CTC-232_FDA

TRAINING MANUAL

CFM56-5A / -5B

FAULT DETECTION &ANNUNCIATION

JUL 2008

CTC-232 Level 3

CFM56-ALL TRAININGMANUAL

general Page Issue 0

CFMI Customer Training CenterSnecma Services

Site de Melun-Montereau,Arodrome de Villaroche

Chemin de Viercy, B.P. 1936,77019 - Melun Cedex

FRANCE

CFMI Customer Training ServicesGE Aircraft Engines

Customer Technical Education Center123 Merchant Street

Mail Drop Y2Cincinnati, Ohio 45246

USA

Published by CFMI

THIS Page InTenTIOnallY leFT BlanK





This CFMI publication is for Training Purposes Only. The information is accurate at the time of compilation; however, no update service will be furnished to maintain accuracy. For authorized maintenance practices and specifications, consult pertinent maintenance publications.

The information (including technical data) contained in this document is the property of CFM International (ge and SneCMa). It is disclosed in confidence, and the technical data therein is exported under a U.S. government license. Therefore, none of the information may be disclosed to other than the recipient.

In addition, the technical data therein and the direct product of those data, may not be diverted, transferred, re-exported or disclosed in any manner not provided for by the license without prior written approval of both the U.S. government and CFM International.

COPYRIGhT 1998 CFM INTERNATIONAl

EFFECTIVITYAll CFM56 ENGINES

CFMI PrOPrIeTarY InFOrMaTIOn


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lExIS


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Aa/C aIrCraFTaC alTernaTIng CUrrenTaCarS aIrCraFT COMMUnICaTIOn aDreSSIng and rePOrTIng SYSTeMaCaU aIr COnDITIOnIng aCCeSSOrY UnITaCMS aIrCraFT COnDITIOn MOnITOrIng SYSTeMaCS aIrCraFT COnTrOl SYSTeMaDC aIr DaTa COMPUTeraDePT aIrlIne DaTa engIne PerFOrManCe TrenDaDIrS aIr DaTa anD InerTIal reFerenCe SYSTeMaDIrU aIr DaTa anD InerTIal reFerenCe UnITagB aCCeSSOrY gearBOXaIDS aIrCraFT InTegraTeD DaTa SYSTeMalF aFT lOOKIng FOrWarDalT alTITUDealTn alTernaTeaMB aMBIenTaMM aIrCraFT MaInTenanCe ManUalaOg aIrCraFT On grOUnDa/P aIrPlaneaPU aUXIlIarY POWer UnITarInC aerOnaUTICal raDIO, InC. (SPeCIFICaTIOn)aSM aUTOTHrOTTle SerVO MeCHanISMa/T aUTOTHrOTTleaTa aIr TranSPOrT aSSOCIaTIOn

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DDaC DOUBle annUlar COMBUSTOrDaMV DOUBle annUlar MODUlaTeD ValVeDar DIgITal aCMS reCOrDerDC DIreCT CUrrenT

DCU DaTa COnVerSIOn UnITDCV DIreCTIOnal COnTrOl ValVe BOeIng DeU DISPlaY eleCTrOnIC UnITDFCS DIgITal FlIgHT COnTrOl SYSTeMDFDaU DIgITal FlIgHT DaTa aCQUISITIOn UnITDFDrS DIgITal FlIgHT DaTa reCOrDIng SYSTeMDISC DISCreTeDIU DIgITal InTerFaCe UnITDMC DISPlaY ManageMenT COMPUTerDMD DeManDDMS DeBrIS MOnITOrIng SYSTeMDMU DaTa ManageMenT UnITDOD DOMeSTIC OBJeCT DaMageDPU DIgITal PrOCeSSIng MODUleDrT De-raTeD TaKe-OFF

