls and directs the output to the headsets and / or the loudspeakers. The Audio Management Unit (AMU) controls the audio level through the ACP. On the ACP, the pilot must push the ILS pushbutton switch and adjust the related potentiometer to the correct audio level. With ILS / DME collocated stations, the DME identification morse code can be listened in sequence with the ILS audio signal when you push the ILS pushbutton switch on the ACP and the ILS pushbutton switch on the FCU.

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Figure 14

MMR ILS Operation

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MANUAL TUNING OF ILS MCDU-T uning A frequency selection is done at the RAD/NAV page via the Alpha-Numeric Keys on the MCDU. To clear a selection, press the CLEAR key (CLR is written in the scratchpad) and press the LSK associated to the system data you want to clear. ILS-TUNING On the RAD/NAV page it is possible to enter a ILS identifier or a frequency and a course. The course will be automatically cleared, if a new ILS station is entered. - After insertion of a new identifier, the FMGEC uses the NAV DATA BASE to search for the new frequency and sent it to the receiver. If the identifier is not in DATA BASE, the NEW NAV AID page is shown. - After insertion of a new frequency, the FMGEC uses the NAV DATA BASE to search for the new identifier to display the data on the MCDU screen. If the identifier is not in DATA BASE, the message NOT IN DATA BASE is shown in the scratchpad and the identifier field is empty. If a flight plan is entered, the system compares this frequency : - on preflight: with the ILS frequency of the Origin-Airport - after preflight: with the ILS frequency of the Destination-Airport If there is no difference, the identifier and the frequency are displayed in cyan (identifier in small fonts, frequency in large fonts). If there is a difference and the frequency is found in the DATA BASE, the frequency is displayed in cyan and the message RWY/ILS MISMATCH is shown in the scratchpad.

RMP Tuning (Radio Navigation Back Up Mode) The RMPs can be used to tune the radio navigation systems: - RMP 1 for VOR 1, DME 1, ILS 1 and ADF 1 - RMP 2 for VOR 2, DME 2, ILS 2 and ADF 2 To do so, the guarded NAV pushbutton must be pressed to switch the RMP in the radio navigation back up mode (green NAV LED on). All navigation systems associated to that RMP now uses the last stored RMP NAV frequencies. After selection of the NAV system via the pushbuttons, a new frequency and a new course can be entered by using rotary knob and the transfer switch. If both RMPs are in navigation back up mode, the ILS frequency and course are exchanged between both RMPs to avoid different ILS settings to the ILS receivers.

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Figure 15

ILS Manual Tuning with MCDU, RMP

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ILS INDICATION Normal Indication on PFD If the ILS pushbutton is pressed and a ILS frequency is sent to the receiver (flight plane insertion on auto tuning, MCDU or RMP insertion on manual tuning), the white ILS deviation scales appear. The magenta deviation indexes appear, when the Localizer or Glide Slope signals are valid. When the deviation is out of range, the index is against one stop and only its outer half remains in view. The scale and the index flash, when the deviation is excessive (ILS deviation warning). The magenta course cursor or dagger shows the ILS course against the heading scale. When the course is out of range, the numeric value is shown on the left or right corner of the heading scale. The magenta ILS Information shows: ILS identifier, if decoded by the ILS receiver. ILS frequency ILS DME distance, if there is a ILS/DME.

Flags or NCD Indication on PFD If the ILS pushbutton is pressed and the ILS receiver fails (LOC or GS) a red ILS message is displayed instead of ILS information in the left bottom corner. Frequency and identifier disappear. With LOC failure, a red LOC flag (flashing 9s, then steady) comes into view in the middle of the LOC scale and the LOC deviation bar goes out of view. With LOC data not available (NCD), the LOC deviation index goes out of view. With G/S failure, a red G/S flag (flashing 9s, then steady) comes into view in the middle of the G/S scale and the G/S deviation bar goes out of view. With G/S data not available (NCD), the G/S deviation index goes out of view. If course input is not available (fail or NCD), the course cursor disappears. The last use frequency will be locked, if the frequency information becomes NCD or fail.

FRA US/T WB 01.08.2000 ______________________________________________________________________________________________________________________________________________________________________________________________ Revision No : 02 Issue Date : 21/05/2013


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G/S Scale G/S Dev. Index

G/S Flag

LOC Scale LOC Dev. Index

ILS Characteristics

Runway Heading

ILS Flag

Figure 16

LOC Flag

ILS Indication on PFD

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Normal Indication on ND ND in Rose ILS Mode The white ILS deviation scales appear. The magenta course cursor or dagger shows the ILS course (runway heading) against the heading scale. The magenta LOC deviation bar appears, when the Localizer signal is valid. It moves perpendicular to the course cursor. When the deviation is out of range, the bar moves against one stop. The scale and the bar flash, when the deviation is excessive (ILS deviation warning). The magenta G/S deviation index appear, when the glide slope signal is valid. When the deviation is out of range, the index moves against one stop and only its outer half remains in view. The scale and the index flash, when the deviation is excessive (ILS deviation warning). The magenta ILS information shows: - ILS system and frequency - ILS course - ILS identifier, if decoded by the ILS receiver. ND in Rose NAV or ARC Mode If the ILS pushbutton on the EFIS control panel is pressed, the magenta course cursor or dagger shows the ILS course (runway heading) against the heading scale.

Flags and NCD Indication on ND If the ILS receiver faILS (LOC or GS) a red ILS message is displayed instead of ILS Information in the right top corner. Frequency and identifier disappear. With LOC failure, a red LOC flag (flashing 9s, then steady) comes into view in the middle of the LOC scale and the LOC deviation bar goes out of view. With LOC data not available (NCD), the LOC deviation bar goes out of view. With G/S failure, a red G/S flag (flashing 9s, then steady) comes into view in the middle of the G/S scale and the G/S deviation bar goes out of view. With G/S data not available (NCD), the G/S deviation index goes out of view. If the course input fails, a vertical red dagger and a red course flags (CRS XXX) is shown. If the course information is NCD, a course of 0is displayed and the LOC deviation bar goes out of view. The last use frequency will be locked, if the frequency information becomes NCD or fail.

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Dagger

LOC Scale and Dev.-Bar

ILS Information

Dagger

ILS Flag

LOC Flag

G/S Scale and Dev.-Index

Figure 17

Course Flag

G/S Flag

ILS Indication on ND

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GPS FUNCTION Normal operation To reduce initialization time, the MMR 1(2) receives position data, LAT/LONG from the ADIRU 1(2) and SET LAT, SET LONG UTC/Date from the FMGC 1(2) through the ADIRU 1(2). Each MMR receives the GPS satellite RF signals from the active antenna to compute and provide the three ADIRUs with: UTC, date position, altitude ground speed, track angle N/S speed, E/W speed, vertical speed horizontal and vertical dilution of precision, figure of merit satellite position satellite measurement (pseudo-range, delta range, range rate, UTC measurement time) GPS measurement status, sensor status real time and predictive integrity data. Within each ADIRU an hybridization function performs the following: monitoring of the MMR using GPS status word and ADIRU BITE generation of failure message for ECAM display use of pseudo-range/delta range data to compute GPS position use of inertial data to smooth GPS position/velocity use of a Kalman filter to estimate and minimize errors use of IR data to improve the robustness of the MMR RAIM algorithm. transmission of GPS and GPIR data to the FMGC for position fixing and display purposes. GPS primary navigation function principle in the FMGC A navigation mode with the least error is chosen based upon the mixed IR position and the best GPIR or radio position available. NOTE:

THE GPIR POSITION USED BY THE FMGC TO DETERMINE THE AIRCRAFT POSITION IS COMPUTED IN THE GPIR PARTITION OF THE ADIRU (HYBRID SOLUTION).

The FMS mode of navigation is selected according to the following hierarchy: GPIR/Inertial DME/DME/Inertial DME/VOR/Inertial Inertial only. The GPIR/INERTIAL mode is selected as long as the following conditions are satisfied: GPIR position is available and with an estimated accuracy consistent with the intended operation. GPIR integrity is available and compatible with the applicable phase of flight requirement. As long as the GPS/INERTIAL mode is active, no DME/DME or VOR/DME radio updating is allowed. However, LOC updating can apply to GPS/INERTIAL position. In this navigation mode, N IR/GPS indication is displayed on the POSITION MONITOR page with N being the number of IRs used to compute mixed IR position. The selected hybrid GPIRS position is displayed on the POSITION MONITOR page in place of the radio position. The mixed IR position and the IR deviations displayed on the POSITION MONITOR page do not change and are still computed using pure IR inputs. Aircraft position is generated by a series of filters which use inertial position, GPIR position or radio position, and aircraft velocity as input. A position bias is computed once every second through the position bias filter. This position bias is computed as the difference between the GPIR position (or radio position) and the inertial position. The aircraft position is finally computed every 200 ms based on the corrected inertial position and the aircraft velocity using the aircraft position filter. The GPS/INERTIAL mode can be manually inhibited by pushing the line key adjacent to the DESELECT GPS indication on the SELECTED NAVAIDS page. FMGC computed integrity: When the GPIR position is available in the FMGC but the GPIR integrity is not delivered by the ADIRS, the FMGC is capable of computing an equivalent integrity called AIM (Alternate Integrity Monitoring), using IR data, during a limited period of time. The goal of this FMGC functionality is to improve the availability of the GPS Primary function in the cockpit.

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Figure 18

MMR GPS Operation

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DISPLAY OF GPS DATA ON MCDU GPS MONITOR The GPS data are displayed on the GPS MONITOR page of the MCDU. To get the GPS MONITOR page, push the DATA key on the MCDU, then the line key adjacent to the GPS MONITOR indication. The upper part is dedicated to GPS 1 data, the lower part to GPS 2 data. The following data are displayed: GPS position (lat/long) true track GPS altitude figure of merit (in meters) ground speed number of satellites tracked mode.

Progress page The progress page indicates whether the GPS is used by the FMGC for navigation. If it is used, the GPS PRIMARY indication is displayed. If it is not used, the GPS PRIMARY LOST message is shown.

PREDICTIVE GPS The integrity prediction results given by the GPS portion of the MMR on Flight Management System (FMS) request are displayed on the PREDICTIVE GPS page of the MCDU (from the progress page which displays required navigation accuracy and estimated position accuracy and GPS PRIMARY indication). The prediction concerns the destination (DEST) and any pilot entered waypoint (WPT) and the integrity availability (HIL < 0.3 Nm) is displayed by Yes (Y) or No (N) for the seven times defined by the five minutes increments for plus or minus 15 minutes around DEST or WPT.

ARRIVAL page To select a GPS approach, use the ARRIVAL page.

MCDU scratchpad GPS PRIMARY in white or GPS PRIMARY LOST in amber show on the MCDU scratchpad. SELECTED NAVAIDS page It is possible to select GPS for navigation computation within the FMS on the SELECTED NAVAIDS page. If you deselected GPS, the message GPS IS DESELECTED appears on the MCDU when a GPS approach starts.



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Figure 19

GPS Data on MCDU

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DISPLAY OF GPS MESSAGES ON ND

WARNING

GPS PRIMARY LOST amber message This message is displayed at the bottom of the image in all the ND modes when the GPS primary is lost (this message cannot be cleared from the MCDU). In this case, the GPS is not used for navigation (accuracy and integrity for the intended operation can still be met by the use of alternate navigation means).

GPS failure The GPSs are monitored by the both FWCs using a status word sent by each GPS. In case of GPS failure, the NAV GPS 1(2) FAULT message is displayed in the lower part of the upper ECAM DU. This message is accompanied by: activation of the MASTER CAUT lights on the glareshield aural warning: Single Chime (SC).

Display of GPS PRIMARY white message This message is displayed at the bottom of the image in all the ND modes when the GPS becomes primary (this message can be cleared from the MCDU). Display of GPS APP green message This approach message is displayed at the top of the image in all the ND modes when a GPS approach is selected in the flight plan.

NOTE:

THE FAILURE IS REMINDED ON THE INOP SYSTEM PAGE OF THE LOWER ECAM DU. THE MESSAGE DISPLAYED IS GPS 1(2).

Loss of the GPS primary navigation When the GPS navigation is lost for any reason, the navigation function is degraded and reverts to the traditional navigation function with IRS positions and radio positions if available (in this case the RNP (Required Navigation Performance) features are still available). Warnings are generated to indicate the loss of GPS PRIMARY navigation: GPS PRIMARY LOST message on the NDs (cannot be cleared) and MCDU (can be cleared) in case of GPS non-precision approach, an aural alert is generated (Triple Click) GPS/FMS position disagreement When GPS Primary is active and either FMGC 1 or FMGC 2 latitude (longitude) deviates from either MMR 1 or MMR 2 latitude (longitude) by more than 0.5 Nm, the NAV FMS/GPS POS DISAGREE and A/C POS...........CHECK messages are displayed in amber and cyan respectively on the ECAM DU. These messages are accompanied by: activation of the MASTER CAUT lights on the glareshield aural warning: Single Chime (SC).

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Figure 20

GPS Messages on ND

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BITE MENU CFDIU Interface BITE description The BITE facilitates maintenance on in-service aircraft. It detects and identifies a failure related to the MMR. The BITE of the MMR receiver is connected to the CFDIU. The BITE: transmits permanently MMR status and its identification message to the CFDIU memorizes the failures which occurred during the last 63 flight legs monitors data inputs from the various peripherals transmits to the CFDIU the result of the tests performed and selt-tests can communicate with the CFDIU through the menus. The BITE can operate in two modes: - the normal mode - the menu mode. Normal mode During the normal mode, the BITE monitors cyclically the status of the MMR. It transmits its information to the CFDIU during the concerned flight. In case of fault detection the BITE stores the information in the fault memories. These items of information are transmitted to the CFDIU. In case of ILS 1 (2) failure, the ILS waring message „NAV ILS 1 (2) FAULT” is shown on the lower part of the Engine/Warning Display, the MASTER CAUTION comes on and the single chime sounds. In case of GPS failure, the „NAV GPS 1(2) FAULT“ message is displayed in the lower part of the Engine/Warning Display (EWD). This message is accompanied by: - activation of the MASTER CAUT lights on the glareshield - aural warning: Single Chime (SC). NOTE:

Menu mode The menu mode can only be activated on the ground. This mode enables communication between the CFDIU and the MMR BITE by means of the MCDU. All the information displayed on the MCDU during the BITE Test configuration can be printed by the printer. The MMR menu mode is composed of: LAST LEG REPORT This menu contains the fault messages (class 1 internal and external) detected during the last flight. PREVIOUS LEGS REPORT This report contains the fault messages related to the external or internal failures (class 1) recorded during the previous 63 flight legs. LRU IDENTIFICATION Allows to display the P/N, the S/N and the SW/N of the equipment. GND SCANNING Based on the monitoring and fault analysis during flight, provides information of the failures detected while using this function. TROUBLE SHOOTING DATA Provides correlation parameters and snapshot data concerning the failure displayed in the LAST LEG REPORT and PREVIOUS LEGS REPORT. CLASS 3 FAULTS Allows to display the class 3 faults recorded during the last flight leg. SYSTEM TEST Allows a check of the correct operation of the MMR interrogator. GROUND REPORT Allows to present the class 1 or 3 internal failures detected on ground.

THESE FAILURES ARE REMINDED ON THE INOP SYSTEM PAGE OF THE SYSTEM DISPLAY (SD). THE MESSAGE DISPLAYED IS ILS 1(2) / GPS 1(2).

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Figure 21

MMR System Test (Fig. 1)

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Figure 22

MMR System Test (Fig. 2)

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1

GS scale and index (white and magenta)

2 LOC scale and index (white and magenta) 3

ILS data (magenta)

4

LOC flag (red)

5

G/S flag (red)

6

ILS flag (red)

1 5

6

4

3

2

6

3 1

ILS

4

5

2

ILS

EFIS mode selector ILS pushbutton

Figure 23

MMR ILS BITE Indication on PFD and ND

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LOCATION

ACTIVATION OF THE FRONT PANEL TEST General The face of the MMR is fitted with a handle, two attaching parts, a TEST pushbutton switch and four Light Emitting Diodes (LEDs). The four LEDs have the following name, color and function: TEST OK (green) indicates that no fault is detected during the initiated (by pushbutton switch or by MCDU) self-test or during the power-up test. MMR FAULT (red) indicates that an internal fault is detected by the MMR itself. BUS IN FAIL (red) indicates that no control input is available. TEST ANT (red) indicates that a failed antenna (or coaxial cable) is detected. The back of the MMR is equipped with one ARINC 600 size one connector, which includes three plugs: Top Plug (TP): connection with the GPS antenna Middle Plug (MP): service interconnection Bottom Plug (BP): connection with the power supply circuit, and the LOC and G/S coaxial interconnections.

GPS Antenna Two L-Band Antennas are mounted on the top of the fuselage, at the centerline, to receive signals from the GPS satellites. The GPS antenna is an active antenna with an integrated preamplifier and filter. It receives GPS signals at 1575.42 MHz and matches to a 50 -ohms coaxial cable at the input to the MMR. The antenna has a right-hand circular polarized and omnidirectional radiation pattern. The power supply of the preamplifier is provided by the MMR through the coaxial cable. NOTE:

THE ANTENNA CONNECTORS HAVE A HOLE TO INSTALL A LOCKWIRE AND SAFETY THE COAXIAL CABLE.

Localizer Antenna The localizer antenna is an airborne antenna used to receive LOC signals in the 108-112MHz range. It is a folded half-loop type driven by capacitive coupling. The antenna has two independent RF connectors used to feed two independent ILS receivers. Connector separation is provided by a hybrid junction in the antenna. Glide Slope Antenna The glide slope antenna is an airborne antenna used to receive GLIDE signals in the 329-335MHz range. It is a folded half-loop type driven by capacitive coupling. The antenna has two independent RF connectors used to feed two independent ILS receivers. Connector separation is provided by a hybrid junction in the antenna.



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Figure 24

MMR Front Panel Test

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Figure 25

MMR Location

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GPS-Antenna

Figure 26

LOC and G/S Antenna

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34-55

VOR/MARKER

DESCRIPTION General The VOR (VHF Omni-Range) system is a medium range navigation aid, which provides, when tuned to a station, the radial of the station the A/C is flying and the A/C angular deviation with respect to a selected course. The A320 uses two independent VOR systems. Both VOR receivers are equipped with a Marker module, but only VOR 1 Marker part is used. Tuning Auto Tuning In normal operation the VOR receiver 1 (2) is automatically tuned by the onside FMGC 1 (2) through the associated RMP 1 (2). In this case, the RMP is only used to transmit the frequency and course information from the FMGCs to the frequency input port A of the receiver. Manual Tuning Frequency and course data can by manually entered on the RAD/NAV page of the MCDUs. The FMGCs sent this information to the receivers in the same way like in then auto-tuning mode. FM Switching If a FMGC fails, a discrete is sent to the receiver (via the RMP) to activate the frequency input port B. This port receives information direct from the opposite FMGC. In this case, one FMGC tunes both VOR receivers. NAV Back Up Tuning If both FMGC fail, each VOR receiver must be tuned directly from the onside RMP. To do so, press the NAV and the VOR pushbutton on both RMPs. The RMP now uses manually entered data and not the data coming from the FMGC. A discrete reselects the frequency input port A, which is directly supplied from the associated RMP. A second discrete inhibits the data display on the RAD/NAV page of the MCDUs to indicate that no FMGC tuning is possible.

Inputs Each LGCIU sends discrete signals to the VOR receiver for internal BITE purposes. Indication All DMCs receive VOR data from both receivers such as VOR bearing, frequency, VOR course, VOR deviation and VOR identifier. Data of both systems are shown on Capts and F/Os ND. Only in ROSE VOR mode, the specific VOR 1 data (characteristics, course, deviation) is shown on Capts ND and specific VOR 2 data on F/Os ND.. All DMCs receive Marker data (OM, MM, AM) from VOR1 receiver to display it on both PFDs. The DDRMI receive VOR bearing from both VOR receivers to show the VOR bearing. Audio The VOR/MKR audio signal is processed by the receiver ,sent to the AMU and can be heard by the crew on headphones or cockpit loudspeaker. Users The FMGCs receives the VOR Data for navigation purpose during various flight phases. The CFDIU is used to communicate with the internal BITE functions of the VOR receivers (tests only available on ground). Warnings and Flags A faulty VOR system results in the following cockpit effects: Flags on PFD and ND Flags on DDRMI

Antenna The VOR antenna is common to both receivers. The antenna has two independent connectors, used to feed the two VOR receivers. VOR1 receiver is connected to the Marker antenna.



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Figure 27

VOR/MKR System Schematic

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INDICATION Marker Indication on PFD At the intersection of the LOC and G/S scale on the PFD the following Marker indication appears: AWY (Airways Marker) in white or OM (Outer Marker) in cyan or MM (Middle Marker) in amber. Marker audio signals are processed by the receiver and sent to the AMU.

Flags and NCD Indication on PFD If the Marker part of the VOR receiver fails or NCD is sent, no Marker indication is displayed.