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FDIU FlIgHT DaTa InTerFaCe UnITFDrS FlIgHT DaTa reCOrDIng SYSTeMFDU FIre DeTeCTIOn UnITFeIM FIelD engIneerIng InVeSTIgaTIOn MeMOFF FUel FlOW (see Wf) -BFFCCV Fan FraMe/COMPreSSOr CaSe VerTICal (VIBraTIOn SenSOr)FI FlIgHT IDle (F/I)FIM FaUlT ISOlaTIOn ManUalFIn FUnCTIOnal ITeM nUMBerFIT Fan InleT TeMPeraTUreFla FOrWarD lOOKIng aFTFlX TO FleXIBle TaKe-OFFFMC FlIgHT ManageMenT COMPUTerFMCS FlIgHT ManageMenT COMPUTer SYSTeMFMgC FlIgHT ManageMenT anD gUIDanCe COMPUTerFMgeC FlIgHT ManageMenT anD gUIDanCe enVelOPe COMPUTerFMS FlIgHT ManageMenT SYSTeMFMV FUel MeTerIng ValVeFOD FOreIgn OBJeCT DaMageFPa FrOnT Panel aSSeMBlYFPI FlUOreSCenT PeneTranT InSPeCTIOnFQIS FUel QUanTITY InDICaTIng SYSTeMFrV FUel reTUrn ValVeFWC FaUlT WarnIng COMPUTerFWD FOrWarD

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OOaT OUTSIDe aIr TeMPeraTUre


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OD OUTleT DIaMeTerOgV OUTleT gUIDe VaneOSg OVerSPeeD gOVernOrOVBD OVerBOarDOVHT OVerHeaT

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QQaD QUICK aTTaCH DeTaCHQeC QUICK engIne CHangeQTY QUanTITYQWr QUICK WInDMIll relIgHT

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TranSFOrMer

SS/n SerIal nUMBerS/r SerVICe reQUeSTS/V SHOP VISITSaC SIngle annUlar COMBUSTOrSar SMarT aCMS reCOrDerSaV STarTer aIr ValVeSB SerVICe BUlleTInSCU SIgnal COnDITIOnIng UnITSDaC SYSTeM DaTa aCQUISITIOn COnCenTraTOrSDI SOUrCe/DeSTInaTIOn IDenTIFIer (BITS) (CF arInC SPeC)SDU SOlenOID DrIVer UnITSer SerVICe eValUaTIOn reQUeSTSFC SPeCIFIC FUel COnSUMPTIOnSFCC SlaT FlaP COnTrOl COMPUTerSg SPeCIFIC graVITYSlS Sea leVel STanDarD (COnDITIOnS : . in.Hg / F)SlSD Sea leVel STanDarD DaY (COnDITIOnS : . in.Hg / F)SMM STaTUS MaTrIXSMP SOFTWare ManageMenT PlanSn SerIal nUMBerSneCMa SOCIeTe naTIOnale DeTUDe eT De COnSTrUCTIOn De MOTeUrS DaVIaTIOnSOl SOlenOIDSOV SHUT-OFF ValVe

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TT oil OIl TeMPeraTUreT/C THerMOCOUPleT/e TraIlIng eDgeT/O TaKe OFFT/r THrUST reVerSerT Fan InleT TOTal aIr TeMPeraTUreT HP COMPreSSOr InleT aIr TeMPeraTUreT HP COMPreSSOr DISCHarge aIr TeMPeraTUreT. eXHaUST gaS TeMPeraTUre T lOW PreSSUre TUrBIne DISCHarge TOTal aIr TeMPeraTUreTaI THerMal anTI ICe TaT TOTal aIr TeMPeraTUreTBC THerMal BarrIer COaTIngTBD TO Be DeTerMIneDTBO TIMe BeTWeen OVerHaUlTBV TranSIenT BleeD ValVeTC(TCase) HP TUrBIne CaSe TeMPeraTUreTCC TUrBIne ClearanCe COnTrOlTCCV TUrBIne ClearanCe COnTrOl ValVeTCJ TeMPeraTUre COlD JUnCTIOnT/e TraIlIng eDgeTeCU eleCTrOnIC COnTrOl UnIT InTernal


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TeMPeraTUreTeO engIne OIl TeMPeraTUreTgB TranSFer gearBOXTi TITanIUMTla THrOTTle leVer angle aIrBUSTla THrUST leVer angle BOeIngTM TOrQUe MOTOrTMC TOrQUe MOTOr CUrrenTT/O TaKe OFFTO/ga TaKe OFF/gO arOUnDT/P TeMPeraTUre/PreSSUre SenSOrTPU TranSIenT PrOTeCTIOn UnITTr TranSFOrMer reCTIFIerTra THrOTTle reSOlVer angle aIrBUSTra THrUST reSOlVer angle BOeIngTrDV THrUST reVerSer DIreCTIOnal ValVe TrF TUrBIne rear FraMeTrPV THrUST reVerSer PreSSUrIZIng ValVeTSI TIMe SInCe InSTallaTIOn (HOUrS)TSn TIMe SInCe neW (HOUrS)TTl TranSISTOr TranSISTOr lOgIC

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VDT VarIaBle DIFFerenTIal TranSFOrMerVIB VIBraTIOnVlV ValVeVrT VarIaBle reSISTanCe TranSDUCerVSV VarIaBle STaTOr Vane

WWDM WaTCHDOg MOnITOrWf WeIgHT OF FUel Or FUel FlOW WFM WeIgHT OF FUel MeTereD WOW WeIgHT On WHeelSWTaI WIng THerMal anTI-ICIng


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IMPERIAl / METRIC CONVERSIONS

mile = ,0 km ft = 0, cm in. = , mm mil. = ,

sq.in. = , cm

USg = , l (dm) cu.in. = . cm

lb. = 0. kg

psi. = .0 kPa

F = . x C +

METRIC / IMPERIAl CONVERSIONS

km = 0. mile m = . ft. or . in. cm = 0. in. mm = . mils.

m = 0. sq. ft. cm = 0. sq.in.

m = . cu. ft. dm = 0. USa gallon cm = 0.0 cu.in.

kg = .0 lbs

Pa = . 0- psi. kPa = 0. psi bar = . psi

C = ( F - ) /.

TABlE OF CONTENTS

EFFECTIVITYAll CFM56-5A/-5B FOR A318-A319-A320-A321


CFM56-5A/-5B TRAININGMANUAL

COnTenTSFaUlT DeTeCTIOn& annUnCIaTIOn

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EFFECTIVITYALL CFM56-5A/-5B FOR A318-A319-A320-A321

CFMI ProPrIetary InForMatIon


ContentsFaULt DeteCtIon& annUnCIatIon

Page 16May 07

seCTIoN PAGe seCTIoN PAGe

LexIs . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5

taBLe oF Contents . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 15

arCHIteCtUre . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 17

InterFaCes . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 27

eCU sIGnaLs . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 55

FaULt DeteCtIon . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 65

enGIne InDICatInG systeMWarnInG InDICatIons . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 75

enGIne MonItorInG & GroUnD oPeratIonMessaGe InterroGatIon . . . . . . . . . . . . . . . . . . . . . . . . . . . . 135CentraLIZeD FaULt DIsPLay systeM . . . . . . . . . . . . . . . . . 139aIrCraFt InteGrateD Data systeM . . . . . . . . . . . . . . . . . . 189




arCHITeCTUreFAUlT DETECTION& ANNUNCIATION

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ARChITECTURE

EFFECTIVITYAll CFM56-5A/-5B FOR A318-A319-A320-321




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ENGINE CONTROl SYSTEM

System components

The CFM-a/-B engines incorporate a computer-based Full authority Digital engine Control (FaDeC) system.

The engine control system is composed of the following elements:

- electronic Control Unit (eCU), containing two identical computers, designated channel a & B.

- Hydro-mechanical Unit (HMU), which converts electrical signals from the eCU into hydraulic pressures to drive the engines valves and actuators.

- eCU alternator.- engine Identification plug (ID plug).- engine pressure, temperature and speed sensors.- Variable Stator Vane (VSV) actuators.- Variable Bleed Valve (VBV) actuators.- High Pressure Turbine Clearance Control (HPTCC).- low Pressure Turbine Clearance Control (lPTCC).- Fuel return Valve (FrV).- Thrust reverser (Tr) control.- Starter air Valve (SaV).- Ignition components / control system.

(-5B):- Transient Bleed Valve (TBV).

(-5A):- Burner Staging Valve (BSV).

Electronic Control Unit (ECU)

The eCU is the prime component of the engine control system.

The eCU governs the engine in response to thrust command inputs from the airplane and provides information to the airplane for flight compartment indication, maintenance reporting and, optionally, engine condition monitoring.

Control system maintenance is assisted by extensive eCU internal software called Built-In-Test-equipment (BITe), which monitors engine data and eCU status flags to detect engine failures.