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(

Figure 28

)

Marker Indication on PFD

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Normal Indication on ND ND in ROSE VOR Mode The white VOR deviation scale appears. The cyan course cursor or dagger shows the preselected course PSC in relation to the heading rose. The cyan deviation bar appears, when the VOR signal is valid. It moves perpendicular to the course cursor (1 dot = 5). When the deviation is out of range, the bar moves against one stop. The arrow shows the direction to the station (TO/FROM). The VOR information (white) shows: - VOR system and frequency - VOR course (PSC) - VOR identifier, when decoded by the VOR receiver - Tuning mode blank if autotuned M for manual tuning via MCDU R for manual tuning via RMP ND in ROSE NAV or ARC Mode When the VOR.D pushbutton on the EFIS control panel is pressed, all VOR/ DME stations contained in the FMGECs NAV DATA BASE are displayed on the ND, depending on the selected range: O for DME + for VOR ND in ROSE VOR, ROSE ILS, ROSE NAV or ARC Mode When the VOR/ADF selector is switched to VOR and the VOR signal is valid, the white VOR pointer appears and shows the bearing to the VOR station. The VOR station characteristics is displayed in the left or right lower corner of the ND and shows: - VOR system - Pointer symbol - VOR/DME frequency or identifier, if decoded by the VOR receiver - Tuning mode

Flags and NCD Indication on ND ND in ROSE VOR Mode If the VOR receiver fails, a red VOR1 (2) flag (flashing 9s, then steady) comes into view instead of VOR deviation scale and VOR information and the deviation bar goes out of view. In case of NCD, the VOR deviation bar goes out of view. If the course input fails, a vertical red dagger and a red course flag (CRS XXX) is shown . If the course information is NCD, the course data shows dashes and the dagger and the VOR deviation bar goes out of view. If the frequency information is fail or NCD, the frequency data and the deviation goes out of view. ND in ROSE VOR,ROSE ILS, ROSE NAV or ARC Mode and VOR/ADF Selector in position VOR If the VOR receiver fails, a red VOR1 (2) flag (flashing 9s, then steady) comes into view instead of VOR characterictics and the VOR bearing pointer goes out of view. In case of NCD, the VOR bearing pointer goes out of view and on VOR characteristics, the VOR identifier is replaced by the frequency. If the frequency information is fail or NCD, the VOR bearing pointer goes out of view. Identifier and frequency data are sent by the DME interrogator. Normal Indication on DDRMI The pointer shows VOR ground station bearing on the heading dial. Flags and NCD Indication on DDRMI If the VOR receiver fails, a red VOR flag comes into view and the bearing pointer is driven in 3 o’clock position. In case of NCD, the bearing pointer shows 3 o’clock position.



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VOR Deviation Bar

VOR Characteristics

Dagger

VOR Information

VOR Flag

Course Flag

VOR Pointer

VOR Flag

VOR Pointer 3 o‘clock

Figure 29

VOR/MKR Indication on ND and RMI

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FAULT ISOLATION AND BITE The different BITE menu selections are: LAST LEG REPORT PREVIOUS LEGS REPORT LRU IDENTIFICATION TEST Faults detected by the System and transfered to the CFDS causes the following messages displayed on the MCDU during BITE. VOR 1(2): NO DATA FROM CONTROL SOURCE There is no correct frequency data input on the active input port of the VOR receiver. RECEIVER The VOR receiver is faulty. NO DATA FROM CFDIU No connection to the CFDS.



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Capt PFD

Capt ND

F/O ND

RMI

F/O PFD

EFIS System

other Systems FMGC 1,2 AMU (Audio)

ÂÂÂÂÂÂÂÂÂ Â ÂÂÂÂÂÂÂÂÂ Â ÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂ

VOR Antenna MKR Antenna

VOR/MKR 1

ÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂ VOR/MKR 2

CFDS

RMP 1

other Systems FMGC 1,2 AMU (Audio)

RMP 2

FMGC 1

FMGC 2

ÂÂÂ ÂÂÂ

CFDS monitored

Figure 30

VOR/MKR BITE Schematic

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1

1

Figure 31

VOR/MKR CFDS BITE Menu

______________________________________________________________________________________________________________________________________________________________________________________________



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Figure 32

VOR/MKR CFDS BITE Test

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1

VOR Pointer (white)

2

VOR Course (Dagger) (cyan)

3

VOR Deviation (cyan)

4

MKR Indication

5

VOR Data (white)

6

VOR flag (red)

6

1

1

4

5

6 1

1

2

3 6

VOR/ADF Switches EFIS mode selector

6

Figure 33

6

5

5

VOR/MKR BITE Indication on PFD, ND and RMI

______________________________________________________________________________________________________________________________________________________________________________________________



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ACTIVATION OF THE FRONT PANEL TEST The front panel test can be activated in ground condition only by pushing the TEST pushbutton switch on the face of the receiver. During the first 3 seconds, all LEDs on the face of the receiver are on. During the next 3 seconds, all LEDs go off. During the last 3 seconds (or until the TEST pushbutton switch is released) the green VOR LED is on (except if a fault has been detected during the test). The name, color and function of the three LEDs are as follows: VOR (red) indicates that an internal fault is detected of the VOR Receiver VOR (green) indicates that no internal fault is detected of the VOR Receiver MKR (red) indicates that a internal fault is detected of the MKR receiver DATA IN (red) indicates that no control input is available (frequency).



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Figure 34

VOR/MKR Front Panel Test

______________________________________________________________________________________________________________________________________________________________________________________________



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LOCATION

Figure 35

VOR/MKR Location Receiver and Antenna

Page ______________________________________________________________________________________________________________________________________________________________________________________________ Revision No : 02 Issue Date : 21/05/2013


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Figure 36

VOR/MKR Location Control and Indication

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34-51

DME

DESCRIPTION General The DME system is a radio aid for medium range navigation, which provides the crew, when tuned to a station, with a digital readout of the slant range distance between the A/C and the selected ground station. The A320 uses two independent DME systems. Each system can work with up to 5 independent stations simultaneously. Tuning Auto Tuning In normal operation the DME interrogator 1 (2) is automatically tuned by the onside FMGEC 1 (2) through the associated RMP 1 (2). In this case, the RMP is only used to transmit the frequency and course information from the FMGCs to the frequency input port A of the receiver. VOR/DME and ILS/DME distance can be calculated and displayed simultaneously. Manual Tuning Frequency and course data can by manually entered on the RAD/NAV page of the MCDUs. The FMGCs sent this information to the receivers in the same way then in auto-tuning mode. FM Switching If a FMGC fails, a discrete is sent to the interrogator (via the RMP) to activate the frequency input port B. This port receives information direct from the opposite FMGC. In this case, one FMGC tunes both DME interrogators. NAV Back Up Tuning If both FMGC fail, each DME Interrogator must be tuned directly from the onside RMP. To do so, press the NAV and the VOR pushbutton on both RMPs. The RMP now uses manually entered data and not the data coming from the FMGC. A discrete reselects the frequency input port A, which is directly supplied from the associated RMP. A second discrete inhibits the data display on the RAD/NAV page of the MCDUs to indicate that no FMGEC tuning is possible. No ILS/DME indication is possible.

Antenna Each DME interrogator uses its own DME antenna for radio transmission and reception. A suppression signal is transmitted by the DME interrogator each time when in transmission mode to inhibit other systems working in same frequency range (ATC, TCAS) and to prevent simultaneous transmission. Inputs Each LGCIU sends discrete signals to the DME interrogator for internal BITE purposes. Indication All DMCs receive DME data from both interrogators such as DME distance, frequency and identifier. Data of both systems are shown on Capts and F/Os ND. ILS DME information is only shown on the associated PFD. The DDRMI receive DME data from both DME interrogators to show the DME distance in the upper left and right corner. Audio The DME audio signal is processed by the Interrogator and sent to the AMU and can be heard by the crew on headphones or cockpit loudspeaker in parallel to the VOR audio signal. Users The FMGCs receive DME Data (5 stations max.) for navigation purpose during various flight phases. The CFDIU is used to communicate with the internal BITE functions of the DME Interrogators (tests only available on ground). Warnings and Flags A faulty DME system results in the following cockpit effects: Flags on PFD and ND blank display in DDRMI



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CAPT PFD

PFD

Figure 37

DME System Schematic

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INDICATION Normal Indication on PFD If the ILS pushbutton on the EFIS control panel is pressed, on the lower left corner of the PFD the ILS DME distance is shown (if a ILS/DME station is available). On Capts PFD ILS/DME1 distance is displayed, on F/Os PFD, ILS/ DME 2 is displayed. If no ILS/DME station is available, the display is blank.

Flags or NCD Indication on PFD If the DME interrogator fails and the ILS pushbutton on the EFIS control panel is pressed, a red DME1 (2) flag (flashing 9s, the steady) comes into view instead of the DME distance. With NCD and ILS pushbutton pressed, dashes are displayed. If the frequency information is fail or NCD, the DME indication is blank.



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ILS DME Flag

ILS/DME Indication Capt PFD DME1 F/O PFD DME2

Figure 38

DME Indication on PFD

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Normal Indication on ND ND in ROSE ILS, ROSE VOR, ROSE NAV or ARC Mode When the VOR/ADF selector is switched to VOR, the DME distance is displayed in green under the associated VOR characteristics in the lower left or right corner. ND in ROSE NAV or ARC Mode When the VOR.D pushbutton on the EFIS control panel is pressed, all VOR/ DME stations around the aircraft and contained in the FMGCs NAV DATA BASE are displayed on the ND, depending on the selected range. Normal Indication on DDRMI The DDRMI shows DME1 and DME2 distance in the upper left and right corner.

Flags or NCD Indication on ND Flags or DME NCD indication are only displayed on ND, when in ROSE or ARC mode and the VOR/ADF selector is switched to VOR. If the DME interrogator fails, a red DME1 (2) flag (flashing 9s, the steady) comes into view instead of DME distance. In case of NCD, two dashes are shown. If the frequency information is fail or NCD, the DME indication is blank. Flags or NCD Indication on DDRMI If the DME interrogator fails, the DME1 (2) window is blanked. There is no DME flag. In case of NCD, dashes are shown. If the frequency information is fail or NCD, the DME indication is blank.



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VOR DME Indication

VOR DME NCD Indication

Figure 39

VOR DME Flag

DME Indication on ND and RMI

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FAULT ISOLATION AND BITE The different BITE menu selections are: LAST LEG REPORT PREVIOUS LEGS REPORT LRU IDENTIFICATION TEST Faults detected by the System and transferred to the CFDS causes the following messages displayed on the MCDU during BITE. DME 1(2) : NO DATA FROM CONTROL SOURCE There is no correct frequency data input on the active input port of the DME interrogator. RECEIVER The DME Interrogator is faulty. NO DATA FROM CFDIU No connection to the CFDS.



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Capt PFD

Capt ND

F/O ND

RMI

F/O PFD

EFIS System

other Systems FMGC 1,2


DME 2 Antenna

DME 1 Antenna

DME 1

ÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂ DME 2

AMU (Audio) ATC 1,2 TCAS (Supp.)

CFDS

RMP 1

other Systems FMGC 1,2 AMU (Audio) ATC 1,2 TCAS (Supp.)

RMP 2

FMGC 1

FMGC 2

ÂÂÂ ÂÂÂ

CFDS monitored

Figure 40

DME BITE Schematic

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1

1

Figure 41

DME CFDS BITE Menu

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Figure 42

DME CFDS BITE Test

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1

ILS DME Indication (magenta)

2

VOR DME Indication (white)

3

DME NCD Indication (white)

4

ILS DME flag (red)

5

VOR DME flag (red)

2

5

4 1

VOR/ADF Switches ILS pushbutton

3 Figure 43

3

2

5

DME BITE Indication on PFD, ND and RMI

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ACTIVATION OF THE FRONT PANEL TEST The front panel test can be activated in ground condition only by pushing the TEST pushbutton switch on the face of the interrogator. During the first 3 seconds, all LEDs on the face of the interrogator are on. During the next 3 seconds, all LEDs go off. During the last 3 seconds (or until the TEST pushbutton switch is released) the green R/T LED is on (except if a fault has been detected during the test). The name, color and function of the three LEDs are as follows: R/T (red) indicates that an internal fault is detected of the DME interrogator R/T (green) indicates that no internal fault is detected of the DME interrogator DATA IN (red) indicates that no control input is available (frequency).



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Figure 44

DME Front Panel Test

______________________________________________________________________________________________________________________________________________________________________________________________



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LOCATION

Figure 45

DME Location Interrogator and Antenna

FRA ______________________________________________________________________________________________________________________________________________________________________________________________ Revision No : 02 Issue Date : 21/05/2013


Page 1078

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Figure 46

DME Location Control and Indication

______________________________________________________________________________________________________________________________________________________________________________________________



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34-53

ADF

DESCRIPTION General The Automatic Direction Finder (ADF) system is a medium range radio navigation aid which provides: an indication of the relative bearing of the aircraft to a selected ground station (frequency range 190 to 1750 KHz). The ADF comprises two independent systems. Each system consists of: one receiver one loop and sense antenna

Antenna Each ADF receiver is connected to its own ADF antenna. The ADF antenna consists of one sense antenna and two loop antenna elements, which are preamplified by integrated amplifiers. A test loop input is not used.

Tuning Auto Tuning In normal operation the ADF receiver 1 (2) is automatically tuned by the onside FMGC 1 (2) through the associated RMP 1 (2). In this case, the RMP is only used to transmit the frequency and course information from the FMGCs to the frequency input port A of the receiver. Manual Tuning Frequency and course data can by manually entered on the RAD/NAV page of the MCDUs. The FMGCs sent this information to the receivers in the same way like in the auto-tuning mode. FM Switching If a FMGC fails, a discrete is sent to the receiver (via the RMP) to activate the frequency input port B. This port receives information direct from the opposite FMGC. In this case, one FMGC tunes both ADF receivers. NAV Back Up Tuning If both FMGC fail, each ADF receiver must be tuned directly from the onside RMP. To do so, press the NAV and the ADF pushbuttons on both RMPs. The RMP now uses manually entered data and not the data coming from the FMGC. A discrete reselects the frequency input port A, which is directly supplied from the associated RMP. A second discrete inhibits the data display on the RAD/NAV page of the MCDUs to indicate that no FMGC tuning is possible.

Indication All DMCs receive ADF data from both receivers such as ADF bearing, frequency and ADF identifier. Data of both systems are shown on Capts and F/Os ND.

Inputs Each LGCIU sends discrete signals to the ADF receiver for internal BITE purposes.

Audio The ADF audio signal is processed by the receiver and sent to the AMU and can be heard by the crew on headphones or cockpit loudspeaker. Users The CFDIU is used to communicate with the internal BITE functions of the ADF receivers (tests only available on ground). Warnings and Flags A faulty ADF system results in the following cockpit effects: Flags on ND



Page 1080

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Figure 47

ADF System Schematic

______________________________________________________________________________________________________________________________________________________________________________________________



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INDICATION Normal Indication on ND ND in ROSE VOR,ROSE ILS, ROSE NAV or ARC Mode When the ADF/VOR selector is switched to ADF and the ADF signal is valid, the green ADF pointer appears and shows the bearing to the ADF station. The ADF station characteristics is displayed in the left or right lower corner of the ND and shows: - ADF system - Pointer symbol - ADF frequency or identifier, if decoded by the ADF receiver - Tuning mode ND in ROSE NAV or ARC Mode When the NDB pushbutton on the EFIS Control Panel is pressed, all ADF stations contained in the FMGC’s NAV DATA BASE are displayed on the ND, depending on the selected range. NOTE:

Flags or NCD Indication on ND Flags or NCD indication are only displayed on ND, when in ROSE or ARC mode and the VOR/ADF selector is switched to ADF. If the ADF receiver fails, a red ADF1 (2) flag (flashing 9s, then steady) comes into view instead of ADF characteristics and the ADF bearing pointer goes out of view. In case of NCD, the ADF bearing pointer goes out of view and on ADF characteristics, the ADF identifier is replaced by the frequency. If the frequency information is fail or NCD, the ADF bearing pointer, the frequency or the identifier goes out of view.

IF ONLY ONE ADF SYSTEM IS INSTALLED, THE SINGLE ADF POINTER IS SHOWN, IF ADF/VOR SELECTOR 2 IS SWITCHED TO ADF.



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ADF Pointer

ADF Flag (Pointer disappears)

ADF Characteristics

Figure 48

ADF Indication on ND

______________________________________________________________________________________________________________________________________________________________________________________________



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COMPONENT DESCRIPTION Loop and Sense Antenna The combined loop and sense antenna operates in the 190 to 1750 kHz frequency range and consists of the following components enclosed in a fiberglass housing: a vertically polarized sense antenna which is omnidirectional in the horizontal plane two horizontally polarized loop antennas which are directional in the horizontal plane a test loop which enables a self-test of the antenna (not used). a printed circuit board which contains three pre-amplifiers used to amplify the loop and sense antennas signals. The pre-amplifiers are energized by plus or minus 12VDC from the ADF receiver. The output impedance of the antenna is 78 Ohm and the Voltage Standing Wave Ratio (VSWR) 1.2:1.



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ADF2

ADF1

(n.u.)

Figure 49

ADF Antenna

______________________________________________________________________________________________________________________________________________________________________________________________



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FAULT ISOLATION AND BITE The different BITE menu selections are: LAST LEG REPORT PREVIOUS LEGS REPORT LRU IDENTIFICATION TEST Faults detected by the System and transferred to the CFDS causes the following messages displayed on the MCDU during BITE. ADF 1(2) : NO DATA FROM CONTROL SOURCE There is no correct frequency data input on the active input port of the ADF receiver. RECEIVER The ADF receiver is faulty. NO DATA FROM CFDIU No connection to the CFDS.



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Capt ND

F/O ND

EFIS System


ADF 2 Antenna

ADF 1 Antenna

ADF 1

AMU (Audio)

ÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂ ADF 2

CFDS

RMP 1

AMU (Audio)

RMP 2

FMGC 1

FMGC 2

ÂÂÂ ÂÂÂ

CFDS monitored

Figure 50

ADF BITE Schematic

______________________________________________________________________________________________________________________________________________________________________________________________



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1

1

Figure 51

ADF CFDS BITE Menu

______________________________________________________________________________________________________________________________________________________________________________________________



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Figure 52

ADF CFDS BITE Test

______________________________________________________________________________________________________________________________________________________________________________________________



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Page 1090

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1

ADF pointer (green)

2

ADF data (green)

3

ADF flag (red)

1

1 3

3

2

2

VOR/ADF Switches

Figure 53

ADF BITE Indication on ND

______________________________________________________________________________________________________________________________________________________________________________________________



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ACTIVATION OF THE FRONT PANEL TEST The front panel test can be activated in ground condition only by pushing the TEST pushbutton switch on the face of the receiver. During the first 3 seconds, all LEDs on the face of the receiver are on. During the next 3 seconds, all LEDs go off. During the last 3 seconds (or until the TEST pushbutton switch is released) the green ADF LED is on (except if a fault has been detected during the test). The name, color and function of the three LEDs are as follows: ADF (red) indicates that an internal fault is detected of the ADF Receiver ADF (green) indicates that no internal fault is detected of the ADF Receiver DATA IN (red) indicates that no control input is available (frequency).



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green red red

Figure 54

ADF Front Panel Test

______________________________________________________________________________________________________________________________________________________________________________________________



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LOCATION

Figure 55

ADF Location Receiver and Antenna

Pag ______________________________________________________________________________________________________________________________________________________________________________________________ Revision No : 02 Issue Date : 21/05/2013


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Figure 56

ADF Location Control and Indication

______________________________________________________________________________________________________________________________________________________________________________________________



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34-42

RADIO ALTIMETER

DESCRIPTION General The LRRA system is a radio system, which determines the distance between the A/C and the terrain. The system is used during climb, approach and landing phase. The A320 uses two independent LRRA systems. Antenna Each transceiver works with one transmission and one reception antenna. Both antennas are identical. Inputs Each LGCIU sends discrete signals to the LRRA transceiver for internal BITE purposes. A function test inhibit signal from the EIU blocks the system test, when the associated engine is operating (engine 1 for system 1, engine 2 for system 2). Indication All DMC receives radio height data from both LRRA transceiver. In normal case, LRRA1 data is displayed on Capts PFD and LRRA2 data on F/Os PFD. If LRRA1 (2) fails, the PFD indication automatically switches over to the offside LRRA.

FAN Each transceiver is cooled by a associated fan. The fan is installed under the transceiver and receives power from the transceiver. Users The LRRA data is sent to the GPWS for different warning activations. The FWCs receive LRRA information to create the altitude call-outs during approach and warnings in case of failure. (The TCAS Computer needs radio height to modify the trigger levels of traffic and collision warnings.) The FMGCs and ELACs use radio height for navigation and automatic flight purposes. The CFDIU is used to communicate with the internal BITE functions of the LRRA transceivers (tests only available on ground and engine off). Warnings and Flags A faulty LRRA system results in the following cockpit effects: Flags on PFD (both systems fail and slats > 0 ) Master Caution Lights on the glareshield Aural Warning (Single Chime) NAV LRRA 1 (2,1+2) FAULT on the upper ECAM display.



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Figure 57

Radio Altimeter Block Diagram

______________________________________________________________________________________________________________________________________________________________________________________________



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INDICATION Normal Indication on PFD Radio Altitude This indication is on the bottom of the attitude sphere for height less than or equal 2500 ft. The dimension and the color of the digits change in relation to the height (RA) and the decision height (DH) as follows: - 400 ft < RA < 2500 ft 3mm green - DH+100 ft < RA < 400 ft 4mm green - RA < DH+100 ft 4mm amber The resolution of the display is also a function of height: - RA > 50 ft 10 ft steps - 5 ft < RA < 50 FT 5 ft steps - RA < 5 ft 1 ft steps Decision Height The DH value is shown on the right top corner of the PFD (FMA indication) as soon as the radio altimeter operates. The DH is entered on APP page of the MCDU and is transmit by the FMGCs if ILS approach is selected. When the radio height is lower then the DH, a amber DH warning message (flashing first, the steady) comes into view at the bottom of the attitude sphere. Rising Runway Indication Below 300 ft, the height is shown by the distance between the horizon line and the limit of the sector 2. The limit of the sector 2 moves up as the aircraft is in the descent phase. The distance between these two lines is proportional to the ground height (sensitivity 5 ft/mm). As it moves up, the limit line erases the graduations on the pitch scale. Red Ribbon When the aircraft is below 500 ft height, a red ribbon comes into view on the bottom and at the right of the altitude scale and moves up as the aircraft is in the descent phase. When the aircraft has touched the ground, the top of the ribbon is at the middle of the altitude window.