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FADEC COMPONENTSCTC--00-0

VBV VSV TBV BSV hPTCCV

lPTCCVFRV

AlTERNATOR

FEEDBACK SIGNAlS

T5PS13

CONTROl SIGNAlS

REVERSERSOlENOIDS+ SWITChES

FUEl hYDRO-MEChANICAl

UNIT

FUElFlOW P25

ECU

T12 T25 T3 T49.5 TCASE TEO P0 PS12 PS3 N1 N2

IGNITION

STARTER AIR VAlVE

ANAlOGDISCRETESIGNAlS

115V400hz 28V

ARINCDATABUSES

TRA

FMV

* **

* -5B ONlY** -5A ONlY





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ECU INPUTS AND OUTPUTS

Electrical interfaces

The following chart is a summary of the eCU electrical interfaces to show which connectors interface with channel a and which interface with channel B.





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ElECTRICAl CONNECTORSCTC--00-0

ChANNEl ACONNECTOR

(ODD)CONNECTORChANNEl B

(EVEN)FUNCTION

J1J3J5J7J9J11

ShAREDJ13J15

ShAREDShARED

J2

J14J12J10

J6J8

J4A/C POWER (28V) AND IGNITER POWER (115V)A/C INPUT/OUTPUT AND TlAThRUST REVERSERSOlENOIDS, TORQUE MOTORS, RESOlVERS, N2AlTERNATOR, SAV, N1 AND T12lVDT'S, RVDT'S, T25, BSV POSITION SWITCh *ENGINE IDENTIFICATION PlUGWF METER, ThERMOCOUPlESTEST INTERFACE

J1 J5 J7 J

14 J15 J8

J6 J2

J3 J9 J1

1 J13

J12 J10 J

4

* BSV POSITION SWITCh ON -5A ONlY





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ENGINE TESTS

ECU initialization

If the engine is not running, the eCU becomes fully operational within a maximum of three seconds after application of airplane power, or an external reset.

If the core speed is greater than 0% n, the eCU performs a short initialization and is fully functional in less than 0 ms after application of airplane power.

each eCU channel performs a reset initialization sequence in response to aircraft-generated resets, or at power-up.

an aircraft-commanded reset occurs when the master lever is toggled from On to OFF.

During reset initialization, all raM variables are initialized, except for a special reserved area. This area of raM is not initialized as it is allocated to parameters critical to engine operation and which must maintain their values prior to the reset operation.

Built-In-Tests

Built-In-Test-equipment (BITe) monitors the system and memorizes failures.

The BITe detects and isolates failures, or combinations of failures, in order to determine the health status of the channels and to transmit maintenance data to the aircraft.

There are two types of Built-In-Test : Initialization test and Periodic test.

The Initialization tests cover functions which cannot be continually tested without disturbing the eCU system operation. The typical tasks of an Initialization test are processor test, memories test and output driver disconnect tests.

The Periodic tests cover functions which can be continually tested. These tests are similar to the Initialization tests, but are run in background as time permits.

Specific tests are available to verify certain engine functions. These tests are the FaDeC test (non-motoring & motoring), ignition test and thrust reverser test.





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TESTSCTC--00-00

POWER UP TEST

BUIlT-IN TESTS PERMANENTMONITORING

ENGINE RUN

OPERATIONAlFUNCTION

POWER UP

YES

NOSPECIFIC TESTS

FADEC TESTThRUST REVERSER TEST

IGNITION TEST





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ENGINE CONTROl SYSTEM

Electronic Control Unit (ECU)

The eCU has two channels, a and B, and both channels are capable of controlling the engine.

The two channels are identical and permanently operational, but they operate independently from each other. each channel has a full complement of sensors, interfaces to the engine and aircraft, central processor and output drivers.

as well as continuously checking and processing their own inputs, the channels compare each others data over a Cross Channel Data link (CCDl), to ensure that there are no anomalies.

The two eCU channels operate their output drivers on an active/standby principle. Both channels always receive inputs and process them, but only the channel in control, called the active channel, delivers control outputs (solenoids/torque motors). The other is called the Standby channel.

The purpose of the dual-redundant architecture is to minimize the effects of control system faults on the engine operation.