Flags and NCD Indication on PFD Failure on one system, the valid system is automatically switched to both Capt and F/O PFDs. - On the upper ECAM display the warning message appears: NAV LRRA 1 (2) FAULT. Failure of both systems, all information go out of view on both PFD. In slat extended configuration, a red RA warning message (flashing 3s, then remains on) appears in place of RA indication. - On the upper ECAM display the warning message appears: NAV LRRA 1+2 FAULT.

AUTOMATIC CALL OUTS (FWC) Altitude Call Outs Height 400 300 200 100 50 40 30 20 10 DH+100 DH

Call Out four hundred three hundred two hundred one hundred fifty forty thirty twenty ten hundred above minimum

If the time between two call outs is greater then 11 sec. (RH > 50 ft) or 4 sec. (RH < 50 ft), intermediate call outs are generated. Retard Announcement The retard-announcement is generated at: 10 ft if Autothrust or one Autopilot in LAND 20 ft, if no Autothrust or Autothrust and Autopilot not in LAND.



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DH Indication

DH Indication and DH Alert RA Flag

Horizon line

Red Ribbon RA Indication

Rising Runway Indication

PERF X 6R

Figure 58

Radio Altimeter Indication on PFD

______________________________________________________________________________________________________________________________________________________________________________________________



Page 1099

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FAULT ISOLATION AND BITE

ECAM WARNING In case of LRRA1 (2) failure, the LRRA waring message ”NAV RA1 (2) FAULT” is shown on the upper ECAM display, the MASTER CAUTION comes on and the single chime sounds.

The different BITE menu selections are: LAST LEG REPORT PREVIOUS LEGS REPORT LRU IDENTIFICATION ARINC TEST RAMP TEST Note: The tests are possible, if no inhibit signal form the EIUs are present. Faults detected by the System and transferred to the CFDS causes the following messages displayed on the MCDU during BITE. RADIO ALTM 1(2) : TRANSCEIVER The LRRA Transceiver is faulty. RECEPTION ANTENNA The Reception Antenna or coaxial cable is faulty. TRANSMISSION ANTENNA The Transmission Antenna or coaxial cable is faulty. NO DATA FROM CFDIU No connection to the CFDS.



Page 1100

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Capt PFD

F/O PFD

ECAM Warning

EFIS System

EIU1

other Systems FMGC 1,2 ELAC 1,2 FWC 1,2 GPWS


EIU2 Test inhibit

LRRA 1

ÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂÂÂ ÂÂÂ ÂÂÂÂÂ ÂÂÂÂÂÂ ÂÂÂÂÂ ÂÂÂÂÂÂ ÂÂÂÂÂ ÂÂÂÂÂÂ ÂÂÂÂÂ ÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂÂ LRRA 2

CFDS

other Systems FMGC 1,2 ELAC 1,2 FWC 1,2

CFDS monitored

ÂÂÂÂÂ ÂÂÂÂÂÂ ÂÂÂÂÂ ÂÂÂÂÂÂ ÂÂÂÂÂ ÂÂÂÂÂÂ ÂÂÂÂÂ ÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂÂ Receive Antenna

Transmit Antenna

Receive Antenna

Figure 59

Transmit Antenna

Radio Altimeter BITE Schematic

______________________________________________________________________________________________________________________________________________________________________________________________



Page 1101

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1

1

Figure 60

Radio Altimeter CFDS BITE Menu

______________________________________________________________________________________________________________________________________________________________________________________________



Page 1102

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Figure 61

Radio Altimeter CFDS BITE Test

______________________________________________________________________________________________________________________________________________________________________________________________



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Page 1104

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1

LRRA Indication

2

LRRA flag (LRRA 1+2 fail and Slats >0)

1 1 2

1

Figure 62

Radio Altimeter BITE Indication on PFD

______________________________________________________________________________________________________________________________________________________________________________________________



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ACTIVATION OF THE FRONT PANEL TEST The front panel test can be activated by pushing the pushbutton switch on the face of the receiver, if the test inhibit discrete is not active. the green SYSTEM OK LED comes on at the beginning of the test. It remains on until the TEST pushbutton switch is released, if no fault is detected. the three red LEDs comes on when a fault is detected. In this case, they remain on until the TEST pushbutton switch is released: The name, color and function of the three LEDs are as follows: - the R/T UNIT LED indicates a transceiver fault - the ANT TX LED indicates a fault in the transmission antenna circuit. - the ANT RX LED indicates a fault in the reception antenna circuit. NOTE : The green LED remains on until the pushbutton switch on the face of the transceiver is released.



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Figure 63

Radio Altimeter Front Panel Test

______________________________________________________________________________________________________________________________________________________________________________________________



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LOCATION

Page: 106 Figure 64 Radio Altimeter Location Transceiver and Antenna ______________________________________________________________________________________________________________________________________________________________________________________________ Revision No : 02 Issue Date : 21/05/2013


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Figure 65

Radio Altimeter Location Control and Indication

______________________________________________________________________________________________________________________________________________________________________________________________



Page 1109

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34-41

WEATHER RADAR

DESCRIPTION General The aircraft is equipped with an X-band weather radar system. This system complies with ARINC Characteristics 708. The weather radar enables detection and localization of the atmospheric disturbances in the area defined by the antenna scanning: plus or minus 90 deg. of aircraft centerline and up to 320 NM in front of the aircraft. In addition the weather radar system enables: detection of turbulence areas caused by the presence of precipitations NOTE: THERE IS NO DETECTION OF TURBULENCE IN CLEAR SKY. presentation of terrain mapping information by the combination of the orientation of the radar beam and of the receiver gain. Five color displays are used to show precipitations and turbulences to the crew. The weather radar system, consists of: a transceiver a single control unit a single antenna drive an antenna a dual transceiver mounting tray with a wave guide switch. Indication The weather radar image is shown on the Captain and First Officer Navigation Displays (ND). The NDs are connected to the three Display Management Computers (DMC) and to the Captain and First Officer EFIS control sections of the FCU. Control Panel The control panel provides mode of operation, antenna tilt and gain information. Discretes enables the transceivers and activates the wave guide switch to connect the antenna to the transceiver.

Antenna The weather radar antenna is energized and monitored in azimuth and elevation by the transceiver. The radio frequency signals are exchanged between transceiver and the antenna via a wave guide. The antenna scans a 180 sector in azimuth and has a tilt coverage of " 15. Inputs The ADIRUs give the pitch and roll angles to ensure antenna stabilization and the ground speed for Doppler mode correction. ADIRU 1 supplies WX TXR 1. ADIRU 3 is automatic back up for the systems and can by manually activated via the ADIRS TXR switch. The EFIS control panels transmit the range selected on Capts and F/Os side in order to create two WX images. A signal from the wave guide switch activates the transceiver, when the wave guide is correctly connected to this transceiver. Each LGCIU sends discrete signals to the WX transceiver for internal BITE purposes. Users The CFDIU is used to communicate with the internal BITE functions of the WX-T ransceivers (tests only available on ground). NOTE:

IF THE ENHANCED GPWS IS OPERATIVE, THE WR IMAGE IS REPLACED BY THE TERRAIN IMAGE, ON THE CAPTAIN AND FIRST OFFICER NDS, DURING A TERRAIN ALERT OR A CREW ACTION.

Warnings and Flags A faulty WX system results in the following cockpit effects: Flags or Fault-Messages on NDs no WX-image on ND (depending on type of failure).

Page: Page: 108

______________________________________________________________________________________________________________________________________________________________________________________________



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Figure 66

Weather Radar Block Diagram

______________________________________________________________________________________________________________________________________________________________________________________________



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CONTROL Weather radar control unit The face of the weather radar control unit includes the following controls:

1 SYS ON/OFF switch This switch enables: the activation of the transceiver the suppression of the radar image on the NDs when these NDs are in the ARC or ROSE mode.

2 Mode selection WX This mode corresponds to the normal operation in weather detection. The radar images are displayed on the NDs in four colors (black,green, yellow, red); their intensity corresponds to the strength of the return signal. WX +T This mode corresponds to operation in weather and turbulence detections. All turbulent (within 50 NM) and non turbulent areas beyond 50 NM are displayed in the conventional black, green, yellow and red as in weather (WX) mode. TURB This mode corresponds to operation in turbulence detection. Turbulence detection is limited to the first 50 NM regardless of the weather radar range selected and displayed. Turbulence areas are displayed on the NDs in magenta. MAP This mode is only used for display of the ground map. A combination of transceiver gain, antenna position (TILT) and range selection enables the display of a larger area. If the image is too bright, due to too great reflection intensity, it can be dimmed by the GAIN potentiometer, item 4.

3 TILT control The TILT selector switch enables the variation of the antenna elevation angle in 1/4 deg. steps on a non-linear scale graduated in degrees, within a range of +15 deg. (UP) to -15 deg. (DOWN) in relation to a horizontal plane defined by the stabilization system. This antenna elevation angle is displayed in cyan in the R lower corner of the ND and progresses in steps of 0.25 degrees. If the antenna position is different from the TILT selector switch position, a red ANT failure warning message replaces the TILT indication in the R lower corner of the ND.

4 Gain control The GAIN potentiometer, item 4, is used to adjust the sensitivity of the receiver in WX, WX+T, TURB and MAP modes. The CAL gain position provides minimum gain setting and corresponds to the normal gain for operation. In this case, the radar system is aligned to give an accurate representation of rain levels corresponding to the real weather situation. The manual use of the GAIN potentiometer mainly allows to see as many precipitations as possible and in particular to view very light rain. In MAX position, the receiver sensitivity, the transmitter pulse width and the antenna beam width are increased. MAX gain is used for the same reasons as manual use and also to better display the leading edges of cells and to view patterns and characteristics of these cells. Therefore the GAIN potentiometer must be in the CAL position to determine the actual calibrated precipitation rate before taking an avoidance decision.

5 GND CLTR SPRS switch Activation of the ground clutter suppression switch in the WX mode reduces the intensity of the ground clutter.

FRA US/T WB 03.01.2001

Page: Page: 110

______________________________________________________________________________________________________________________________________________________________________________________________



Page 1112

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4

3

2

1

Figure 67

5

Weather Radar Control Panel

______________________________________________________________________________________________________________________________________________________________________________________________



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EFIS control panel (on the FCU) Only the controls related to the selection and use of the radar image on the NDs are described. A mode selector switch, item 1, made up of a rotary switch enables the selection of the ROSE or ARC function for display of a weather radar image on the CAPT and F/O NDs. A scale selector switch, item 2, common to EFIS, FMGS and radar systems, enables the selection of 10, 20, 40, 80, 160 or 320 operation range in nautical miles (NM) for display of the weather radar image on the CAPT and F/O NDs. EFIS switching panels The CAPT and F/O EFIS switching panels which are connected to the CAPT and F/O NDs, include ND concentric potentiometers which enable the brightness adjustment of the image displayed on the NDs. The outer knob of each potentiometer controls the brightness of the radar image only, item 3.

Utilization of the EFIS control panels and EFIS switching panels

1 Mode selector switch This mode selector switch enables the image display on the corresponding ND whenever the ARC or ROSE mode is selected and the transceiver is supplied. In that case, the radar image is displayed in the background of the navigation image.

2 Scale selector switch This selector switch enables the display of the range selected for an optimum use of the radar image on the corresponding ND. For each of the following ranges: 10, 20, 40, 80, 160 and 320, four concentric range arcs are displayed respectively spaced 2.5, 5, 10, 20, 40 and 80 NM, when the mode selector switch is in the ARC position. Only 2 range arcs are displayed in the ROSE mode.

3 Radar image brightness control The ND outher potentiometer enables the adjustment of brightness and contrast of radar echoes in relation to the navigation image, which is superimposed. However, the adjustment range does not allow total extinction of the image. The OFF position of the potentiometer corresponds to the minimum brightness. The BRT position corresponds to the maximum brightness. NOTE:

A PHOTOELECTRIC CELL ASSOCIATED WITH EACH ND ALSO ADJUSTS IMAGE BRIGHTNESS AS A FUNCTION OF AMBIENT LIGHT VARIATIONS.

Page: Page: 112

______________________________________________________________________________________________________________________________________________________________________________________________



Page 1114

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Figure 68

Weather Radar EFIS and ND Control

______________________________________________________________________________________________________________________________________________________________________________________________



Page 1115

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INDICATION Normal Indication on ND The weather radar image is shown on the ND, when the EFIS mode selector is switched to a ROSE or the ARC mode. The image depends on range selection. The system shows the disturbance intensity through the use of colors which vary with the atmospheric precipitation rate. The disturbances are shown to the crew members on the NDs with different colors: black, green, yellow, red to quantify the precipitation rates magenta to represent the turbulence areas up to 50 NM.: PRECIPITATION RATE | COLOR OF ECHOES --------------------——— |----------------------———————— less than 0.76 mm/h | black from 0.76 to 3.81 mm/h | green from 3.81 to 12.7 mm/h | yellow from 12.7 to 50.8 mm/h | red ---------------------------------------------—————————— | PRECIPITATION RATE | COLOR OF ECHOES —————————————————————————————————— Turbulence | 50.8 mm/h and above | magenta -----------------------------------------------------————— The actual operating range of the system is 320 NM. The peak power emitted is approximately 100 W. The antenna scans the 180 deg. sector in azimuth 15 times per minute. Additionally, the weather radar may be used as a navigation aid. In the mapping mode, it allows identification of major changes in the ground map: (e.g. a sea coast, an estuary, a lake, a mountain, an island, a big city, etc.). The brightness of the weather radar image is separately adjustable with a brightness control potentiometer. In the lower right corner, the antenna tilt is displayed in green. If the gain potentiometer is not in the AUTO position, MAN is displayed in green in front of the tilt indication.

Flags and NCD Indication on ND When the WX system fails, in the lower right corner a warning message comes into view and shows the faulty components. The tilt indication goes out of view. There are failures which result in the loss of the radar image (red warnings) and failures, which do not effect the radar image (amber warnings). If there is no range input, the radar image is lost and the WXR RNG message appears. Failures which result in the loss of the radar image. The corresponding messages are displayed in red WXR R/T : indicates a failure of the weather radar transceiver WXR ANT : indicates a failure of the weather radar antenna WXR CTL : indicates a failure of the weather radar control unit WXR RNG : indicates an error of comparison between the range from the EFIS control panel and the copy data received on the symbol generator via the radar data bus. Failures which do not affect the radar image. The corresponding messages are displayed in amber WXR WEAK : indicates the loss of the transceiver calibration WXR ATT : indicates an attitude failure from the ADIRU WXR STAB : indicates the loss of the radar antenna stabilization WXR TEST : indicates the selection of the radar TEST mode. NOTE:

IN CASE OF VENTILATION FAILURE OR WHEN THE BLOWER AND EXTRACT PUSHBUTTON SWITCHES ARE BOTH RELEASED, THE COLORED BACKGROUNDS OF THE WEATHER RADAR IMAGE DISAPPEAR.

FRA US/E Mk 18.12.95

Page: Page: 114

______________________________________________________________________________________________________________________________________________________________________________________________



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Weather Image

WX-Mode Tilt Indication

Ground Image Fault Message Manual Gain Indication

-

MAP-Mode Figure 69

Weather Radar Indication on ND

______________________________________________________________________________________________________________________________________________________________________________________________



Page 1117

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FAULT ISOLATION AND BITE The different BITE menu selections are: LAST LEG REPORT PREVIOUS LEGS REPORT LRU IDENTIFICATION TEST Faults detected by the Weather Radar System or the connected systems causes the following CFDS and/or BITE messages. RADAR 1 TRANSCEIVER Failure of the Weather Radar Transceiver RADAR 1 CONTROL UNIT Failure of the Weather Radar Control Unit RADAR 1 ANTENNA Failure of the Weather Radar Antenna RADAR 1 MOUNTING TRAY Failure of the Waveguide Switch RADAR 1 NO DATA FROM ADIRU Loss of ADIRU Input Data NO RADAR 1 DATA Loss of the Bus Output from Weather Radar Transceiver to the CFDIU DMC 1: NO WXR 1 DATA Loss of WXR Data or Bus Fault DMC 2: NO WXR 1 DATA Loss of WXR Data or Bus Fault DMC 3: NO WXR 1 DATA Loss of WXR Data or Bus Fault

FRA US/T WB 03.01.2001

Page: Page: 116

______________________________________________________________________________________________________________________________________________________________________________________________



Page 1118

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Capt ND

F/O ND

EFIS System

ÂÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂ ÂÂÂÂÂÂ ÂÂÂÂÂÂ ÂÂÂÂÂÂ ÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂ WXR Control

other Systems: -FCU (EFIS Ctrl Capt) (EFIS Ctrl F/O) ADIRU 1,3

WXR 1

WXR 2 MOUNT

CFDS

-ATT,HDG Transfer

WG

Antenna

Figure 70

ÂÂÂ ÂÂÂ

CFDS monitored

Weather Radar BITE Schematic

______________________________________________________________________________________________________________________________________________________________________________________________



Page 1119

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Figure 71

Weather Radar CFDS BITE Test

______________________________________________________________________________________________________________________________________________________________________________________________



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BITE TEST The TEST mode becomes active when it is selected through the CFDS via MCDU 1 or 2. It enables an operational check of the main circuits which constitute the system. The transmission/reception channel is tested for less than one second, then a special test pattern is displayed on the NDs as long as TEST mode is active. Moreover the elevation and azimuth control circuits of the antenna drive are excited during the test period. The complete test period lasts 15 seconds approximately. When the TEST mode is activated: The antenna carries out an elevation scanning sequence from up to down positions (+ 15 deg., - 15 deg.) then an azimuth scanning sequence from right to left, then stabilizes at 0 deg, perpendicular to the aircraft centerline. It should be noted that: the special test pattern, with TEST indication displayed in the R lower corner of the NDs, can only be displayed if no fault is detected the antenna no longer responds to the stabilization signals from the ADIRU when the TEST mode is active. At the first ground supply, the antenna test sequence is performed independently of the TEST mode selection and SYS switch. The test image is not displayed on the ND. This antenna test sequence is the same as the Bite test sequence.

Figure 72

Weather Radar Test Pattern on ND

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Figure 73

Weather Radar Location Transceiver and Antenna

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Figure 74

Weather Radar Location Control and Indication

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34-41

WXR/PWS

DESCRIPTION General The aircraft is equipped with an X-band weather radar system with Predictive Windshear capability. The weather radar system enables: detection and localization of the atmospheric disturbances in the area defined by the antenna scanning: plus or minus 90 deg. of aircraft centerline and up to 320NM in front of the aircraft, detection of turbulence areas caused by the presence of precipitations up to a distance of 40NM, presentation of terrain mapping information by the combination of the orientation of the radar beam and of the receiver gain, detection of a microburst windshear event in the area defined by the antenna scanning: plus or minus 60 deg. presentation of windshear events within an area plus or minus 30 deg. of aircraft centerline and up to 5NM in front of the aircraft. NOTE:

A MICROBURST IS A COOL SHAFT OF AIR, LIKE A CYLINDER, BETWEEN 1000 AND 3000 FT. WHEN IT ENCOUNTERS THE GROUND (AIRFLOW VELOCITY FROM 40 TO 110 KTS) THE DOWNWARD MOVING AIRFLOW IS TRANSLATED TO A HORIZONTAL FLOW (FROM 80 TO 220 KTS), AT THE BASE OF THE AIR SHAFT. TWO TYPES OF MICROBURST EXIST, WET AND DRY. Five color displays are used to show precipitations, turbulence and ground mapping to the crew. The location of the windshear events is indicated by an icon (symbol consisting of alternating red and black arcs).

System Architecture The weather radar system is composed of items closely associated with its operation, such as peripherals supplying parameters, EFIS display units or maintenance functions. The weather radar image is shown on the CAPT and F/O Navigation Displays (ND).The NDs are connected to the three Display Management Computers (DMC) and to the CAPT and F/O EFIS control panels of the FCU. The weather radar and windshear detection image is shown on the Captain and First Officer Navigation Displays (ND) and the windshear warning is shown on Captain and First Officer Primary Flight Displays (PFD) and on the upper ECAM DU. The weather radar system consists of: a transceiver, a dual control unit, a dual antenna drive, an antenna, a transceiver dual mounting tray with a wave guide switch. NOTE:

ONLY THE ANTENNA DRIVE OF THE WXR/PWS SYSTEM IS THE SAME AS THAT OF THE NORMAL WXR SYSTEM.

NOTE:

IF THE ENHANCED GPWS IS OPERATIVE, THE WR IMAGE IS REPLACED BY THE TERRAIN IMAGE ON THE NAVIGATION DISPLAYS, DURING A TERRAIN ALERT OR A CREW ACTION.