Channel selection and fault strategy

active and Standby channel selection is performed at eCU power-up and during operation.

active and Standby selection is based upon the health of the channels and each channel determines its own health status. The healthiest is selected as the active channel.

When both channels have an equal health status, active/standby channel selection alternates with every engine start, if n was greater than ,000 rPM during the last run.





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ECU DESIGNCTC--00-00

ChANNElA

ChANNElB

ECU

INPUTS

INPUTS

CCDl

ACTIVE

STAND-BY






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InTerFaCeSFAUlT DETECTION& ANNUNCIATION

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INTERFACES





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RElATED ENGINE AND AIRCRAFT INTERFACE COMPUTERS

There are many aircraft computers and systems that interface directly or indirectly with the engine FaDeC systems.

These computers can be classified into various types of activity:

Interface and Data storage for communication and data retrieval

eIU = engine Interface UnitCFDIU = Centralized Fault Display Interface UnitDMU = Data Management UnitFDIMU = Flight Data Interface Management UnitDFDrS = Digital Flight Data recording system

Warning and display for Crew and maintenance

FWC = Flight Warning ComputerSDaC = System Data acquisition ConcentratorDMC = Display Monitoring ComputereCaM = electronic Centralized aircraft Monitoring

Air Data for engine control

aDIrU = air Data Inertial reference UnitaDC = air Data Computer

Bleed and pneumatic data for engine control

eCS = environmental Control SystemBMC = Bleed Monitoring ComputernaI = nacelle anti-IceWaI = Wing anti-Ice

Aircraft system information for engine control

FlSCU = Fuel level Sensing Control UnitlgCIU = landing gear Control Interface UnitSFCC = Slats and Flaps Control Computer

Auto pilot for engine power targeting

FMgC = Flight Management and guidance ComputerFCU = Flight Control Unit





Page May 0

INTERFACE COMPUTERSCTC--00-0

AIR DATA & INERTIAl REFERENCE SYSTEM(ADIRS)

ENGINE INTERFACE UNIT(EIU)

FlIGhT CONTROl UNIT(FCU)

FlIGhT MANAGEMENT & GUIDANCE COMPUTER(FMGC)

CENTRAlIzED FAUlT DISPlAY INTERFACE UNIT(CFDIU)

ENGINE VIBRATION MONITORING UNIT(EVMU)

DATA MANAGEMENT UNIT(DMU)

MUlTIPURPOSE CONTROl & DISPlAY UNIT(MCDU)

DIGITAl FlIGhT DATA RECORDING SYSTEM(DFDRS)

ElECTRONIC CENTRAlIzED MONITORING SYSTEM(ECMS)

DISPlAY MANAGEMENT COMPUTER(DMC)

FlIGhT WARNING COMPUTER(FWC)

SYSTEM DATA ACQUISITION CONCENTRATOR(SDAC)

SlAT/FlAP CONTROl COMPUTER(SFCC)

lANDING GEAR CONTROl INTERFACE UNIT(lGCIU)

ENVIRONMENTAl CONTROl SYSTEM(ECS)

FADEC

SYSTEM





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FADEC INTERFACES - INPUTS

The FaDeC interface arInC input busses are shown on the two next pages. The eCU () is interfaced with:

- The air data computers (aDC, aDIrU ()).- The engine interface unit (eIU).

The eIU receives data from arInC busses; the data received are provided to the eCU under the same format.

The eIU also receives discrete values, which are transformed into arInC information and delivered to the eCU. This information comes from:

- The cockpit switches (Master lever, rotary selector, Fadec ground power, etc.).

- The autothrottle loop (from FMgC () and the FCU)

- The environmental control system (eCS), the Bleed system.

- The Fuel level Sensing Control Unit (FlSCU ()).- The landing gear Control Interface Unit (lgCIU).- The Slat and Flap control Computer (SFCC).- Wing anti-ice (WaI) and nacelle anti-Ice (naI).

The eCU manages power according to thrust modes:

- Manual mode, depending on Throttle lever angle. (not shown).

- autothrust mode, according to the autothrust function generated by the auto-Flight System (aFS).

The FMgC computes the autothrust order and sends it to the eCU via the FCU and eIU. The FCU is the interface for transmission of engine data from the FMgC to the eIU. Thrust limit computation is performed by the eCU, except when the alpha floor protection is activated.