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Figure 75

WXR/PWS Block Diagram

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Weather Radar Transceiver The receiver-transmitter is the heart of the WR/PWS, the additional necessary wiring and interfaces enable the weather radar transceiver to operate as a PWS (WR/PWS). The receiver-transmitter ensures the following functions: generation of the very short intense pulses of microwave energy via an X-band wave guide to the antenna, and the processing of their echoes (radio frequency signals) to obtain the desired information, the receiver signal is formatted into 1600-bit ARINC 453 words and sent to the DMCs, acquisition of data from Radio Altimeters (RA1 and RA2) and other specific interfaces, windshear event detection and generation of the appropriate signal, BITE function the system. The radar transceiver is de-activated when the 1/OFF/2 switch on the weather radar control unit is set to the OFF position or by placing the mode selector switches on both EFIS control sections in any position other than ROSE or ARC (exception windshear function). Weather Radar Control Unit The dual control unit generates a 32-bit (label 270) serial control word which describes the selected operating modes (1/OFF/2, WX, WX + T, TURB, MAP, GND CLTR SPRS, PWS). WR Antenna Drive The weather radar system has one dual antenna drive which is the interface of the transceiver to control and monitor the azimuth and elevation of the antenna. Weather Radar Antenna The antenna is used for transmitting and receiving radar radio frequency signals. Transceiver Dual Mounting Tray with a Wave Guide Switch It allows to install each transceiver on the aircraft rack and connects the activated transceiver to the wave guide. The wave guide switch is integral with the mounting tray. It ensures switching of the RF signal from the antenna to each transceiver.

Peripherals (Inputs) The transceiver receives digital serial data inputs from the components: Radio Altimeter - The Radio Altimeter provides altitude information over two ARINC 429 bus inputs to the WR/PWS. This data is used for automatic activation of the windshear function. Air Data Reference - true airspeed data and computed airspeed used for velocity calculations, - altitude data used for sensitivity time control (STC) calculations, - corrected altitude data used only when altitude data not available. Inertial Reference - pitch and roll data for the stabilization and control of the antenna, - east/west velocity and north/south velocity used for velocity calculations, - ground speed used for velocity calculations, - track angle and drift angle used for velocity calculations, - true heading, - Magnetic heading. Centralized Fault Display Interface Unit (CFDIU) EFIS Control Section - The receiver/transmitter receives one bus from the CAPT EFIS control section and another one from the F/O EFIS control section. The transceiver receives the following discrete inputs: Ground/flight signal and landing gear extended signal - is used to determine the identifying flight phase for BITE - Landing gear extended signal is used to determine transition from landing mode to takeoff mode to identify a GO AROUND condition. In this case the appropriate aural message is generated. Qualifiers A and B signals - In order to activate automatically the windshear function, one of each qualifier A and one of each qualifier B have to be valid. Windshear function enable signal - This discrete input provided by the WXR control unit through PWS/OFF/ AUTO switch enables the windshear function. Also transmitted to the DMCs which use it for the logic of windshear messages displayed.

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Figure 76

WXR/PWS Data Acquisition

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Peripherals (Outputs) The transceiver provides the following outputs: Displays The WXR/PWS is connected to the DMCs by an ARINC 453 bus to transmit the weather radar data and windshear data on the data word. All the weather and windshear data received by the DMCs are processed to display weather radar image and windshear events by the Electromic Instrument System (EIS). The Navigation Display (ND) provides the following indications: - weather radar image - windshear events location for advisory, caution or warning alert - windshear failures. The Primary Flight Display (PFD) provides all visual alerts for caution or warning alert. NOTE:

THE FLIGHT WARNING COMPUTERS (FWC) AND THE FDIU RECEIVE WR/PWS DATA THROUGH THE DMCS. THESE DATA ARE USED BY THE FWCS TO DISPLAY PWS FAILURE AND WINDSHEAR FUNCTION OFF.THE FDIU RECORDS THE WINDSHEAR ALERT AND FAILURE. Centralized Fault Display System (CFDIU) The WR/PWS is connected to the CFDS to transmit the following words: label 354: LRU identification P/N and S/N (coded in ISO5), label 356: fault message (coded in ISO5), label 377: equipment identification. Audio Mixing Box An analog audio output allows to transmit the aural alert windshear (synthetic voice message) to an audio mixing box connected to loud speakers. Enhanced Ground Proximity Warning System (Enhanced GPWS) The Enhanced GPWS receives WR/PWS alerts from WXR1 Hazard bus to determine the alert priorities. Predictive Windshear alerts override a terrain display and revert to the WR display with the corresponding windshear data. The alert priorities between the WR/PWS and the Enhanced GPWS have been defined as follows:

123456-

WR/PWS Warning, WR/PWS Caution, Terrain Warning, Terrain Caution, WR/PWS advisory (no audio), Terrain background (no audio).

Audio Inhibit Discrete Signals These discretes are used to indicate whether the aural alert output has to be active or not. predictive windshear aural alerts (audio inhibit discrete input) are inhibited by the Reactive Windshear System and stall warning. predictive windshear audio inhibit discrete output is used to inhibit other aural alerts generated by systems such as: Traffic Alert and Collision Avoidance System (TCAS) or Ground Proximity Warning System (GPWS) or other FWC warnings. This inhibition occurs each time there is a PWS aural alert. Pin Programming audio level program pins set the audio output level of the synthetic voice (windshear aural alert). SDI (Source Destination Identification Encoder) program pins encode the location of the WR/PWS unit on the aircraft. qualifier polarity program pins: for both qualifiers, this pin program indicates the validity of the signal. CFDIU interface program pins: when the second WR/PWS is installed on the aircraft, this program pin is activated. The WR/PWS can communicate with the CFDIU. caution alert audio program: two program pins are provided to select the type of windshear caution aural alert. The MONITOR RADAR DISPLAY synthetic voice is generated instead of the chime. windshear function enable program pins is used to activate the windshear function. windshear function bite enable signal allows to send to CFDIU failure related to the predictive windshear function.

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Figure 77

WXR/PWS Block Diagram

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CONTROLS AND INDICATING The various system controls are grouped on the weather radar control unit and on the EFIS control sections of the FCU. Weather radar control unit The face of the weather radar control unit includes the following controls:

1

A mode selector switch, which enables the selection of the WX, WX+T, TURB or MAP function.

2

A TILT selector switch, which enables the control of the antenna elevation. Antenna position is read in degrees, opposite the notch on the switch: - either from 0 to 15 deg. upwards (UP) - or from 0 to 15 deg. downwards (DN). In windshear mode, the tilt control is automatic in the WR/PWS for the scanning. However, the tilt displayed on ND is in accordance with the one selected on the radar control unit.

3

A GAIN potentiometer, which enables the manual adjustment of the transceiver gain. In the windshear position, the gain control is automatic in the WR/PWS for the scanning.

4

A switch, with three stable positions 1/OFF/2, which enables the selection of the transceiver 1 or 2 and the deactivation of the transceivers.

6

A GND CLTR SPRS switch, which enables the selection of the ground clutter suppression (ON/OFF).

EFIS control section (on the FCU) In this part, the controls related to the selection of WX and W/S functions are described.

7

A mode selector switch, made up of a rotary switch enables the selection of the ROSE or ARC function for display of a weather radar image on the CAPT and F/O NDs. If neither ROSE or ARC mode is selected, the message W/S CHANGE MODE is shown on both NDs, if there is a windshear alert. The color depends on the W/S alert level.

8

A scale selector switch, common to EFIS, FMGS and radar systems, enables the selection of 10, 20, 40, 80, 160 or 320 operation range in nautical miles (NM) for display of the weather radar image on the CAPT and F/O NDs. Windshear information is presented in the 10 NM range only. If a windshear alert is generated but the selected range is greater than 10 NM, the message W/S: SET RNG 10 NM is shown on the NDs. The color depends on the W/S alert level. Lighting/LOUDSPEAKER control panel CAPT and F/O lighting/LOUDSPEAKER control panels 301VU and 500VU which are connected to CAPT and F/O NDs, include ND concentric potentiometers for adjusting the brightness of the image displayed on the NDs.

5

A PWS/AUTO/OFF switch, which enables the selection of the windshear function. In AUTO position, the windshear detection is automatic if altitude is lower than 2300 ft and qualifiers A and B are valid. This automatic operation can be inhibited when the switch is in the OFF position.

9

The outer knob of each potentiometer controls the brightness of the radar image only.

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3

7

2

5

8

9 4

1

6

Figure 78

WXR/PWS Controls and Indicating

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WEATHER RADAR DATA DISPLAYED ON ND’S Indication on ND On the figure, the details A) and B) respectively correspond to the ARC and ROSE ND modes for which the display of the radar image is possible. Messages inform the crew of the tilt angle and gain selected on the weather radar control unit. Other messages indicate the failures which affect the operation of the radar system. All these messages are displayed in the R lower corner of each ND whenever a radar image is selected. Tilt information and gain selection are displayed on the ND when no failure warning message is generated, or when the TEST mode is not selected. The various failures which can affect the radar image are listed in decreasing order of importance. If several failures occur, only the most important one is displayed (Ref. details C) and D) on the figure). Two types of failures can affect the radar system: Failures which result in the loss of the radar image The corresponding messages are displayed in red - WR : indicates a failure of the weather radar transceiver R/T - WR : indicates a failure of the weather radar antenna ANT - WR : indicates a failure of the weather radar control unit CTL - WR : indicates an error of comparison between the range RNG from the EFIS control section and the copy data received on the symbol generator via the radar data bus. NOTE:

Failures which do not affect the radar image The corresponding messages are displayed in amber - WR : indicates the loss of the transceiver calibration WEAK - WR : indicates an attitude failure from the ADIRU ATT - WR : indicates the loss of the radar antenna stabilization STAB - WR : indicates the selection of the radar TEST mode. TEST

IN CASE OF INSUFFICIENT AVIONICS VENTILATION, THE WEATHER RADAR IMAGE IS LOST IN ORDER TO PREVENT NDS OVERHEAT.

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Figure 79

WXR/PWS Weather Radar Indication on ND

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WINDSHEAR INFORMATION ON PFD AND ND Windshear indications The location of a windshear phenomenon is indicated to the crew by means of an icon superimposed on the radar image. This icon consists of alternating red and black arcs. For 10 NM range selection and above, yellow radial lines appear at the edges and start beyond the windshear event. These lines, superimposed on the radar image, continue to the edge of the display area to provide directional information for the event. The windshear data are always displayed even if the 1/OFF/2 selector switch on the radar control unit is set to OFF. The PWS/OFF/AUTO switch on the radar control unit has to be set to AUTO. Logic of scanning mode The antenna scan pattern varies depending on the mode of operation. Weather radar scan pattern In weather radar mode, the antenna scans a 180 deg. in azimuth and has tilt (pitch) coverage of plus or minus 15 deg. Stabilization limits are plus or minus 25 deg. in the pitch axis and plus or minus 40 deg. in the roll axis. The antenna scans the zone 15 times per minute. Beam opening is 3.6 deg. in elevation and 3.5 deg. in azimuth. An antenna scanning is performed in 4 seconds, this causes the transmission of 720 data words to the data bus lines. When the two ranges selected on both EFIS control panels are identical, the radar images displayed on the NDs are refreshed every 4 seconds. On the contrary, when two different ranges are selected on the EFIS control panels, the images are refreshed every 8 seconds. Weather and windshear scan pattern When the system is placed into alternate weather/windshear mode, the weather processing is operating during left-to-right scans and windshear processing is operating during right-to-left scans. In this case, the antenna scans only plus or minus 60 deg. But windshear the targets are displayed only in the area plus or minus 30 deg. from the aircraft centerline. An antenna scanning is performed in 3 seconds three other seconds are used to refresh data inside the CPU.

Then, when the two ranges selected on both EFIS control panels are identical, the radar images displayed on the NDs are refreshed every 6 seconds. On the contrary, when two different ranges are selected on the EFIS control panels, the images are refreshed every 12 seconds. Windshear scan pattern The windshear processing is operating during right-to-left scans and left-toright scans. The antenna scans only plus or minus 60 deg. The detected windshear targets are displayed only in the area plus or minus 30 deg. From the aircraft centerline. When the two ranges selected on both EFIs control panels are identical, the radar images displayed on the NDs are refreshed every 6 seconds. On the contrary, when two different ranges are selected on the EFIS control panels the images are refreshed every 12 seconds. NOTE:

FOR ROLL ANGLE EXCEEDING 28 DEG., THE WINDSHEAR DETECTION IS NOT ACTIVE.

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Figure 80

WXR/PWS Windshear Indications and Scanning Logic

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WINDSHEAR ALERT Alert Levels There are three alert levels defined in function of event seriousness and distance from the aircraft. The weather radar provides the crew with visual and aural warnings which vary in function of the level detected. Windshear warning alert (level 3) This alert corresponds to the most dangerous phenomenons. It is generated for windshear events detected within +/- 0.25 NM from the longitudinal axis of the aircraft and within +/- 30 deg. scan of the aircraft heading. On the ground, the maximum range is 3 NM. In flight, the maximum range is reduced to 1.5 NM. During takeoff, level 3 covers ranges from 0 to 1.5 NM, from 50 to 1200 ft Above Ground Level (AGL). During landing, this coverage is from 1.5 to 0.5 NM, from 370 to 50 ft. Range reduction is a linear function of altitude: at 370 ft, range is equal to 1.5 NM and reaches 0.5 NM at 50 ft. During takeoff, this warning is inhibited from the time the aircraft attains 100 kts and until it reaches 50 ft AGL. Level 3 warning is inhibited below 50 ft (in approach phase) and above 1200 ft. The windshear warning alert is announced by: an aural warning message: GO AROUND WINDSHEAR AHEAD in approach or WINDSHEAR AHEAD, WINDSHEAR AHEAD at takeoff, generated by the radar synthesized voice. a visual warning: red W/S AHEAD message on the PFD. Display priority on PFD is given to level 3. The computer has to determine whether the aircraft is taking-off or landing to generate the aural warning message ”GO AROUND, WINDSHEAR AHEAD” or ”WINDSHEAR AHEAD, WINDSHEAR AHEAD”. Transition between the aural warning messages is controlled by the GEAR UP discrete input.

Windshear caution alert (level 2) This level covers the events detected in a region from 0 to 3 NM, within +/- 30 deg. of the aircraft heading but outside the windshear warning alert region (level 3). This caution alert is inhibited: during takeoff, from the time the aircraft attains 100 kts and until it reaches 50 ft AGL, during landing, below 50 ft AGL. There should be no windshear caution alert (level 2) above 1200 ft. The windshear caution alert is announced by: an aural warning: MONITOR RADAR DISPLAY a visual warning: amber W/S AHEAD message on the PFD. Windshear advisory alert (level 1) This level covers the events located within 5 NM from the aircraft, within +/- 30 deg. of the aircraft heading but outside the windshear warning and caution alert regions (levels 2 and 3). There should be no windshear advisory alert (level 1) above 1500 ft. No aural or visual warnings are provided for this advisory alert: only the windshear icon is superimposed on the radar image. The weather radar transmits the windshear alerts following their detection order. A maximum of 8 events can be transmitted. Therefore, alerts of different levels can be generated simultaneously.

ALERTS

PFD

Advisory (Level1)

ND

AURAL WARNING

windshear icon

Caution (Level 2)

W/S AHEAD (AMBER)

windshear icon

MONITOR RADAR DISPLAY

Warning (Level 3)

WINDSHEAR AHEAD (RED)

windshear icon

during takeoff: WINDSHEAR AHEAD WINDSHEAR AHEAD during landing: GO AROUND WINDSHEAR AHEAD

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Figure 81

WXR/PWS Alert Ranges and Alert Levels

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WINDSHEAR WARNING DISPLAYED Windshear flags on NDs When a windshear fault occurs, an amber PRED W/S message comes into view. The radar image remains available if this fault does not affect the radar modes or detection function. A detected fault is displayed when: the aircraft is on the ground or the flap and slat control lever is in a position different from 0. the windshear PWS/OFF/AUTO switch on the radar control unit is set to AUTO (the fault message is not displayed whent the switch is set to OFF). Warning display on Upper ECAM Display Unit A detected windshear fault is indicated by the following amber messages: NAV: PRED. W/S DET FAULT on EWD. PRED. W/S DET on SD INOP SYSTEM area. This message is associated to the indications presented on the NDs. When the PWS/OFF/AUTO switch is set to OFF on the weather radar control unit, a green or amber PRED W/S OFF memo message is presented to the crew. The color of this message depends on the flight phases.

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Figure 82

WXR/PWS Indication of Failure messages

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FAULT ISOLATION AND BITE General The BITE facilitates maintenance on in-service aircraft. It detects and identifies a failure related to the system. The BITE of the WR/PWS is situated in the radar transceiver and through two ARINC 429 low-speed buses (an input bus from the CFDIU and an output bus to the CFDIU). The BITE: transmits permanently weather radar system status and its identification message to the CFDIU, memorizes the failures which occurred during the last 63 flight legs, monitors data inputs from the various peripherals (EFIS control section, ADIRUs, RAs), transmits to the CFDIU the result of the tests performed and self-tests, can communicate with the CFDIU through the MCDU menus. acquires the general maintenance parameters (UTC, date, A/C ident...) and command codes from the CFDIU. Normal mode During the normal mode the BITE monitors cyclically the status of the WR/ PWS. It transmits its information to the CFDIU during the concerned flight. In case of fault detection the BITE stores the information in the fault memories. These items of information are transmitted to the CFDIU by an ARINC 429 message with label 356. Interactive mode The interactive mode can only be activated on the ground and through the line key adjacent to the RADAR 1 indication, presented on the SYSTEM REPORT/ TEST/NAV page of any MCDU. This mode enables communication between the CFDIU and the BITE of the weather radar transceiver by means of the MCDU. The interactive mode is composed of: LAST LEG REPORT This report contains the fault messages related to the external or internal failures (class 1 and 2) recorded during the last flight leg.

PREVIOUS LEGS REPORT This report contains the fault messages related to the external or internal failures (class 1 and 2) recorded during the previous 63 flight legs. LRU IDENTIFICATION Allows to display the P/N, the S/N and the SW/N of the equipment. GND SCANNING Based on the monitoring and fault analysis during flight, provides information of the failures detected while using this function. The WR/PWS peripheral monitoring and internal cycle tests are used to detect transient failures. The peripheral monitoring and WR/PWS internal cyclic tests are used to detect transient failures. TROUBLE SHOOTING DATA Provides correlation parameters and snapshot data concerning the failure displayed in the LAST LEG REPORT and PREVIOUS LEGS REPORT. CLASS 3 FAULTS Allows to display the class 3 faults recorded during the last flight leg. GROUND REPORT Allows to present the class 1, 2 or 3 internal failures detected on ground. These failures differ from those displayed on the LAST LEG REPORT and CLASS 3 FAULTS. By pressing the line key adjacent to the failure message, the operator is allowed to access to the corresponding Trouble Shooting Data. TEST Allows a check of the correct operation of the WR/PWS on ground. This test can be performed through the CFDS by selecting on the MCDU the test function on the RADAR 1 main menu page. At the end of the BITE TEST, the test pattern comes into view.

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Figure 83

WXR/PWS BITE Test (figure 1)

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Figure 84

WXR/PWS BITE Test (figure 2)

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Figure 85

WXR/PWS Component Location

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34-52

ATC/MODE S

DESCRIPTION General The Air Traffic Control (ATC) system is based on the replies provided by the airborne transponders in response to interrogations from the ATC secondary radar. The ground ATC secondary radar uses technics which provide the air traffic control with information that cannot be acquired by the primary radar. This system enables to distinguish between aircraft and to maintain effective ground surveillance of the air traffic. The system provides the air traffic controllers with : Mode A : transmission of aircraft identification or, Mode C : transmission of aircraft barometric altitude or, Mode S : aircraft selection and transmission of flight data for the ground surveillance. The mode S is fully compatible with the other modes, A and C. The mode S bas been designed as an evolutionary addition to the ATC system to provide the enhanced surveillance and communication capability required for air traffic control automation. NOTE:

THE ATC/MODE S WILL BE ABLE TO PROVIDE THE TRAFFIC COLLISION AVOIDANCE SYSTEM (TCAS) WITH THE AIRCRAFT ADDRESS. The interrogation frequency is 1030 MHz. The reply frequency is 1090 MHz. An airborne transponder provides coded reply signals in response to interrogation signals from the ground secondary radar and from aircraft which will be eventually equipped with the TCAS. This ground interrogation is transmitted in the form of pair of pulses P1 and P3 for the mode A or C and in the form of pulses P1, P3 and P4 for the mode S. The A320 uses two independent ATC Mode S systems. The Control Panel is common for both ATC transponders.

Control Panel The control panel is used to activate one of the two transponders and for mode switching. It transmits the ATC code and the ident pushbutton activation to the active transponder. The TCAS computer is controlled through the active transponder by the ATC/TCAS control panel. Antenna Each transponder works with a TOP and a BOTTOM antenna. The system receive signals from both antennas. Replies are transmitted to the antenna with the strongest receive level. A squitter (ATC mode S address) is transmitted alternatively by TOP and BOTTOM antenna for TCAS purpose. A suppression signal is transmitted by the ATC transponder each time when in transmission mode to inhibit other systems working in same frequency range (DME, TCAS) and to prevent simultaneous transmission. Inputs The ADIRUs provide baro altitude to the associated transponder for transmission to a ATC ground station or to other TCAS equipped aircraft. ADIRU 1 supplies ATC 1, ADIRU 2 supplies ATC 2. ADIRU 3 is a back up and will be used according to the AIR DATA switch status. The FMGC provide flight identification. This data will be transmitted to an ATC ground station after a mode S interrogation. Each LGCIU sends a discrete signal to the ATC control panel which is used for transponder activation and for internal BITE purposes. The TCAS computer transmits data to the ATC transponder to reply to a mode S interrogation and coordination messages during a coordinated resolution advisory (RA). Users The CFDIU is used to communicate with the internal BITE functions of the ATC transponders. The operative ATC transponder transmits data to the TCAS such as baro altitude, TCAS operation mode from the control panel, TCAS BROADCAST MESSAGE received, coordination messages during a corrective resolution advisory (RA) and max. A/S capability of the own aircraft. Warnings and Flags ATC FAIL indicator light (control panel) which indicates transponder failure.



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ICAO Addr. (Mode S Addr.)