The eCU provides two Idle Selection modes:

- approach idle, when flaps are extended (SFCC), or landing gear is down (lgCIU).

- Modulated idle up to approach idle, depending on oil temperature, (IDg cooling), air conditioning and anti-ice demand (Zone and pack controller).

The eCU allows fuel to be returned to the a/C for IDg cooling and engine oil cooling if the a/C fuel management computers (FlSCU) allow it.





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FADEC INTERFACES - ARINC 429 INPUT BUSSESCTC--0-00

J3

5

CFDIU

EIU 1ECS Ch A

ECU 1

Ch B

DISCRETE lINKS

EIUFMGC

FCUECSBMCCFDIU

FlSCUlGCIU

SFCC

NAIWAI

WhERE :

= ENGINE INTERFACE UNIT= FlIGhT MANAGEMENT AND GUIDANCE COMPUTER= FlIGhT CONTROl UNIT= ENVIRONMENTAl CONTROl SYSTEM= BlEED MONITORING COMPUTER= CENTRAlIzED FAUlT DISPlAY INTERFACE UNIT= FUEl lEVEl SENSING CONTROl UNIT= lANDING GEAR CONTROl INTERFACE UNIT= SlATS AND FlAPS CONTROl COMPUTER= NACEllE ANTI-ICE= WING ANTI-ICE

ARINC 429 lINKS

FlSCU 1(2)

lGCIU

SFCC

COCKPITSWITChES

zONECONTROllER

FCUFMGCAFS

BMC 1(2)

PACKCONTROllER

NAIWAI





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FADEC INTERFACES - ARINC 429 INPUT BUSSES

The aDIrUs send air data parameters to the eCU for power management and engine control. This information is shared by the two eCU channels.

These parameters are combined with engine data, with a weighted average, to compute altimetric correction for ambient temperature, pressure, and Mach effects.

TlA information

The throttle is directly hardwired to the eCU; therefore, even in the event of a major computer breakdown, it is possible to continue almost normal engine control. as shown on the drawing, this information is also shared information. Some functions will be inhibited, like the autothrust control. The engine will operate under continuous ignition and the minimum idle speed will be approach idle.





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FADEC INTERFACES - ARINC 429 INPUT BUSSES AND TlACTC--00-00

ADCs

ADIRU

WhERE :

= AIR DATA INERTIAl REFERENCE UNIT

J3

J4

ADIRU 1

ADIRU 2

Ch A

ECU 1

Ch B

OThERENGINE

hARDWIRED lINKS

ARINC 429 lINKS

3

3

IDENTICAlFOR OThERENGINES

ThROTTlESETENG 1





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FADEC INTERFACES - ARINC 429 OUTPUTS

The FaDeC interface arInC output busses are shown on the opposite page. The main eCU interface is the eIU, however it is not the only one. The eCU () is interfaced with:

- The engine interface unit (eIU)- The Display monitoring computers (DMC, , )- The Flight Warning computers (FWC , )- The Flight Management Computers (FMgC , )- The Data Management Unit (DMU)

There are two low-bit-rate synchronized arInC output busses per eCU channel. The information provided to the aircraft system is used for the following purposes:

- The CFDS is a Centralized maintenance aid system, which gives the maintenance technicians a means to read the information related to most of the aircraft systems and to initiate tests of these systems from the cockpit. The CFDS communicates with the eCUs through the eIU.

- The DFDr collects various engine and a/C system parameters through the FDIMU and the CFDS system, and processes them internally. The DFDr stores the data collected over the last hours.

- The BMC (Bleed Monitoring Computer) receives the bleed information from the engine through the eIU.

The same bus provides primary parameters (n, n, egT, fuel flow), which are sent by the eCU directly to the eCaM through the DMC computers. Secondary parameters are sent to the eCaM by the SDaC. The FWC are involved in warning messages displayed on the eCaM system if any abnormal conditions are sensed.

The DMU collects, stores and processes various a/C and engine data and generates condition reports.

The FMgC receives the actual n and uses it to compute the autothrust order, and sends it to the FCU, which provides the autothrust target as explained on the previous page.

nOTe:The FDIMU (Flight Data Interface Management Unit) is a computer that centralizes all the data prior to having it written in the DFDr.





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FADEC INTERFACES - ARINC 429 OUTPUT BUSSESCTC--0-00

J3

J4

Ch A

ENGINE 1 (2)

Ch B

FDIMU FlIGhT DATA INTERFACE MANAGEMENT UNIT

EIU 1 (2)

BMCCFDS

DFDRS

FDIMU

DMC 1

FWC 1 FWC 2 DMU

DMC 3 DMC 2 FMGC 1&2





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ENGINE SENSORS DIRECTlY CONNECTED TO AIRCRAFT COMPUTERS AND DISPlAY lOGIC

Some sensors are directly connected to aircraft computers; their readings are not provided through the eCU. This means that fault reports will be only available through the CFDS.

These computers are:

- The SDaCs- The eIU- The FWCs- eVM

The SDaCs receive:

- Fuel filter clog engine (SDaC)- Fuel filter clog engine (SDaC)- Oil filter clog engine & - Oil Pressure transmitter- Oil quantity transmitter

The engine Interface Unit receives the oil temperature parameter taken on the lube unit for engines & . This information is just used for cockpit indication.

The low oil pressure switch information is delivered to the eIU and Flight Warning Computers (FWCs).

an aircraft logic described later in this course provides the crew with warning and advisory information.

Maintenance messages will be found in the in the aircraft Centralized Fault and Display System (CFDS).

The aircraft is equipped with a single engine and Vibration Monitoring computer. It receives the speeds and vibrations from both engines.

The low oil pressure switch information is delivered to eIU and Flight Warning Computers (FWCs).

an aircraft logic described further in that course provide the crew with warning and advisory information.

Maintenance message will be found in the in the aircraft Centralized Fault and Display System (CFDS).

The aircraft is equiped with a single engine and Vibration Monitoring computer. It receives the speeds and vibrations from both engines.





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ENGINE SENSORS DIRECTlY CONNECTED TO AIRCRAFT COMPUTERSCTC--0-00

ENGINE AIRCRAFT

SDAC 1

4

DUAl

DUAl

SDAC 2

EIU

lOW PRESSUREAND GROUND

lOGIC

FWC 1

FWC 2

EVMU

FUEl FIlTER ClOG ENG1

OIl FIlTER ClOG

N1 SPEED SENSOR

N2 SPEED SENSOR

N1 BRG ACC

TRF ACC

OIl PRESSURE xMTR

OIl TEMP INDIC.

lUBE UNIT

OIl QTY xMTR

lOW OIl PRESS SW

FUEl FIlTER ClOG ENG2





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FADEC INTERFACES

The eCaM monitors operational data in order to display warnings and system information. FaDeC system data is processed by the SDaCs, FWCs and DMCs before being presented on the eCaM engine Warning Display (eWD) and System Display (SD).

The eWD is dedicated to the primary engine parameters and engine warning messages. The SD is dedicated to the propulsion system parameters when the engine system page is called either automatically, or manually.

The SDaCs digitalize systems data and transmit it to the DMCs. The SDaCs receive systems information concerning amber cautions and transmit it to the FWCs.

The FWCs receive systems data concerning red warnings and memos, generate messages and activate attention getters. Both FWCs have the same engine monitoring capability.

The DMCs use outputs from the FWCs to display information on the lower part of the eWD.

The eVMU provides vibration information to the SDaCs for real time monitoring on the eCaM and to the DMU for condition monitoring.





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ENGINE DATA DISPlAY PROCESSINGCTC--00-00

DMC 3

FWC 2

2

3

ECU

EIU

FWC 1

MASTER

CAUT

MASTER

WARN

DMC 2DMC 1

UPPERECAMEWD

lOWERECAM

SD

SDAC 2 3

EVMU

SDAC 1





Page 0May 0

ON-BOARD MAINTENANCE SYSTEM

acquisition of aircraft system data is performed by the Centralized Fault Display System (CFDS). This includes the eCaM to display warnings and system information, the DFDrS, which is an obligatory recording system, the CFDIU and the DMU, which is the main component of the aircraft Integrated Data System (aIDS).

In each aircraft system computer, a BITe monitors the system and memorizes the failures. after failure detection, the BITe is able to identify the possible failed lrUs and give a snapshot of the system environment when the failure occurred. all information necessary for maintenance and troubleshooting is memorized in nVM.