ICAO Addr. (Mode S Addr.)

Figure 86

ATC System Schematic

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CONTROL 1 ATC mode of operation

4 IDENT pushbutton switch

The mode of operation of the transponder is selected by a switch with three positions STBY, AUTO, ON. STBY mode When the transponder is in standby it does not transmit either squitters or replies to ground station or other aircraft interrogations. AUTO mode In flight, the aircraft operates as in the ON mode : all its functions are active. When the aircraft touches down, the landing gear ground/flight relay disables the Mode A and C replies of the selected transponder from ground station interrogations. ON mode The Mode S transponder operates permanently, both in flight and on the ground. It periodically transmits squitters (at 1 second intervals) to be detected by other aircraft and replies to their interrogations and those from ground stations. This function permits, on ground, to override the inhibition of replies from interrogations in Mode A or C. It is used by the air traffic controller to check the correct operation of the aircraft Mode A or C transponder prior to takeoff.

On ground station request, an addition pulse must be included in the Mode A and Mode C replies transmitted by the transponder to enable a more precise location. This operation is performed by pressing the IDENT pushbutton switch on the control unit.

5 XPDR Fault Light The Fault light illuminates,if the selected transponder is fail.

6 ALT Reporting Switch The ALT RPTG switch inhibits altitude information when in the OFF position.

2 Selection of SYS 1/2 active transponder The SYS 1/2 switch permits selection of the active transponder. The non-selected transponder is placed in standby.

3 Identification code in Mode A The Mode S transponder also replies to Mode A interrogations from ground stations. Nine numeric keys permit the pilot to set the Mode A octal code assigned to the aircraft by the ATC ground station controller and included in the transmitted replies. A window on the control unit displays this code permanently as long as the content of the digital output message complies with the displayed data.

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5

6

3

3

2

4 Figure 87

ATC Control Panel

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FAULT ISOLATION AND BITE The different BITE menu selections are: LAST LEG REPORT PREVIOUS LEGS REPORT LRU IDENTIFICATION TEST Faults detected by the System and transferred to the CFDS causes the following messages displayed on the MCDU during BITE. ATC 1(2) : NO DATA FROM CONTROL SOURCE There is no correct data input from the Control Panel to the ATC Transponder. NO DATA FROM ADIRU There is no correct data input from the ADIRU1 (2) to the ATC Transponder. TRANSPONDER The ATC Transponder is faulty. NO DATA FROM CFDIU No connection to the CFDS. ANTENNA BOT The Bottom Antenna or the coaxial cable is faulty. ANTENNA TOP The Top Antenna or the coaxial cable is faulty.

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ÂÂÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂÂ ATC Control

other Systems: ADIRU 1,3 (FMGC 1)


TCAS

ATC 1

ATC 2

CFDS

other Systems: ADIRU 2,3 (FMGC 2)

CFDS monitored

ÂÂÂÂÂ ÂÂÂÂÂÂ ÂÂÂÂÂ ÂÂÂÂÂÂ ÂÂÂÂÂ ÂÂÂÂÂÂ ÂÂÂÂÂ ÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂÂ Top Antenna

ÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂÂÂ ÂÂÂ ÂÂÂÂÂ ÂÂÂÂÂÂ ÂÂÂÂÂ ÂÂÂÂÂÂ ÂÂÂÂÂ ÂÂÂÂÂÂ ÂÂÂÂÂ ÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂÂ

Bottom Antenna

Top Antenna

Figure 88

Bottom Antenna

ATC BITE Schematic

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1

1

Figure 89

ATC CFDS BITE Menu

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Figure 90

ATC CFDS BITE Test

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ACTIVATION OF THE FRONT PANEL TEST The front panel test can be activated in ground condition only by pushing the TEST pushbutton switch on the face of the transponder. During the first 3 seconds, all LEDs on the face of the transponder are on. During the next 3 seconds, all LEDs go off. During the last 3 seconds (or until the TEST pushbutton switch is released) the green TPR LED is on (except if a fault has been detected during the test). The name, color and function of the three LEDs are as follows: TPR (red) indicates that an internal fault is detected of the transponder TPR (green) indicates that no internal fault is detected of the transponder ALT (red) indicates that no altitude data input is available DATA IN (red) indicates that no control panel input is available (same as CTL) ANT TOP (red) indicates that a upper antenna or associated circuitry fault is detected (same as UPPER ANT) ANT BOT (red) indicates that a lower antenna or associated circuitry fault is detected (same as LOWER ANT ) TCAS (red) indicates that the TCAS System connection is not available (if installed) MAINTENANCE (red) indicates that the CFDIU System connection is not available.

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Figure 91

ATC Front Panel Test

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LOCATION

Figure 92

ATC Location Transponder and Antenna



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Figure 93

ATC Location Control and Indication

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34-43

TCAS

DESCRIPTION General The TCAS II (Traffic Collision Avoidance System) is a system whose function is to detect and display aircraft in the immediate vicinity and to provide the flight crew with indications to avoid these intruders by changing the flight path in the vertical plane only. The TCAS periodically interrogates their transponders, computes their trajectories and constantly determines their potential threat. When an aircraft is airborne, its TCAS periodically transmits interrogation signals for all ATC mode A/C and Mode S transponder-equipped aircraft in the vicinity. These interrogations are received by the ATC ground stations and by the transponders of the other aircraft. In response to these interrogations, the transponders of nearby aircraft return signals containing their altitude value. The TCAS computes the range between the two aircraft by measuring the elapsed time between transmission of the interrogation and reception of the reply. The altitude, altitude rate, range and range rate are determined by a periodic tracking of these exchanges and the data are used for intruder threat assessment. Visual and aural advisories are supplied by the TCAS computer whenever assessment of the relative position of two aircraft reveals a potential collision hazard. Control Panel The Control Panel is used to activate the operating and display mode of the TCAS system. This information is transmitted across the active ATC system. Antenna The TCAS system works with a TOP and a BOTTOM antenna. The antennas are used for transmission and reception. The antenna consist of four independent elements. Based on a comparison of signal phases received by the four independent elements or on a phase shifting transmission it is possible to get a directional antenna characteristics. The system activates the TOP and BOTTOM antenna alternatively. Inputs The ADIRU 1 transmits aircraft attitude and heading for ND display calculation.

The LRRAs provides radio altitude for system activation and sensitivity modulation (sensitivity level). One LRRA signal is active, the other is standby. The LGCIU 1 sends Air/Gnd and LDG extend information for TCAS mode control and for internal BITE purposes. The active ATC transponder transmits control panel data and baro altitude for TCAS mode control and sensitivity modulation. Discrete signals from the FWCs and the GPWS are used for the inhibition of certain advisories by equipment with higher priority than TCAS. This is also possible by pressing the Master Warning pushbutton. Indication All DMCs receives TCAS data to display them on Capts PFD, ND and on F/Os PFD, ND. The NDs displays the bearing and relative height of aircraft in the immediate vicinity (< 40NM). The PFDs shows the necessary V/S to reach sufficient vertical separation in case of a resolution advisory (red and green sectors). Warnings and advisories are accompanied by synthesized voice announcements via the Cockpit loudspeakers. Users The ATC mode S transponder gets TCAS status information and RA calculation data for transmission to other TCAS equipped aircraft or mode S ground stations. The FWCs receive TCAS status data to create ECAM warnings in case of failure. The CFDIU is used to communicate with the internal BITE functions of the TCAS computer (tests only available on ground). Warnings and Flags A faulty TCAS system results in the following cockpit effects: Flags on PFD and ND Master Caution Lights on glareshield Aural Warning (Single Chime) NAV TCAS FAULT on the upper ECAM display.

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Figure 94

TCAS Block Diagram

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CONTROL ATC/TCAS Control Unit Operational Use The TCAS is a cooperative system whose operating mode is very close to the ATC Mode S transponder associated to it. The main controls are thus grouped on the ATC/TCAS control unit and the traffic and conflict resolution information is presented on the EFIS displays. The manual operating modes of the TCAS are selected via the ATC/TCAS control unit. TCAS modes of operation The TCAS mode of operation is selected by means of a switch with three positions: STBY, TA, TA/RA. STBY mode In the STBY Mode, the advisory generation and surveillance functions are inhibited. No TCAS information can be displayed on the PFDs and NDs. The green TCAS STBY message is displayed in the memo section of the EWD. TA mode In this mode, intruders are displayed on the ND according to their position in the airspace. The TCAS does not generate any vertival orders. The RA type intruder symbols are converted into TA type symbols. The TA ONLY message is displayed in white on the ND at the bottom. TA/RA mode The TCAS performs all TA mode functions and also issues preventive or corrective resolution advisories, represented in the form of colored sectors along the vertical speed scale on the PFD. The sensitivity level is determined automatically in function of altitude.

TCAS modes of indication The TCAS mode of indication is selected by means of a switch with 4 positions: THRT, ALL, ABV, BLW. THRT Proximate and other intruders are displayed on the ND only, if a TA (Traffic Advisory) or RA (Resolution Advisory) is present, and they are within 2700 feet above and 2700 feet below the aircraft. ALL In this mode, all intruders are displayed on the ND according to their position in the airspace. The altitude range is -2700 feet to +2700 feet. ABV and BLW modes This selection controls the above and below vertical altitude for traffic advisory: -ABV : altitude range is set to 9900 ft above the aircraft and 2700 ft below -BL W: altitude range is set to 9900 ft below the aircraft and 2700 ft above.

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2 TCAS Display Switch 1 TCAS Mode Switch Figure 95

TCAS Control Panel

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INDICATION Normal Indication on PFD A TCAS indication appears on PFD only in case of a resolution advisory (RA). So the system must be active and switched to TA/RA mode. If a RA is initiated, red and/or green sectors are shown on the vertical speed scale of the PFD. There are two different types of RAs: Preventive resolution advisory In this case, the advisory instructs the pilot to avoid certain deviations from current vertical speed rate to avoid a risk of collision. On the PFD vertical speed scales the forbidden values are indicated by red sectors. Corrective resolution advisory In this case, the advisory instructs the pilot to change current flightpath (vertical plan only) to avoid a collision. On the vertical speed scale of the PFD, colored sectors indicate avoidance maneuvers to be performed: - red sector -> forbidden vertical speed (v/s) - green ”fly to” sector -> a v/s range to be respected

Fail Indication on PFD If the TCAS system fails, a red TCAS message (flashing 9s, the steady) comes into view to the left of the vertical speed scale on the PFD and no advisories can be displayed. In case of vertical speed fail, the RAs (red and/or green sectors) are displayed without a reference to a vertical speed scale to inform the crew about the necessary action (climb or descend).

Aural Alerts Trajectory correction or holding visual orders are accompanied by synthesized voice announcements. These announcements are generated by the TCAS computer. Preventive resolution advisory ” monitor vertical speed, monitor vertical speed ” Corrective resolution advisory depending on situation (climb or descend): ” climb, climb, climb ” ” climb, crossing climb, climb, crossing climb ” ” reduce climb, reduce climb ” ” increase climb, increase climb ” ” climb, climb now, climb, climb now ” If separation is achieved and situation is save, the following announcement is generated: ” clear of conflict ”

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Corrective RA

FLY TO Vertical Speed Sector (green)

Preventive RA

Fail Indication

FORBITTEN Vertical Speed Sector (red)

FORBITTEN Vertical Speed Sector (red)

Corrective RA with V/S Fail

FLY TO Vertical Speed Sector (green) TCAS Flag (red) FORBITTEN Vertical Speed Sector (red)

Figure 96

TCAS Indication on PFD

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Normal Indication on ND A TCAS indication on ND appears, when: the TCAS mode switch is in TA or TA/RA mode, and the ALT RPTG switch is ON, and the ATC transponder is not in STBY, and a ROSE or the ARC mode is selected on EFIS control panel. The aircrafts present in the surveillance zone are represented by symbols whose shape and color correspond to the type of intruder defined in the TCAS: Other Traffic -> white diamond (7mm) no collision threat Proximate Traffic -> white filled diamond (7mm) no collision threat; intruder in vicinity to A/C (closer than 6 NM in lateral and +- 1200 ft in vertical direction) Traffic Advisory -> amber filled circle (5mm) potential collision threat; time to intercept appr. 40 s Resolution Advisory -> red filled square (5mm) real collision threat; time to intercept appr. 25 s The symbols are positioned on the ND so as to depict their relative bearing and range. Data tags are associated with intruders. These tags consist of: two digits indicating their relative altitude in hundreds of feet a symbol indicating whether the intruder is above (+) or below (-) the aircraft. an arrow to the right of the symbol indicates the vertical trend of the aircraft (v/s > "500ft/min). These indications are only present for the 10, 20 and 40 NM range selection. If a TA or RA type intruder is detected at wrong range or mode selection, messages come into view on the ND. If the range is 20 or 10 NM, a white range ring with markings at each of the twelve clock positions is placed around the own aircraft symbol at a radius of 2.5 NM. Only the 8 most threatening intruders are displayed.

No Bearing Indication on ND Without bearing acquisition, the intruder characteristics are displayed on the bottom of the NDs (range, relative altitude, v/s tendency). Only the two most threatening intruders are displayed. The most dangerous is displayed on the left side. If the intruder becomes a potential threat, the characteristics is displayed in amber (TA) or in red (RA). Advisory messages If a TA or RA intruder is detected and the display range is greater then 40 NM, the following message comes into view at the center of the ND, in red for RA and in amber for TA: TCAS : REDUCE RANGE If a TA or RA intruder is detected and the display is in PLAN mode, the following message comes into view at the center of the ND, in red for RA and in amber for TA: TCAS : CHANGE MODE Status messages If the TCAS is in TA mode, the following message comes into view at the bottom of the ND to indicate, that no RA are possible: TA ONLY Fault Indication on ND If the TCAS system fails, a red TCAS message (flashing 9s, the steady) comes into view at the bottom of the NDs.

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Advisory Messages: TCAS: CHANGE MODE (amber or red) TCAS: REDUCE RANGE (amber or red)

Other Traffic (white)

Message Window Status Messages:

Proximate Traffic (white) Traffic Advisory (amber) Resolution Advisory (red)

Figure 97

TA ONLY (white, Crew activated) TA ONLY (amber, ATC activated) TCAS

(red)

TCAS Flag

No Bearing Indication 10.3NM+10

12.4NM-09

TCAS Indication on ND

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SENSITIVITY LEVELS The notion of sensitivity level is very important in the TCAS as many of the operating modes depend on it. The TCAS separates the surrounding airspace into altitude layers. A different Sensitivity Level (SL) threshold for issuing advisories is applied to each altitude layer. The sensitivity level is decreased at low altitude to prevent unnecessary advisories in higher traffic densities such as terminal areas. Generally, the level is determined automatically by the TCAS in function of: altitude values from the radio altimeter up to 2500 ft AGL barometric altitude values in the 2500 ft to 36,000 ft range. Time to intercept (TAU) values corresponding to each sensitivity level indicate the TA and RA thresholds. The vertical separation thresholds at closest point of approach (CPA) also vary in function of the sensitivity level for the different types of advisory. The following table summarizes these data: --------------------------------------------------------------------I I TAU thresholds I Vertical separation I I I I thresholds I -----------------------------I-----------------I---------------------I I Source I Altitude I SL TA RA I S0 S1 S2 I I Altitude I I I TA RA RA I I I I I prev cor I I I I (sec) (sec) I (ft) (ft) (ft) I I--------------------------------------------------------------------I I Radio Alt Iless than 500 I 2 20 I 1200 I I Radio Alt I 500-2500 I 4 35 20 I 1200 750 400 I I Baro I 2500-10000 I 5 40 25 I 1200 750 400 I I Baro I 10000-20000 I 6 45 30 I 1200 750 500 I I 20000-30000 I 7 45 35 I 1200 850 640 I I Baro I Baro Imore than 30000I 7 48 35 I 1200 950 750 I I--------------------------------------------------------------------I

the ATC Mode S equipped ground stations may modify the sensitivity level of the aircraft TCAS via the uplink without, however, having the capability to force the Standby Mode. If several ground stations command sensitivity levels, the TCAS logic selects the lowest level. Definition of priority logic: First a sensitivity level based on altitude is selected. Level 2 is selected if the radio altimeter altitude is less than 500 ft. Level 2 is also selected if own aircraft is configured such that both CLIMB and DESCEND RAs are inhibited (e.g., below 1000 ft AGL with insufficient climb performance). Level 4 is selected if the aircraft is above 500 ft and below 2500 ft AGL. Level 4 is the least sensitive of the levels selected automatically by the TCAS ; in fact in this altitude layer, the numerous inhibitions reduce the appearance of RA. If the aircraft is above 2500 ft AGL, barometric altitude is used to select either level 5 (below 10,000 ft), 6 (from 10,000 to 20,000 ft), and 7 (above 20,000 ft). ATC/TCAS control unit input is read by the TCAS computer. If the pilot has selected Automatic Mode (TA/RA), then the altitude-based sensitivity level will be used in comparisons to determine the final level. From all sensitivity level commands, if any, received from ground stations, the lowest is selected. If the TA ONLY mode is selected, either manually via the control unit or by a ground station, the altitude-based sensitivity level is used for TA thresholds and the RAs are inhibited. Otherwise, the lowest of all inputs is chosen.

Level 1 corresponds to Standby Mode in which no advisory is generated and Level 3 is reserved for the future TCAS III which will have the capability to generate horizontal maneuver advisories. There are two other means of modifying the sensitivity level: selecting TA only mode on the ATC/TCAS control unit forces level 2. In this case, intruders of all types are displayed but will not be transformed into RA symbols and no vertical speed modification indications will be issued.

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Figure 98

TCAS Sensitifity Levels

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ECAM WARNING In case of TCAS failure, the TCAS warning message ”NAV TCAS FAULT” is shown on the upper ECAM display, the MASTER CAUTION comes on and the single chime sounds.

FAULT ISOLATION AND BITE The different BITE menu selections are: LAST LEG REPORT PREVIOUS LEGS REPORT LRU IDENTIFICATION GND SCANNING CLASS 3 FAULTS TEST GROUND REPORT

Faults detected by the system and transferred to the CFDIU causes the following messages displayed on the MCDU during BITE. Internal Faults TCAS (1SG) The TCAS computer is faulty. TCAS TOP ANTENNA (7SG1) The TOP antenna is faulty. TCAS TOP ANTENNA (7SG1) COAXIAL JX The TOP antenna segment X or coaxial connection or cable is faulty. TCAS BOT ANTENNA (7SG2) The BOT antenna is faulty. TCAS BOT ANTENNA (7SG2) COAXIAL JX The BOT antenna segment X or coaxial connection or cable is faulty. External Faults RA X (1SAX) / TCAS (1SG) No connection to the LRRA X. ATC X (1SHX) / TCAS (1SG) No connection to the ATC X system. ADIRU1 (1FP1) / TCAS (1SG) No connection to the ADIRU 1 system. ATC-TCAS CTL PNL (3SH) / TCAS (1SG) No data from control panel. CFDIU (1TW) / TCAS (1SG) No connection to the CFDIU. POWER SUPPLY INTERRUPT A power supply interrupt has occurred.

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Capt Loudspeaker

Capt PFD

Capt ND

F/O ND

ECAM

F/O PFD

F/O Loudspeaker

Warning

EFIS System

CFDIU

ÂÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂ ÂÂÂÂÂÂ ÂÂÂÂÂÂ ÂÂÂÂÂÂ ÂÂÂÂÂÂ ÂÂÂÂÂÂ ÂÂÂÂÂÂ ÂÂÂÂÂÂ ÂÂÂÂÂÂ ÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂÂ

other Systems LRRA 1,2 ADIRU 1

other Systems FWC 1,2

TCAS

ATC 1

Top Antenna

Bottom Antenna

Audio

ATC 2

ÂÂÂ ÂÂÂ

CFDIU monitored

ATC/TCAS Control

Figure 99

TCAS BITE Schematic

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A

A

Figure 100

TCAS CFDIU Test Procedure

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Self Test A quick check of the correct operation of the TCAS installation can be performed by activating the TEST function : either by pressing the pushbutton switch on the front of the TCAS computer or through the CFDIU by applying the procedure TCAS Functional Test on the Multipurpose Control and Display Unit (MCDU). The self-test sequence checks the main functions of the computer and transmits to the displays: resolution advisory characteristics (0 ft/min advisory, up corrective advisory, don’t descend, don’t climb > 2000 ft/min, rate to maintain). four intruder data according to the following table:

Failure indication At the end of the test sequence the system generates a synthesized voice message: TCAS SYSTEM TEST OK if the system operates correctly or TCAS SYSTEM TEST FAIL if an anomaly has been detected.

------------------------------------------------------------------------I INTRUDER I TYPE I RANGE I REL ALT I BEARING I VERTICAL RATE I I I I (NM) I (FEET) I (DEG) I I ------------------------------------------------------------------------I 1 I RA I 2.00 I +200 I +90 I no vertical rate I I 2 I TA I 2.00 I -200 I -90 I climbing I I 3 I PROX I 3.625 I -1000 I +33.75 I descending I I 4 I OTHER I 3.625 I +1000 I -33.75 I no vertical rate I -------------------------------------------------------------------------

ND image The ND must display the images corresponding to the four types of intruders : Other, Proximate, TA and RA. The shapes and colors of the traffic symbols are: white outlined diamond for Other traffic white diamond for Proximate traffic amber circle for TA traffic red square for RA traffic. PFD image At the beginning of the test sequence, green and red sectors must appear sequentially on the vertical speed scale of the PFD. Then a resolution advisory display is shown.

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BITE TEST INDICATION 1

TCAS prox. traffic (white)

2

TCAS traffic adv (yellow)

3

TCAS resolution adv (red)

4

TCAS resolution adv. on V/S scale

5

TCAS flag (red)

4 5

1

2

3

5

TCAS Display Switch TCAS mode selector

TCAS

______________________________________________________________________________________________________________________________________________________________________________________________


Figure 101

TCAS BITE Indication on PFD, ND


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ACTIVATION OF THE FRONTPANEL TEST The front panel test can be activated in ground condition only by pushing the TEST pushbutton switch on the face of the TCAS computer. During the first 3 seconds, all LEDs are on. During the next 3 seconds, all LEDs turn off. During the last seconds (or until the TEST pushbutton switch is released) the TTR PASS LED is shown (except if a fault has been detected during the test). If a fault exists, the faulty component is coded as follows: -----------------------------------------------------————— LED I COMPONENT -----------------------------------------------------————— TTR FAIL I TCAS PROCESSOR FAIL XPNDR I MODE S TRANSPONDER or DATA LINK FAIL UPPER ANT I UPPER TCAS ANTENNA FAIL LOWER ANT I LOWER TCAS ANTENNA FAIL RAD ALT I RADIO ALTIMETER or DATA LINK FAIL HDNG I HEADING DATA FAIL R/A I RA DISPLAY FAIL T/A I TA DISPLAY FAIL ------------------------------------------------------—————

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Test Push Button

Figure 102

TCAS Front Panel Test

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LOCATION

Figure 103

TCAS Location Computer and Antennas

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Revision No : MK 02 21.2.96 FRA US/E Issue Date : 21/05/2013


P:

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34-48

GPWS

DESCRIPTION General The GPWS system generates aural and visual warnings, if the A/C adopts a potentially hazardous configuration (excessive descent rate or unsafe terrain clearance, below G/S) The A320 uses one GPWS systems. The system is active between 2450 ft and 10 ft for mode 1 and 3 warnings and between 2450 ft and 30 ft for mode 2, 4 and 5 warnings.

Outputs All warnings activations are monitored by the FWCs to inhibit simultaneous altitude call outs an for recording by the DFDR. When the system fails, a discrete switches on the FAULT legend on the GPWS control panel. A ECAM message appear via the SDAC. The CFDIU is used to communicate with the internal BITE functions of the GPWS computer (test only available on ground).

Control Panel Switches on the control panel are used to switch the system on and off and to control the different modes of operation.

Warnings and Flags A faulty GPWS system results in the following cockpit effects: Master Caution Lights on glareshield Aural Warning (Single Chime) NAV GPWS FAULT on the upper ECAM display Fault Light on GPWS control panel.

Inputs The radio height from LRRA1 is used for mode 1-4 warning calculation. The ADIRU1 sends data for warning profile calculation and position data for envelope modulation (profile modulation of known critical airports, runways). The ILS receiver transmits G/S deviation for mode 5 warning. The FMGC1 provides data for envelope modulation and the landing configuration 3 signal (activated via the MCDU). The SFCC sends flap position (conf 3 or full) for landing config relay. A signal from the FWCs inhibits all aural GPWS warnings as long as a stall or windshear warning is active. Pressing the GPWS-GS warning light activates the cockpit self test (AIR and RA>1000 ft or GND) or inhibits the mode 5 warning (AIR and RA<1000 ft). The audio suppression on the ECAM control panel cancels all GPWS aural warnings. The LCGIU sends a discrete signal to the GPWS which is used for internal BITE purposes. Warning Outputs The GPWS generates visual warnings through associated lights (red GPWS light for mode 1-4 and amber G/S light for mode 5) and synthetic voice through the loudspeakers.

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Figure 104

GPWS System Schematic Inputs digital

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Figure 105

GPWS System Schematic Inputs discrete

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Figure 106

GPWS System Schematic Outputs

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CONTROL 1

SYS pushbutton switch This pushbutton switch when pressed (in) (white OFF legend on), inhibits all the GPWS warnings an no GPWS self test is possible. The SYS pushbutton switch provides a FAULT warning indicating that a system fault has been detected by the GPWC. When the FAULT legend comes on, these messages are displayed on the EWD: NAV GPWS FAULT (amber) associated with action requested. GPWS .... OFF (cyan). A GPWS message (amber) is also displayed on the STATUS page on the System Display (SD).

2

G/S MODE pushbutton switch This pushbutton switch when pressed (in) (white OFF legend on), inhibits the glide slope mode.

3

FLAP MODE pushbutton switch This pushbutton switch when pressed (in) (white OFF legend on), inhibits flap abnormal condition input and generates the GPWS FLAP MODE OFF message (green) in the MEMO area of the EWD.

4

LDG FLAP 3 pushbutton switch The LDG FLAP 3 pushbutton switch when pressed (in) (white ON legend on), selects the landing flap 3 position, when released (out) , the FULL position. The GPWS FLAP message is permanently displayed (green) in the MEMO of the EWD.

5

CAPT and F/O GPWS/G/S pushbutton switches These pushbutton switches have two functions when pressed (in): - they cancel the glide slope alert or - they initiate the self test sequence if the aircraft is on ground or above 1000 ft.

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5

5

Figure 107

GPWS Contol and Indication 1

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OPERATION AND INDICATION Warning modes The serial digital data and discrete data inputs are interfaced and used in different combinations to monitor aircraft vertical performance. The following functions are monitored: Mode 1 - Excessive rate of descent Mode 2 - Excessive closure rate with terrain Mode 3 - Descent after take off and minimum terrain clearance Mode 4 - Unsafe terrain clearance Mode 5 - Descent below glide slope Warning messages Each mode computes and compares the aircraft behavior with a corresponding warning envelope. If the warning envelope is penetrated, visual and aural warnings are generated. The aural message is broadcast through the cockpit loudspeakers. The basic messages are as follows : MODE AURAL WARNING VISUAL WARNING ---------------------------------------------------1 SINK RATE GPWS 1 WHOOP WHOOP PULL UP GPWS 2 TERRAIN GPWS 2 TERRAIN TERRAIN GPWS 2 WHOOP WHOOP PULL UP GPWS 3 DON’T SINK GPWS 4 TOO LOW TERRAIN GPWS 4 TOO LOW GEAR GPWS 4 TOO LOW FLAPS GPWS 5 GLIDE SLOPE G/S Each of these warnings inhibits the automatic call out.

Inhibitions You can cancel the warnings if: You press the EMER CANC key on the ECAM control panel (aural warning only). You press the GPWS/G/S pushbutton switch for mode 5 (glide slope) visual and aural warning. This inhibition is temporary and the mode will be automatically reactivated for a new envelope penetration. You press the G/S MODE pushbutton switch for mode 5 (glide slope) visual and aural warning (permanent inhibition). You press the SYS pushbutton switch for inhibition of all the modes (visual and aural warnings). NOTE:

ALL AURAL MESSAGES ARE INHIBITED IF A STALL OR A WINDSHEAR (IF WR/PWS IS INSTALLED) IS IN PROGRESS.

Landing configuration switching To avoid the nuisance warnings during the approach, the GPWC needs to know at which flap position (FULL or 3) the crew intends to land. If the flap 3 position is selected from the control panel switch, it switches the ground signal from the SFCC 1 to the GPWC. Fault indication The faults which prevent the normal operation of the GPWC are stored in the BITE memory. They can be read on the BITE display if you operate the STATUS/HISTORY switch located on the GPWC face below the BITE display. When a class 1 fault is stored, the FAULT legend of the SYS pushbutton switch comes on.

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Figure 108

GPWS Control and Indication 2

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ECAM WARNING In case of GPWS failure, the GPWS waring message ”NAV GPWS FAULT” is shown on the upper ECAM display, the MASTER CAUTION comes on and the single chime sounds.

FAULT ISOLATION AND BITE A GPWS FAULT light shows the status of the system. The test is initiated by pressing GPWS G/S pushbutton for at least .5 sec (Ground Test) or at least 5 sec. (Ground vocabulary test). BITE Test The different BITE menu selections are: LAST LEG REPORT PREVIOUS LEGS REPORT LRU IDENTIFICATION GROUND SCANNING TROUBLE SHOOTING DATA CLASS 3 FAULTS TEST GROUND REPORT

Faults detected by the System and transferred to the CFDS causes the following messages displayed on the MCDU during BITE. GPWC (1WZ) The GWPS Computer is faulty. WRG: PIN PROG/GPWC (1WZ) Any program pins change while in flight mode. CAPT/FO GPWS GS PB SW(4WZ1/2) / GPWC(1WZ) G/S cancel > 15s, self test > 60s. ECP (2WN) / GPWC (1WZ) audio cancel > 30s. GPWS FLP MODE PB SW(7WZ) / SFCC(21VC) / GPWC(1WZ) Flap discrete signal not valid. GPWS SYS PB SW(9WZ) / GPWC(1WZ) / FWC 1/2(1WW1/2) Permanent GPWS inhibition signal. LGCIU 1 (5GA1) / GPWC (1WZ) Gear discrete signal not valid CFDIU / GPWC (1WZ) No connection to the CFDS RA1 (1SA1) / GPWC (1WZ) There is no correct data input form the LRRA system 1. ADIRU 1 (1FP1) / GPWC (1WZ) There is no correct data input form the ADIRU system 1. ILS 1 (1RT1) / GPWC (1WZ) There is no correct data input form the ILS system 1. FMGC 1 (1CA1) / GPWC (1WZ) There is no correct data input form the FMGC system 1.

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Capt Loudspeaker

ÂÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂÂÂÂÂ

F/O Loudspeaker

GPWS Control

other Systems: LRRA 1 ADIRU 1 ILS 1 FMGC 1 DMU (AIDS)

GPWC

CFDS

ÂÂÂÂÂ ÂÂÂÂÂ ÂÂÂÂÂ ÂÂÂÂÂ ÂÂÂÂÂ GPWS

other Systems: SFCC 1 LGCIU FWC 1,2 ECAM Ctrl SDAC 1,2 FWC 1,2

ÂÂÂÂÂÂ ÂÂÂÂÂÂ ÂÂÂÂÂÂ ÂÂÂÂÂÂ ÂÂÂÂÂÂ GPWS

G/S

G/S

Figure 109

GPWS BITE Schematic

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1

1

Figure 110

GPWS CFDS BITE Menu

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1

1

Figure 111

GPWS CFDS BITE Test

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Internal Built In Test Equipment (BITE) Capability The interruptive self test facility of the GPWC System provides the following test modes : airborne self test ground self test ground vocabulary test status/history test (GPWS Frontpanel). 1) Airborne self test The airborne self test is enabled when these conditions are met : the radio altitude input is greater than 1000 ft. and valid and the airspeed is greater than 90 Kts the GPWS/G/S pushbutton switch is pressed. With no system faults present, the GPWC System generates this warning sequence : A single soft GLIDE SLOPE aural warning is broadcast. A single WHOOP WHOOP PULL UP aural warning is broadcast. 2) Ground self test The ground self test presents the same test sequence as the airborne test but also includes an internal check. The internal test is enabled by a radio altitude input indicating an altitude below 5 ft. and computed airspeed below 60 kts. The test is initiated if you press and hold the TEST pushbutton switch for 0.5 seconds.

3) Ground vocabulary test The ground vocabulary test is initiated if you press the GPWS/G/S pushbutton switch continuously or during the PULL UP portion of the ground test. The test sequence is as follows : Verify that the system status is correct for the ground test (radio altitude < 30 ft., landing gear downlocked). Carry out the internal tests as detailed in the ground self test. Carry out the output sequence of the ground self test of the GPWC. Generate all aural warnings in this sequence : ORDER TRIGGER

WARNING

-----------------------------------------------------------1 Sink rate SINK RATE 2 Pull up WHOOP WHOOP PULL UP 3 Terrain TERRAIN 4 Pull up WHOOP WHOOP PULL UP 5 Don’t sink DON’T SINK 6 Too low terrain TOO LOW TERRAIN 7 Too low gear TOO LOW GEAR 8 Too low flaps TOO LOW FLAPS 9 Too low terrain TOO LOW TERRAIN 10 Glide slope GLIDE SLOPE 11 Minimums MINIMUMS MINIMUMS

NOTE:

IF YOU HOLD THE SWITCH PRESSED EITHER CONTINUOUSLY OR DURING THE PULL UP SEQUENCE OF THE GROUND TEST YOU WILL INITIATE THE VOCABULARY TEST. Upon initiation of the test, if no fault is present, the sequence will start. Between the first and second sequence steps (soft GLIDE SLOPE and WHOOP WHOOP PULL UP broadcasts) an internal check is made of the following : processor instruction set program memory contents voice memory contents The internal clock must not exceed four seconds.

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Fault Light

press button for test 1) RA > 1000 ft and TAS > 90 kts -----> Airborne self test 2) RA < 5 ft and TAS < 60 kts and t < 1s -----> Ground self test 3) RA < 30 ft and LG downlocked and t > 5s -----> Ground vocabulary test

1) +2) GLIDE SLOPE WHOOP WHOOP PULL UP 3) SINK RATE WHOOP WHOOP PULL UP TERRAIN WHOOP WHOOP PULL UP DON’T SINK# TOO LOW TERRAIN TOO LOW GEAR TOO LOW FLAPS TOO LOW TERRAIN GLIDE SLOPE MINIMUMS MINIMUMS

Figure 112

GPWS Cockpit Self Tests

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ACTIVATION OF THE FRONT PANEL TEST Status history test The test is controlled by the STATUS/HISTORY switch on the GPWC. The results are shown on an 8-character BITE display also located on the face of the GPWC. The STATUS test is initiated if you momentarily select PRESENT STATUS with the STATUS/HISTORY switch.The STATUS test will show the status of the latest flight information. The HISTORY test is initiated if you momentarily select FLIGHT HISTORY with the STATUS/HISTORY switch. The HISTORY test will depict the information related to the last ten flights. Both tests commence with ALL SEGMENTS TEST and terminate with END TEST messages. The computer failures are indicated by the message GPWC FAILED or FLT HIST INVALID. All other failure messages indicate an incorrect input condition. The desired information is read from memory and converted to alphanumeric data for presentation on the BITE display. Messages are presented by means of this vocabulary : IN TEST INVALID LAMPTEST END TEST INACTIVE AUDIO GPWS OK INHIBIT SELECT GPWC PREVIOUS FLIGHT-0 RADIO TEN FLIGHT-1 ALTIMETR FLIGHTS FLIGHT-2 BARORATE OK FLIGHT-3 AIRSPEED FAILED FLIGHT-4 ALTITUDE AIR DATA FLIGHT-5 GLIDE ILS DATA FLIGHT-6 SLOPE IRS DATA FLIGHT-7 CANCEL EXTERNAL FLIGHT-8 COURSE FLT HIST FLIGHT-9 GEAR INPUT CORRECTD RUNWAY PROCESSR LOCALIZR HEADING OUTPUT LATITUDE FLAPS ASSEMBLY LONGITUD

Examples of the STATUS test vocabulary are : GPWS OF GPWC FAILED GEAR INVALID Examples of the HISTORY test vocabulary are : RADIO ALTIMETR INACTIVE FLIGHT-1 PREVIOUS TEN FLIGHTS OK FLAPS INVALID FLIGHT-8 BARORATE INVALID FLIGHT -2 ILS DATA INACTIVE FLIGHT-3 If the history test message you read is too long, you can stop the test. To do this, you must set the STATUS/HISTORY switch to PRESENT STATUS until the BITE display is blank. The display will then show CANCEL followed by END TEST. This operation will have no adverse effect on the data held in the flight history memory. The sequences of the flight history test and those of the present status test will not activate the cockpit voice outputs.

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PRES STAT

FLT HIST

Figure 113

GPWS Status/History Test

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LOCATION

Figure 114

GPWS Location Computer

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Figure 115

GPWS Location Control and Indication

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34-48

ENHANCED GPWS

SYSTEM DESCRIPTION General The purpose of the Enhanced Ground Proximity Warning System (Enhanced GPWS) is to alert the flight crew of potentially hazardous conditions with respect to the terrain. The system achieves this objective by accepting a variety of aircraft parameters and providing the flight crew with aural alert messages and visual annunciations and displays in the event that the boundaries of any alerting envelope are exceeded. Enhanced features have been added to existing basic Ground Proximity Warning Modes 1 to 5 which are the backbone of the system. Several main alerting functional areas are integrated into the Enhanced GPWC. The functional areas are: basic Ground Proximity Warning System (GPWS) (Modes 1 to 5), Terrain Awareness and Display (TAD) function, Terrain Clearance Floor function (TCF). Basic GPWS Modes The basic GPWS modes generate aural and visual warnings if the aircraft adopts a potentially hazardous condition with respect to: Mode 1 - Excessive rate of descent. Mode 2 - Excessive closure rate with terrain. Mode 3 - Descent after takeoff and minimum terrain clearance. Mode 4 - Unsafe terrain clearance. Mode 5 - Descent below glide slope.

Enhanced Features The Enhanced GPWC includes enhanced features which complete the basic GPWS modes: Terrain Awareness alerting and Display (TAD) function - A major new feature of the Enhanced GPWS is the incorporation of the terrain awareness alerting and display functions. These functions use aircraft geographic position, aircraft altitude and a terrain data base to predict potential conflicts between the aircraft flight path and the terrain, and to provide graphic displays of the conflicting terrain. The terrain awareness alerting algorithms continuously compute terrain clearance envelopes ahead of the aircraft. If the boundaries of these envelopes conflict with terrain elevation data in the terrain database, then alerts are issued. Terrain Clearance Floor (TCF) function - The Terrain Clearance Floor (TCF) function adds an additional element of protection to the basic GPWS modes. - It creates an increasing terrain clearance envelope around the intended airport runway directly related to the distance from the runway. - TCF alerts are based on current aircraft location, nearest runway center point position and radio altitude. - TCF is active during takeoff, cruise and final approach. - This alert mode complements existing Mode 4 protection by providing an alert based on insufficient terrain clearance even when in landing configuration. - TCF function generates aural and visual alert.

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Figure 116

Enhanced GPWS

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TERRAIN AWARENESS ALERTING AND DISPLAY (TAD) Description The terrain awareness component of the Enhanced GPWS is divided into functional blocks with an interface to Navigation Display. The highlighted blocks monitor aircraft position with respect to local terrain data base and provide rapid audio and visual alerts when a terrain threat is detected. Terrain threats are recognized and annunciated when terrain violates specific computed envelope boundaries forward of the aircraft path. The terrain awareness alert lamps and audio outputs behave in the same manner as the standard GPWS mode alerts. A terrain caution alert or terrain warning alert initiates a specific audio alert phrase. Complementing the terrain threat alerts, the Enhanced GPWS also maintains a synthetic image of local terrain forward of the aircraft for display on EFIS Navigation Displays (ND). The Enhanced GPWS is configured to automatically de-select the weather display and pop-up a display of the terrain threats when they occur. The logic used provides an external input for predictive windshear alerts that can override a terrain display and revert to the weather display with the corresponding windshear data (if WR/PWS installed). The Enhanced GPWS provides two external display outputs, each with independent range-scaling control in the same fashion as a weather radar on both NDs. Changes of range scaling to one ND do not affect the other display. Each of these two independent outputs may be used to drive more than one display.

Input Processing and Signal Selection The Enhanced GPWS Input Processing and Signal Selection function conditions and formats aircraft data into proper form for use by the Enhanced GPWS while insulating the Enhanced GPWS from variations in aircraft type and configuration. Aircraft Data Inputs - Aircraft position latitude and longitude are required for terrain awareness operation and are received from the Flight Management System (FMS). The terrain threat detection and display processing are automatically disabled in some particular conditions. This is indicated to the flight crew by an ECAM memo (TERR STBY). Additionally, aircraft ground track and ground speed data are received from the IR portion of the ADIRU 1 (IRS). The aircraft altitude MSL is received from the air data portion of the ADIRU 1. Other aircraft inputs include aircraft heading (from the IR), roll attitude (from IR) and flight path angle (derived by EGPWS). Local Terrain Processing The local terrain processing block extracts and formats local topographic data and terrain features from the related data bases creating a set of digital elevation matrix overlays for use by the terrain threat detection and display processing functions. Additionally, data for the nearest runway are also extracted for use by the terrain threat detection and display processing functions. Terrain Threat Detection The terrain threat detection and display processing block performs the threat analysis on the terrain data within computed caution and warning envelope boundaries below and forward of the aircraft path. Results of these threat assessments are combined with background terrain data and data for the nearest runway and formatted into a terrain display image which can be displayed on Navigation Display in place of the weather image. In the event of terrain caution or warning conditions, a specific audio alert is triggered and the terrain display image is enhanced to highlight each of the types of terrain threats.



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Figure 117

EGPWS Terrain Awareness Functions

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TAD AUDIO ALERTS AND INDICATION Terrain Displays and Alerts The terrain awareness alerting and display function maintains a background display of local terrain forward of the aircraft for cockpit display. In the event of terrain caution or warning conditions, an aural alert and lamp outputs are triggered. The background image is then enhanced to highlight related terrain threats forward of the aircraft. NOTE:

TERRAIN IS NOT SHOWN IF MORE THAN 2000 FT BELOW REFERENCE ALTITUDE AND / OR TERRAIN IS NOT SHOWN IF TERRAIN ELEVATION IS WITHIN 400 FT OF RUNWAY ELEVATION NEAREST THE AIRCRAFT.

Threat

Color

Warning terrain (approx. 30 sec. from impact); audio alert TERRAIN AHEAD, PULL UP Caution terrain (approx. 60 sec. from impact); audio alert: TERRAIN AHEAD

Solid Red Solid Yellow

Terrain that is more than 2000 ft. above aircraft altitude Terrain that is between 1000 and 2000 ft. above aircraft altitude

High Density Red High Density Yellow Medium Density Yellow

Terrain that is 500 (250 with gear down) ft. below to 1000 ft. above aircraft altitude

Medium Density Green

Terrain that is 500 (250 with gear down) ft. below to 1000 ft. below aircraft altitude

Black

Terrain that is 1000 to 2000 ft. below aircraft altitude No close terrain

Light Density Magenta

Unknown terrain

Light Density Green

Terrain Caution Alert A specific audio alert and light output is triggered and the background image is enhanced to highlight the terrain caution threats. At the start of a terrain caution alert, the terrain awareness function triggers the caution audio alert phrase TERRAIN AHEAD. The phrase is repeated after seven seconds if still within the terrain caution envelope. During a terrain caution alert, the GPWS legend of pushbutton switches is on. During a terrain caution alert, areas where terrain violates the terrain caution envelope along the aircraft track, and within plus or minus 90 deg. of the aircraft track, are painted with the caution color 100 per cent yellow. Terrain Warning Alert When the conditions have been met to generate a terrain warning alert, a specific audio alert and light output is triggered and the background image is enhanced to highlight the terrain caution and warning threats. At the start of a terrain warning alert, the terrain awareness function triggers the warning audio alert phrase TERRAIN AHEAD, PULL UP. The phrase is repeated continuously while within the terrain warning envelope. During a terrain warning alert, the GPWS legend of pushbutton switches is on. During a terrain warning alert, areas where terrain violates the terrain warning envelope along the aircraft track, and within plus or minus 90 deg. of the aircraft track, are painted with the warning color 100 per cent red. NOTE:

-WHEN AN ALERT OCCURS (CAUTION OR WARNING) AND THE FCU MODE IS NOT IN A CORRECT MODE (ARC OR ROSE), THE MESSAGE TERR. CHANGE MODE IS DISPLAYED ON ND’S. -WHEN AN ALERT OCCURS (CAUTION OR WARNING) AND THE FCU RANGE SELECTED IS 160 OR 320NM THE MESSAGE TERR. REDUCE RANGE IS DISPLAYED ON ND’S.

TAD inhibitions manually by the GPWS/TERR P.B. switch automatically when the FMS aircraft position accuracy is not accurate enough. This is indicated to the crew by the automatic deselection of terrain display and the illumation of the TERR STBY memo on the ECAM display unit.

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TERR AHEAD

-message TERR AHEAD in RED: Warning Terrain -message TERR AHEAD in AMBER: Caution Terrain -message TERR in CYAN: normal indication

Figure 118

EGPWS Terrain Indication on ND

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TERRAIN CLEARANCE FLOOR (TCF) Description The Terrain Clearance Floor (TCF) alert function adds an additional element of protection to the standard GPWS. It creates an increasing terrain clearance envelope around the airport runway to provide protection against Controlled Flight Into Terrain (CFIT) situations beyond that which is currently provided. TCF alerts are based on current aircraft location, nearest runway center point position and radio altitude. TCF is active during takeoff, cruise and final approach. This alert mode complements the existing Mode 4 protection by providing an alert based on insufficient terrain clearance even when in landing configuration. Alerts for TCF illuminate GPWS cockpit lamps and produce aural messages. System Inputs Input

Source

Radio Altitude

External: Radio Altimeter

Latitude

External: FMS

Longitude

External: FMS

Runway Center Latitude

Internal: data base

Runway Center Longitude

Internal: data base

Navigation Mode

FMS

Alert Envelope Parameters

Internal: data base

1/2 Runway Length

Internal: data base

System Error Factor

Internal: data base

System Outputs When an aircraft penetrates the TCF alert envelope the following aural message occurs: TOO LOW TERRAIN. This aural message occurs once when initial envelope penetration occurs, and one time thereafter for each 20 per cent degradation in radio altitude. At the same time the GPWS legend of pushbutton switches comes on.

Runway Data Base The TCF runway data base consists of data records containing the position of airport runway center points along with 1/2 the runway length. The data base includes all hard surface runways in the world greater than or equal to 3500 ft in length. The process of generating this data base is certified and includes an end check that validates that the data was not corrupted in the translation process. This data base can be updated without affecting the customer certified system part number. The design of the data base and related software is such that additional runway records can be added in the future without altering the code. The data base provides a means of accessing the runway record of the runway closest to the current aircraft position. Alert Envelope The TCF alert envelope is a circular band centered over the nearest runway. The distance from the runway center to the inner envelope edge is equal to 1/2 the runway length plus the envelope bias factor. Thus the inner and outer radius of the envelope are modulated based on the runway length and envelope bias factor. Runway length varies from one runway to the next, and the envelope bias factor is typically 1/2 NM to 2 NM and varies with position accuracy. The outer alert envelope boundary extends to infinity, or until it meets the outer alert envelope boundary of another runway. The alert is inhibited below radio altitude of 30 ft. TCF inhibitions manually by the GPWS/TERR P.B. switch automatically when the FMS aircraft position accuracy is not accurate enough. This is indicated to the crew by the automatic deselection of terrain display and the illumation of the TERR STBY memo on the ECAM display unit.

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Figure 119

EGPWS Terrain Clearance Floor alert envelope

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INPUTS OUTPUTS EGPWC

EGPWC Digital Data Inputs The Enhanced GPWC receives serial digital data inputs from: radio altimeter transceiver 1 (radio altitude), Air Data/Inertial Reference Unit 1 (ADIRU), ADR portion (barom. altitude, barom. altitude rate, computed airspeed), IR portion (latitude, longitude, magnetic heading), IR portion (latitude, longitude, magnetic heading), ILS receiver 1 (glide slope dev., localizer dev., selected runway heading), FMGC (latitude, longitude, track, navigation modes), CFDIU (command word, date, flight number, UTC), FCU 1 and 2 (CAPT and F/O ranges), Weather Radar 1 (hazard bus). Discrete Data Inputs Discrete data inputs are received from the following: Slat Flap Control Computer 1 (SFCC) (3 and FULL flap position), Flight Warning Computer 1 and 2 (FWC) (all audio inhibition), Landing Gear Control and Interface Unit (LGCIU), ECAM control panel (audio suppression), GPWS/FLAP MODE pushbutton switch which, when pressed (in) (white OFF legend on), overrides a flap abnormal condition input, GPWS/SYS pushbutton switch which, when pressed (in) (white OFF legend on), inhibits Modes 1 to 5 warnings, GPWS/G/S MODE pushbutton switch which, when pressed (in) (white OFF legend on), overrides the glide slope mode, GPWS/G/S pushbutton switch which, when pressed (in), enables the Enhanced GPWC to perform test, GPWS/TERR pusbutton switch which, when pressed (in) (white OFF legend on), inhibits TAD and TCF functions, TERR ON ND (CAPT or F/O) pushbutton switches allow the crew to select or deselect terrain display on ND’s, Weather Radar control unit.

Warning Outputs Two discrete outputs from the EGPWC control the GPWS/G/S pushbutton switches located on Captain and First Officer main instrument panels. - The upper legend identified GPWS, controlled by the first output, comes on red when a ground proximity warning is generated by the EGPWC for Modes 1 to 4 or TAD and TCF warnings. - The lower legend identified G/S, controlled by the second output, comes on amber when a glide slope (Mode 5) caution alert is generated by the Enhanced GPWC. The pushbutton switch provides a facility to cancel a glide slope warning, if in progress, or to initiate an Enhanced GPWS self-test. Both discrete outputs are also used to inhibit TCAS and automatic call out when the GPWS or G/S warnings are in progress. Both discrete outputs are also used for the Digital Flight Data Recorder (DFDR). Monitor Outputs There are two monitor outputs: GPWS monitor output controls the FAULT legend of the SYS pushbutton switch and indicates a failure of Modes 1 to 5, TERR monitor output controls the FAULT legend of the TERR pushbutton switch and indicates a failure of TAD and TCF functions. Bus Output The bus output is used by the Aircraft Integrated Data System (AIDS), Data Management Unit (DMU) and by the Centralized Fault Display Interface Unit (CFDIU) for test causes. Audio Output The audio output is used by the cockpit loud speakers for aural warning messages.

POWER SUPPLY The Enhanced GPWC power supply circuits receive 115VAC, 400 Hz, single phase (22 W max.) supply from the AC Bus1.

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Figure 120

EGPWC Inputs and Outputs

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CONTROL Pushbutton Switches

1

4

GPWS/TERR pushbutton switch

When this pushbutton switch is pressed (white OFF legend on), the TAD and TCF functions are inhibited (visual display and audio inhibition). provides a FAULT warning indicating that a failure of TAD and / or TCF functions has been detected by the Enhanced GPWC. When the FAULT legend comes on, the following messages are displayed: - on the upper ECAM display unit if they are not inhibited by the FWC: NAV - GPWS TERR DET FAULT (amber) GPWS TERR.............OFF (cyan) - on the STATUS page (INOP SYS) of the lower ECAM display unit: GPWS TERR (amber)

2

GPWS/SYS pushbutton switch

When this pushbutton switch is pressed (white OFF legend on), all ground proximity alerts (Mode 1 to 5) are inhibited (visual and audio) and no Enhanced GPWC self-test is possible. provides a FAULT warning indicating that a failure in Modes 1 to 5 has been detected by the Enhanced GPWC.When the FAULT legend comes on, the following messages are displayed: - on the upper ECAM display unit if they are not inhibited by the FWC: NAV - GPWS FAULT (amber) GPWS ........OFF (cyan) (associated with action requested) - on the STATUS page (INOP SYS) of the lower ECAM display unit: GPWS (amber)

3

GPWS/FLAP MODE pushbutton switch

This pushbutton switch, when pressed in (white OFF legend on), overrides flap abnormal condition input and generates the GPWS FLAP MODE OFF message (green) in the memo area of the upper ECAM display unit.

5

GPWS/LDG FLAP 3 pushbutton switch

To avoid nuisance warnings during approach, the Enhanced GPWC needs to know at which flap position the crew intends to land. When pressed in (white ON legend on), indicates to the Enhanced GPWC that the pilot intends to land in flap 3position. When released out, the pushbutton switch indicates to the Enhanced GPWC that the pilot intends to land in flap FULL position. The GPWS FLAP message is permanently displayed in green on the MEMO of the ECAM display unit if no warning is in progress.

6

CAPT and F/O TERR ON ND pushbutton switches

These pushbutton switches allow the crew to select or deselect terrain display on ND. The ON legends indicate that terrain data is displayed on ND (following manual or automatic pop up selection).

7

CAPT and F/O GPWS/G/S pushbutton switches

These pushbutton switches, located on panels 301VU and 500VU, have two functions when pressed in: - they cancel the glide slope alert, or - they initiate the self-test sequence if the aircraft is on ground or above 2000 ft Above Ground Level (AGL)

GPWS/G/S MODE pushbutton switch

This pushbutton switch, when pressed in (white OFF legend on), inhibits the glide slope mode.

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7

6

Figure 121

EGPWS Control and Indication

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ALERTS AND INHIBITITIONS Aural alert messages Each mode computes and compares aircraft behaviour with a corresponding alert envelope. If the alert envelope is penetrated, visual and aural alerts are generated. The aural message is broadcast through the cockpit loud speakers. The messages are as follows: MODE

AURAL ALERTS

VISUAL ALERTS

1

SINK RATE

GPWS

1

PULL UP

GPWS

2

TERRAIN

GPWS

2

TERRAIN TERRAIN

GPWS

2

PULL UP

GPWS

3

DON’T SINK

GPWS

4

TOO LOW TERRAIN

GPWS

4

TOO LOW GEAR

GPWS

4

TOO LOW FLAPS

GPWS

5

GLIDE SLOPE

G/S

TAD

TERRAIN AHEAD

GPWS

TAD

TERRAIN AHEAD PULL UP

GPWS

TCF

TOO LOW TERRAIN

GPWS

Each of these alerts inhibits the automatic call out.

Inhibitions Alerts may be cancelled by: Pressing the EMER CANC key on the ECAM control panel (aural alert only). Pressing the GPWS/G/S pushbutton switch on the main instrument panel for Mode 5 (glide slope) visual and aural alert. This inhibition is temporary and the mode is automatically reactivated for a new envelope penetration. Pressing the G/S MODE pushbutton switch on the overhead panel for Mode 5 (glide slope) visual and aural alert (permanent inhibition). Pressing the SYS pushbutton switch on the overhead panel for inhibition of GPWS Modes 1 to 5 (visual and aural alerts). Pressing the TERR pushbutton switch on the overhead panel for inhibition of TAD and TCF functions (visual and aural alerts). NOTE:

ALL AURAL MESSAGES ARE INHIBITED IF A STALL OR A WINDSHEAR (IF WR/PWS IS INSTALLED) IS IN PROGRESS.

ECAM WARNING Failure in Mode 1 to 5 upper ECAM display unit: NAV - GPWS FAULT (amber) GPWS ........OFF (cyan) lower ECAM display unit (INOP SYS): GPWS (amber) Failure of TAD and / or TCF functions upper ECAM display: NAV - GPWS TERR DET FAULT (amber) GPWS TERR.............OFF (cyan) lower ECAM display unit (INOP SYS): GPWS TERR (amber)

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Figure 122

EGPWS Messages on upper and lower ECAM-DU

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SELF-TESTS Description On the ground only, the Enhanced GPWC provides self-test capability, providing an indication of the ability of the Enhanced GPWC to perform its intended function. The Enhanced GPWC self-test is initiated by momentarily pressing GPWS/G/S pushbutton switches or activated via the MCDU on the ground. When activated by pressing the GPWS/G/S pushbutton switches, the self-test is enunciated.This self-test can also be accessed via the headphone jack on the front panel of the Enhanced GPWC and it has been divided into six different levels to help with Enhanced GPWC testing and troubleshooting. Level 1 Level 1, functional testing, provides an overview of the current operational functions selected and provides an indication of their operational status. A long level 1 self-test sequence is initiated when the GPWS/G/S pushbutton switch is not released while self-test voices start. Level 1 self-test sequence: the FAULT legend of the GPWS/SYS pushbutton switch comes on. the FAULT legend of the GPWS/TERR pushbutton switch is on, G/S legends of both GPWS/G/S pushbutton switches are on, the GLIDESLOPE aural warning operates then stops. G/S legends of both GPWS/G/S pushbutton switches go off, GPWS legends of both GPWS/G/S pushbutton switches are on, the PULL UP aural warning operates then stops. the TERRAIN AHEAD PULL UP aural warning operates then stops. GPWS legends of both GPWS/G/S pushbutton switches go off, ON legends of both TERR ON ND pushbutton switches are on, terrain self-test pattern and TERR TST amber message is displayed on both NDs, the above warnings operate, then all these warnings, one after the other: GLIDE SLOPE, PULL UP, TERRAIN AHEAD, PULL UP, SINK RATE, PULL UP, TERRAIN, PULL UP, DON’T SINK, DON’T SINK, TOO LOW TERRAIN, TOO LOW GEAR, TOO LOW FLAPS, TOO LOW TERRAIN, GLIDE SLOPE,

TOO LOW TERRAIN, TERRAIN AHEAD, TERRAIN AHEAD, TERRAIN AHEAD PULL UP terrain self-test pattern turns off, ON legends of both TERR ON ND pushbutton switches and the images displayed on both NDs revert to the configuration selected before the test, the FAULT legend of the GPWS/SYS pushbutton switch goes off, the FAULT legend of the GPWS/TERR pushbutton switch goes off. Level 2 Level 2, current faults, provides a listing of the internal and external faults currently detected by the Enhanced GPWC. Level 2 self-test is initiated by pressing the GPWS/G/S pushbutton switch within 3 seconds after the end of level 1 self-test. Level 3 Level 3, Enhanced GPWS configuration, indicates the current configuration by listing the current hardware, software, data bases and program pin inputs detected by the Enhanced GPWC. This level is initiated by pressing the GPWS/G/S pushbutton switch when PRESS TO CONTINUE message is enunciated. Level 4 Level 4, fault history, provides a historical record of both internal and external faults detected by the Enhanced GPWC. This level is initiated by pressing the GPWS/G/S pushbutton switch when PRESS TO CONTINUE message is enunciated. Level 5 Level 5, warning history, provides a historical record of the warnings and cautions given by the Enhanced GPWC. This level is initiated by pressing the GPWS/G/S pushbutton switch when PRESS TO CONTINUE message is enunciated. Level 6 Level 6, discrete test, provides annunciation of discrete input transitions to be used for maintenance support. This level is initiated by pressing the GPWS/ G/S pushbutton switch when PRESS TO CONTINUE message is enunciated.

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Figure 123

EGPWS Test Pattern

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FAULT ISOLATION AND BITE The BITE: continuously transmits Enhanced GPWS status and its identification message to the CFDIU, memorizes the faults which occurred during the last 63 flight segments, monitors data inputs from the various peripherals (FMGC, RA transceiver, ILS receiver, ADIRU, SFCC, LGCIU, ECAM control panel and CFDIU), transmits to the CFDIU the result of the tests performed, can communicate with the CFDIU through the menus.

CLASS

MESSAGE

CLASS

ATA

GPWS FLP MODE PB SW (7WZ)/SFCC1(21CV)/GPWC(1WZ)

1

34-48-08

LGCIU1 (5GA1)/GPWC (1WZ)

1

32-31-71

CFDIU (1TW)/GPWC (1WZ)

3

31-32-34

RA1 (2SA1)/GPWC (1WZ)

1

34-42-33

ADIRU1 (1FP1)/GPWC (1WZ)

1

34-12-34

ADIRU1 (1FP1)/GPWC (1WZ)

3

34-12-34

ATA

ILS1 (2RT1)/GPWC (1WZ)

1

34-36-31

Internal Failures: MESSAGE

External Failure:

GPWC (1WZ)

1

34-48-34

ILS1 (2RT1)/GPWC (1WZ)

3

34-36-31

GPWC (1WZ)

3

34-48-34

FMGC1 (1CA1)/GPWC (1WZ)

1

22-83-34

WRG:PIN PROG/GPWC (1WZ)

1

34-48-34

FMGC1 (1CA1)/GPWC (1WZ)

3

22-83-34

CAPT/FO GPWS GS PB SW(4WZ/5WZ)/GPWC(1WZ) 1

34-48-34

FCU (3CA) BUS CP-L/GPWC (1WZ)

1

22-81-12

FO GPWS GS PB SW (5WZ)/GPWC(1WZ)

1

34-48-08

FCU (3CA) BUS CP-R/GPWC (1WZ)

1

22-81-12

ECP(6WT)/GPWC(1WZ)

1

31-61-12

WXR1 (1SQ1) BUS HAZARD/GPWC (1WZ)

1

34-41-33

GPWS SYS PB SW (9WZ)/GPWC (1WZ)

1

34-48-08

WXR1 (1SQ1)/GPWC (1WZ)

1

34-41-33

GPWS TERR PB SW (31WZ)/GPWC(1WZ)

1

34-48-08

GPWC(1WZ)/FWC1/2(1WW1/2)/WXR1/2(1SQ1/2)

1

34-48-34

CAPT TERR ON ND PB SW (30WZ1)/GPWC (1WZ)

1

34-48-08

FO TERR ON ND PB SW (30WZ2)/GPWC (1WZ)

1

34-48-08

Installation of EGPWC After installation of EGPWC do the correct test and perform the loading of the Enhanced GPWC Data Base by using the PCMCIA interface.

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Capt Loudspeaker

F/O Loudspeaker

ÂÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂ E/GPWS Control

ÂÂÂÂÂÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂÂÂÂÂ

other Systems: LRRA1; ADIRU1; ILS1; FMGC1; FCU1,2; WX1

DMU (AIDS)

EGPWC

ÂÂÂÂÂ ÂÂÂÂÂ ÂÂÂÂÂÂ ÂÂÂÂÂ ÂÂÂÂÂÂ ÂÂÂÂÂ ÂÂÂÂÂÂÂÂÂÂÂ ÂÂÂÂÂÂ ÂÂÂÂÂÂ GPWS G/S

TERR ON ND o

o

ON

Figure 124

other Systems: SFCC 1; LGCIU1 FWC1,2; ECAM Ctrl; WX1 SDAC 1,2 FWC 1,2

CFDS

ÂÂÂÂÂÂ ÂÂÂÂÂÂ ÂÂÂÂÂÂ ÂÂÂÂÂÂ ÂÂÂÂÂÂ ÂÂÂÂÂÂ ÂÂÂÂÂÂ ÂÂÂÂÂÂ ÂÂÂÂÂÂ ÂÂÂÂÂÂ TERR ON ND o

o

GPWS G/S

ON

EGPWS BITE Schematic

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CFDS BITE MENU Mode The menu mode can only be activated on the ground. This mode enables communication between the CFDIU and the Enhanced GPWC BITE by means of the MCDU. The Enhanced GPWS menu mode is composed of: - LAST LEG REPORT - PREVIOUS LEGS REPORT - LRU IDENTIFICATION - GROUND SCANNING - TROUBLE SHOOTING DATA - CLASS 3 FAULTS - GROUND REPORT - TEST

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Figure 125

EGPWS CFDS BITE

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Figure 126

EGPWS Component Location Cockpit

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Figure 127

EGPWC Location

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Training Manual A 319/320/321 ATA 46 Information Systems 46-21 ATIMS

Level 3 B12E B2E

Lufthansa Issue: July 2000 Technical Training GmbH For Training Purposes Only Book No: A320 46 L3 E Lufthansa Base Lufthansa 1995 ______________________________________________________________________________________________________________________________________________________________________________________________



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ATA 46

INFORMATION SYSTEM

46-21

ATIMS

HAM US/F JC 01.07.2000 ______________________________________________________________________________________________________________________________________________________________________________________________ Revision No : 02 Issue Date : 21/05/2013


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GENERAL Up to now, flight crews have communicated with air traffic controllers using HF and VHF radio communications which are subject to atmospheric disturbances and so, often difficult to understand. Furthermore, the transmission networks become saturated due to the air traffic increase, and to the limited capability to exchange complex data (routes, weather information...). Consequently, the Air Traffic and Information Management System (ATIMS) has been developed to enable datalink communications and the exchange of complex data or specific reports between the aircraft and the ground centers: controller-pilot datalink communications (HF voice in backup) for air traffic management, automatic reporting (position, intention) for air traffic surveillance, specific airline-aircraft communications (operational control) to improve airline operational costs and flexibility.



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CNS

Figure 1

ATIMS - Concept / Architecture

______________________________________________________________________________________________________________________________________________________________________________________________



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SYSTEM GENERAL DESCRIPTION System Configuration The Air Traffic and Information Management System is organized around a host platform which integrates datalink applications and the routing function. The ATIMS system is configurated in Pre-FANS configuration: Air/ground communication Router Function (ARF) Airline Operational Control applications (AOC). NOTE:The Datalink Control and Display Units (DCDU) are not fitted and the ATC MSG pushbuttons switches are not operational in Pre-FANS configuration. In addition, the FANS A configuration contains the ATSU FANS A applications (Air Traffic Control) with the Datalink Control and Display Units (DCDU) and the ATC MSG pushbuttons switches operational. System Architecture The ATIMS is composed of: an Air Traffic Service Unit (ATSU), two Datalink Control and Display Units (DCDUs), two ATC MSG illuminated pushbuttons switches. NOTE:The two DCDUs are not fitted in Pre-FANS configuration and the ATC MSG pushbuttons switches are not operational. The ATSU is connected to the following units and uses the services of these multipurpose devices for interface needs: the Multipurpose Control and Display Units (MCDUs) the Printer the Flight Warning Computers (FWCs) the Radio Management Panels (RMPs).



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Figure 2

ATIMS - General Schematic

______________________________________________________________________________________________________________________________________________________________________________________________



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LOCATIONS ------------------------------------------------------------------------------FIN | FUNCTIONAL DESIGNATION | PANEL|ZONE|ACCESS | ATA | | | | DOOR | REF. -------------------------------------------------------------------------------

1TX1 4TX1 4TX2 20TX 22TX 24TX

ATSU P/BSW-ATC MSG, CAPT P/BSW-ATC MSG, F/O SOFTWARE-ATSU A/C INTERFACE SOFTWARE-AOC SOFTWARE-AOC DATABASE

80VU 122 131VU 211 130VU 212 211 211 211

811 831 831 831 831 831

46-21-34 46-21-00 46-21-00 46-21-00 46-21-00 46-21-00

Figure 3

ATIMS - Locations



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Figure 4

ATIMS - Locations

______________________________________________________________________________________________________________________________________________________________________________________________



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COMPONENT DESCRIPTION ATSU Description The ATSU is the main component of the system.

- ATSU A/C interface software

Its architecture is based on: an Air Traffic Service Unit (ATSU) hardware case an ATSU A/C INTERFACE software uploaded in the ATSU through the Multipurpose Disk Drive Unit (MDDU) or the portable data loader an AOC software uploaded in the ATSU through the MDDU or the portable data loader an ATSU FANS A APPLICATIONS software uploaded in the ATSU through the MDDU or the portable data loader (only in FANS A configuration). The main functions performed by the ATSU are: to host the various datalink applications, including Airline Operational Control and Air Traffic Services, to provide management and access to the different datalink services available, to provide management and access to the various datalink networks available.

Its

different functions are : monitoring of the system (power supply and BITE functions) acquisition of the aircraft parameters for applications software use management of the air/ground communications (ARF function) management of the communication with the on-board peripheral units management of the human/machine interface (MCDU, DCDU, Printer and alert function) - AOC software The AOC software consists in hosted AOC applications which are depending on airline definition. These datalink applications concern operations related to the flight such as flight plans, weather, behaviour of aircraft elements transmitted for maintenance reasons, fuel quantity, personnel management, gate management... - ATSU FANS A applications software The different Air Traffic Control (ATC) applications contained in this package are: (a)ATS Facilities Notification (AFN) application The purpose of this application is to establish the contact with the ATC ground center, then to provide the ATC center with the aircraft registration, the data-link applications available on the aircraft with the corresponding addresses. (b)Controller-Pilot Data Link Communications (CPDLC) application The aim of this application is to provide dialog between ground controllers and flight crews, using datalink instead of voice communications. Each CPDLC message is composed of a set of message elements which correspond to the existing phraseology used by current ATC procedures. (c)Automatic Dependent Surveillance (ADS) application The ADS function is to provide the ATC ground center with aircraft surveillance data through periodic, event or on-demand reports.



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Figure 5

ATSU

______________________________________________________________________________________________________________________________________________________________________________________________



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DCDU Description Datalink Control and Display Unit (DCDU) (in FANS A configuration only) The DCDUs are the interface means dedicated to the ATC applications. They provide the flight crew with display capabilities and control means, allowing the display of messages received from ATC ground center and the sending of answer and messages to the ground center. The DCDUs are managed by the ATSU which processes and organizes the data in screen pages to be displayed and translates received key codes into crew orders (soft keys). Physical Description The DCDU is equipped with a LCD screen and twice four push-buttons located on each side of the screen. Electrical Characteristics power supply: 28VDC dissipated power : 16.5 Watts (average value). one back connector composed of: - one ARINC 429 input bus from the ATSU. - one ARINC 429 output bus to the ATSU. - eleven discrete inputs (P- type) for pin programming, power supply, ground and air/ground signal. - one analog input.

Functional Operation The DCDU has two main functions: Display function: display of the air traffic system information to the flight crew The DCDU is equipped with an LCD flat panel. The DCDU displays the messages formatted by the ATSU on a black background, in eight different colors: amber, black, cyan, green, magenta, red, white and yellow. These messages are in semi-graphical format and include alphanumerical text and simple graphical attributes such as boxes, arrows, separation lines, ”inverse video”... Control function: a response device for the flight crew. The DCDU has: - four pushbuttons switches associated to menu keys named ”soft keys” - four engraved pushbuttons switches dedicated to: page up and down functions. message up and down functions . print function . manual brightness control. In addition, the pushbuttons switches are lighted for night vision in accordance with the general cockpit selection. Each DCDU has a ”black screen” function in order not to disturb the flight crew in case of abnormal display. Any action on a pushbutton switch is transmitted to the ATSU. Any action on a menu key is acknowledged by the DCDU itself on the display, prior to and independently of a possible message change from the ATSU, by a reverse video display. The time elapsed between pilot action and both sending of the data on the ARINC bus and local feedback displayed on the screen is not greater than 100 ms. ATC MSG Illuminated Pushbutton Switches Description These pushbutton switches provide the flight crew with a visual alert in case of ATC message reception and with an alert cancel means by pressing them. In addition, an aural alert is triggered by the FWCs, single or repetitive chime sounds according to the priority of the message.



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Figure 6

DCDU

______________________________________________________________________________________________________________________________________________________________________________________________



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MCDU USE The MCDU gives access to these three independent menus: for air/ground communication management functions: - configuration/initialization. - VHF3 control. - communication statistics display. - test/audit mode control for AOC hosted applications: - configuration/initialization. - downlink message entry/selection/transmission. - uplink message display for ATC applications in FANS A configuration:. - preparation and modificaton of all the messages initiated by the crew, - justifications to negative replies or to a particular request, - editing of text, - configuration of applications (ADS activation, AFN initialization) and - configuration of systems (ATSU management, automatic or on-request printing).



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Figure 7

MCDU MENU

______________________________________________________________________________________________________________________________________________________________________________________________



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RMP USE Each of the three RMPs is an interface device for the VDR3 operation. The frequency range is from 118000 to 136975 KHz by 25 KHz or 8.33KHz steps. Each RMP enables the crew to ask for a switching between the RMP and the ATSU to control the VDR3 frequency: When the RMP controls the VDR3 frequency, only the Voice mode is available and the selection of the VDR3 frequency is done through the RMP by displaying the frequency in the ACTIVE window. When the ATSU controls the VDR3 frequency, the Voice mode can be accessed through the ATSU menu on the MCDU and the selection of the VDR3 frequency is done on the MCDU either in Data mode or in Voice mode. NOTE: The RMP sends to the ATSU the pilot request of switching the system controlling the VDR3 between the RMP and the ATSU. In return, the ATSU indicates to the RMP which system, between the RMP and the ATSU, is controlling the VDR3 frequency.

______________________________________________________________________________________________________________________________________________________________________________________________

Revision No : 02 US/F JC HAM Issue Date : 21/05/2013

01.07.2000


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Figure 8

ATIMS - RMP

______________________________________________________________________________________________________________________________________________________________________________________________



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POWER SUPPLY The ATIMS system is supplied with this(these) following circuit breaker(s): PANEL DESIGNATION FIN LOCATION 121VU ATSU 1 3TX1 L16 121VU ATSU 1/SWTG 5TX1 L15 ATSU The ATSU is supplied with 115VAC from the main 115VAC BUS1 bar 101XP-C via circuit breaker 3TX1 and 28VDC from the main 28VDC BUS1 bar 101PP via circuit breaker 5TX1. DCDU The DCDU1 is supplied with 28VDC from the main 28VDC BUS1 bar 103PP via circuit breaker 6TX1. The DCDU2 is supplied with 28VDC from the main 28VDC BUS1 bar 202PP via circuit breaker 6TX2.



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Figure 9

ATIMS - Power Supply

______________________________________________________________________________________________________________________________________________________________________________________________



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INTERFACE General The ATSU uses the data transparent protocol, defined in ARINC 429 Specification, when it communicates with the on-board avionics systems.

The ATSU is interfaced with the following peripheral units: the Flight Management and Guidance Computer 1 and 2 (FMGC) (Ref. AMM 22-83-00) the Multipurpose Control and Display Unit 1 and 2 (MCDU) (Ref. AMM 22-82-00) the VHF Data Radio 3 (VDR3) transceiver (Ref. AMM 23-12-00) the Satellite Data Unit (SDU) (Ref. AMM 23-28-00) the Cabin Terminals: the Cabin Management System (Ref. AMM 23-74-00) and the Digital Interface Unit (Ref. AMM 23-34-00) the Radio Management Panels 1, 2 and 3 (RMP) (Ref. AMM 23-81-00) the Clock (Ref. AMM 31-21-00) the Data Management Unit (DMU) (Ref. AMM 31-36-00) the Flight Warning Computer 1 and 2 (FWC) (Ref. AMM 31-52-00) the System Data Acquisition Concentrator 1 and 2 (SDAC) (Ref. AMM 31-54-00) the Display Management Computer 1, 2 and 3 (DMC) (Ref. AMM 31-62-00) the Centralized Fault Display Interface Unit (CFDIU) (Ref. AMM 31-32-00) the Multipurpose Disk Drive Unit (MDDU) (Ref. AMM 31-38-00) the printer (Ref. AMM 31-35-00)

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Figure 10

ATIMS - Interface

______________________________________________________________________________________________________________________________________________________________________________________________



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ATIMS/VDR3 interface This interface is in accordance with ARINC 750 specifications. The ATSU uses the services provided by the VDR3 to communicate with the ground in DATA or VOICE mode. The ATSU receives uplink messages and transmits downlink messages through the VHF3 Data Radio. Functional split The functional split between ATSU and VDR3 is the following: - in Voice mode The ATSU controls the VHF3 transfer switch between Data and Voice mode. - in Data mode The ATSU configures the VDR3 in the appropriate protocol, the ARINC 750 data mode (VDR mode control and VDR data mode setting). The ATSU controls the VHF operational parameters of the VDR (frequency). Voice/Data select discrete The ATSU has direct control of VDR3 switching between Voice and Data mode. The VDR3 Voice/Data mode selection is controlled through: - any of the three RMPs by displaying DATA indication for DATA mode or the selected frequency for VOICE mode in the ACTIVE display. - the MCDU in VHF3 CONTROL page through COMM menu.

Port select discrete The VDR3 has two frequency control interfaces: - Port A is a digital input linked to the ATSU - Port B is a digital input linked to the RMPs. The ATSU applies a command signal to the VDR3: - when the port select discrete is a ground signal, the VDR3 takes into account the digital input port A and operates on the frequency transmitted by the ATSU - when the discrete is in open circuit, the VDR3 takes into account the digital input port B and operates on the frequency transmitted by the RMPs. NOTE:When the ATSU is faulty or not supplied, the VDR3 operates in Voice mode and the frequency is controlled by the RMPs. Switching to Voice mode 1) If the switching to Voice mode is initiated from one RMP, the ATSU router sends: - a VDR Voice mode order - a VDR port B select order. 2) If the switching to Voice mode is initiated from one MCDU, the ATSU router sends: - a VDR Voice mode order - a VDR port A select order - the voice frequency to be used by the VDR (frequency in 8.33 KHz or 25 KHz resolution range depending on the system configuration). Switching to Data mode Whether the switching to Data mode is initiated from one RMP or through one MCDU , the ATSU router sends: - a VDR Data mode order - a VDR port A select order. NOTE:Sending a VDR port A select order has for consequence the display of ACARS or DATA indication in the ACTIVE window on the RMPs.



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Figure 11

ATIMS / VDR3 interface

______________________________________________________________________________________________________________________________________________________________________________________________



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ATIMS / RMP interface Port select information The ATSU indicates to the RMPs which system between the ATSU and the RMP is controlling the frequency by means of the port select discrete. Each RMP receives the port select discrete: - when this discrete is grounded, the VDR3 frequency is controlled by the ATSU - when this discrete is in open circuit, the VDR3 frequency is controlled by the RMPs. Remote port select information The ATSU acquires the remote port select discrete from each of the three RMPs. Each RMP sends to the ATSU this signal to transmit the pilot request of switching the system controlling the VDR3 frequency. When a switching request is issued, the ATSU activates the VHF voice mode function to determine which system between the ATSU and the RMP will control the VDR3 frequency and in which mode.

______________________________________________________________________________________________________________________________________________________________________________________________



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Figure 12

ATIMS / RMP interface

______________________________________________________________________________________________________________________________________________________________________________________________



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Interface with the ground network Ground network architecture The airline can exchange data with its aircraft through a ground network which is managed by four world service providers: These providers are: ARINC, in the USA CANADIAN, in Canada JAPANESE, in Japan SITA, in the other regions. In this ground network, each service provider is responsible for its own network. The networks are interconnected, therefore the data is transferred over any network. The aircraft can be in liaison with the network through the VHF. On the ground, each service provider works on a special frequency: ARINC network 131,550 MHz CANADIAN network 131,475 MHz JAPANESE network 131,450 MHz SITA network 131,725 MHz ( SITA 725 ) 131,550 MHz ( SITA 550 )

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Figure 13

ATIMS - Ground Network

______________________________________________________________________________________________________________________________________________________________________________________________



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OPERATION System Configuration

ATSU DATALINK applications

(1)ATSU A/C INTERFACE software loading The loading procedure is ensured by the ” Boot software” included in the ATSU hardware case. The ATSU is first to be loaded with the A/C INTERFACE software which takes A/C configuration and adaptation to different system interfaces into account. It contains namely air/ground communication services and the router function. The adaptation to the A/C configuration is also ensured through pin programming acquisition, either from hard ATSU pin programming, or from pin programming information received fron CFDIU (configuration label). NOTE:The ATSU is automatically reseted after a software loading operation through an internal mechanism.The LRU IDENTIFICATION function from the BITE enables to display the software version reference in order to check the correct uploading. (2)AOC software and database loading In Pre-FANS configuration, the AOC application software and the associated database are then up-loaded in the ATSU. The AOC software package contains the specific airline datalink applications. Remote AOC applications are also embedded in peripherals which some of them are optional such as DMU, Cabin Terminals (remote AOC peripherals). (3)ATSU FANS A applications loading In FANS A configuration, the ATSU FANS A application software package is also up-loaded. In this case, the ATSU sends information to the FWCs indicating that ATC applications are active for ATC datalink alarm generation.

Three menus are available: Air/Ground Communication Menu The COMM MENU is available in pre-FANS and FANS A configurations and gives access to: - the COMM INIT page to initialize the datalink communications - the VHF3 DATA MODE page to select the Data mode and the VDR3 frequency - the VHF3 VOICE DIRECTORY page to select the Voice mode and the VDR3 frequency - the COMM STATUS page to display the availability of the VHF3 and SATCOM networks - the COMPANY CALL page to display the message content and to validate a VOICE-GO-AHEAD - the MAINTENANCE page which comprises a TEST part and an AUDIT part These menus are developped in AMM 46-24-00, P. Block 001.



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Figure 14

COMM menu

______________________________________________________________________________________________________________________________________________________________________________________________



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______________________________________________________________________________________________________________________________________________________________________________________________



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Figure 15

AOC Menu

______________________________________________________________________________________________________________________________________________________________________________________________



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Controller-Pilot Data Link Communication ATC MENU The ATC menu is only comprised in FANS A configuration and gives access to: - the LAT REQ page for request of lateral trajectory changes to the ATC center - the WHEN CAN WE page for time estimation request to the ATC center - the LOGBOOK page to display any message closed and stored on the DCDU - the LOGON STATUS page to initialize the ATS Facilities Notification with an ATC center and establish contact - the EMERGENCY page to generate emergency messages to the ATC center - the VERT REQ page for request of vertical trajectory changes to the ATC center - the OTHER REQ page for miscellaneous request like voice contact request with ATC center - the TEXT page to send justifications to negative replies to the ATC center - the REPORT page to generate automatically position reports at each ATC waypoint - the ADS page to activate the Automatic Dependent Surveillance These menus are developped in AMM 46-22-00, P. Block 001.



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Figure 16

ATC Menu

______________________________________________________________________________________________________________________________________________________________________________________________



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BITE ARCHITECTURE The ATIMS BITE is used to facilitate the aircraft maintenance in compliance with ABD048 specifications. It detects, identifies and memorizes the internal and external failures related to the ATIMS system: ATSU internal failures external interface failures with ATSU peripherals. The ATIMS BITE is ensured by the ATSU which concentrates the failure information provided by the ATSU internal monitoring. This BITE is of type 1 and operates in two modes: normal mode MENU mode Normal mode During the normal mode, the BITE: monitors the ATSU status monitors data inputs from the various ATIMS peripherals (FMGC, MCDU, CFDIU,...) permanently transmits ATIMS system status and its identification message to the CFDIU. In case of fault detection, the BITE stores the information in the fault memories and transmits it to the CFDIU by an ARINC 429 message (Label 356). The BITE memorizes the failures which occurred during the last 63 flight legs.



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Figure 17

ATIMS - BITE Architecture

______________________________________________________________________________________________________________________________________________________________________________________________



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Menu mode System Report/Test Function The BITE information (system report) and the test request (system test function) are available through MCDU menus which allows to communicate with ATIMS BITE via the CFDIU. To gain access to the BITE, it is necessary to use one MCDU (Ref. 22-82-00) All the information displayed on the MCDU during the BITE test configuration can be printed by the printer. ATIMS maintenance menu is only accessible on ground from the general maintenance menu and the SYSTEM REPORT/TEST page. This mode enables communication between the CFDIU and the ATIMS BITE by means of the MCDU.

1L 2L 3L 4L 5L 6L

ATIMS menu mode is composed of: LAST LEG REPORT CLASS 3 FAULT PREVIOUS LEGS REPORT SYSTEM TEST LRU IDENTIFICATION GROUND SCANNING TROUBLE SHOOTING DATA GROUND REPORT RETURN SPECIFIC DATA

1R 2R

5R 6R

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Figure 18

ATIMS - SYSTEM REPORT / TEST page access

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LAST LEG REPORT page This report contains the fault messages (internal and external, Class 1 and 2) recorded during the last flight. PREVIOUS LEGS REPORT page This report contains the fault messages related to the external or internal failures (Class 1 or 2) recorded during the previous 63 flight legs.

GROUND REPORT page This function is used to present Class 1, 2 or 3 internal failures when they are detected on ground. The relevant trouble shooting data are displayed by pressing the line key adjacent to the failure indication. These failures differ from those displayed on the LAST LEG REPORT page.

LRU IDENTIFICATION page This menu enables to display the Part Numbers of the different components (ATSU, DCDU, Software packages). GROUND SCANNING page This function is based on the monitoring and the fault analysis during the flight and enables consultation of the ATIMS failure recordings. The ATSU peripheral monitoring and internal cyclic tests are used in order to detect transient failures. TROUBLE SHOOTING DATA page This function provides correlation parameters and snapshot data concerning the failure displayed in the LAST LEG REPORT and the PREVIOUS LEGS REPORT pages. CLASS 3 FAULTS page This menu enables to display the Class 3 faults recorded during the last flight leg. SYSTEM TEST page The ATIMS BITE test is initiated when pressing the line key adjacent to the SYSTEM TEST indication. The test ends with the display of the following message on the MCDU: TEST OK indication when all the tests are completed and no failure has been detected or the failure message(s) when one or more failures have been detected.



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Figure 19

ATIMS - System Test Page

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SPECIFIC DATA page This menu enables acces to different functions: 2L : PIN PROGRAMMING to check the ATSU configuration with its parity validity. NOTE:The order of the pin programming display is in accordance with the one of the ATSU input connector: - a pin programming not defined (spare) is displayed with a zero value. - an active pin programming is displayed with a 1 value. 3L : DUMP TSD 4L : SW P/N PRINT OUT

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Figure 20

ATIMS - Specific Data Page

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VHF 3 LINK TEST OF THE ATSU

Figure 21

VHF 3 LINK TEST



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Figure 22

VHF 3 LINK TEST

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Figure 123

EGPWS Test Pattern

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A320 B2 Avionic

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