The eCU is able to distinguish between faults external and internal to the FaDeC system. external faults are defined as those detected on aircraft interfaces not dedicated to the FaDeC system. external functions include the aDIrUs, the eIU and aircraft power supplies. all other faults in the system (eCU, HMU, sensors, cables, components, etc.) are considered internal faults.

The eVMU is a built-in test equipment. all internal or external failures are monitored with the CFDS. Specific maintenance procedures are available for troubleshooting.

The Data Management Unit (DMU) also reads vibration information. It builds report in which Vibration, and phase angle for n and n rotors are specified.

The main components of the CFDS are the CFDIU, which has a main channel and a standby channel, and the aircraft system BITeS. The CFDIU continuously scans the busses from the aircraft systems and if a failure message from a system BITe is present on a bus, the CFDIU copies and stores it.

The CFDIU also stores the eCaM messages generated by the FWCs and acts as an interface for some class failures, transmitted by the DMU, and used for the eCaM maintenance status.





Page May 0

ON-BOARD MAINTENANCECTC--00-0

FWC 2FWC 1

MCDU 2MCDU 1

SDAC 2SDAC 1

PDlDARACARSMDDUPRINTER

DMU (AIDS)

CFDIU

EVMU

EIU

ECU





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ON-BOARD MAINTENANCE SYSTEM (CONTINUED)

The aircraft systems are divided into types , and , depending on their capabilities and connection to the CFDIU. Most systems are type and these can memorize failures which have occurred in the last flights. The engine (FaDeC) is a type system.

The MCDU is the operators interface with the CFDIU.

The DMU records significant operational parameters in order to monitor the engines, the aircraft performance and to analyze specific aircraft problems. a Portable Data loader (PDl) can be connected to the DMU for up and down loading. an optional Digital aIDS recorder (Dar) enables data to be stored on a replaceable cassette.

Most reports may be printed and data can also be transmitted to the ground, manually or automatically, through the aCarS. Data may also be loaded into the maintenance computers, through the Multi-purpose Disk Drive unit (MDDU).






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ON-BOARD MAINTENANCE SYSTEM - FAUlT DISPlAY PRINCIPlE

During flight, all faults, failures, or abnormal situations are recorded by the aircraft.

The faults are classified by level; there are three levels of failure:

- Class faults are indicated to the crew in real time.

- Class faults are not indicated to the crew in real time. at the end of the flight, when the aircraft is on the ground for more than 0 seconds after 0 kt, a status box comes on the upper eCaM.

- Class faults are not indicated to the crew. These faults will be retrieved in the CFDS memory by the maintenance specialist.

There are two levels of class faults. The class long dispatch faults concerning items, and the SM (scheduled maintenance) concerning faults that must be repaired within 00 hours.

The faults provided to the crew in real time are displayed on the UPPer eCaM in the message area:

- eng OVerTeMPeraTUre- eng COMPreSSOr Vane

In addition, there are some audio warnings. The single tone indicates an advisory information, a parameter drifting from its normal value. The continuous chime indicates a more important failure. In both cases, if existing the relevant page comes on the lower eCaM.

When the eCaM displays an abnormal information item, it is displayed with a certain logic that will be described for each parameter in this course.

In addition after landing a post flight report is printed automatically 0 seconds after 0 kt. This report is for the maintenance people, in order to initiate the troubleshooting.

The post flight report displays only class and faults, it must be also reported in the a/C logbook.

The post flight report indicates the eCaM WarnIngS indicated to the crew during the flight and the FaUlT MeSSageS for the maintenance specialist.

The PFr indicates which system has a fault and when it occurred with a specific grammar, which is useful to find the troubleshooting procedure.





Page May 0

ON-BOARD MAINTENANCE SYSTEM - FAUlT DISPlAY PRINCIPlECTC--0-00

1.296 56 C 56 C

2300 2185

EGT0C

F.FKG/HKG/H

104.99 99.9N2%%

105N1%% FLX

FOB: 6700KG 6700KG

FLAPS F

2

670

105

952

105

105

ENG 1 EGT OVERLIMITENG 1 COMPRESSOR VANEENG 1 EGT OVERLIMITENG 1 COMPRESSOR VANE

T.0 INHIBITIGNITIONLDG LT

T.0 INHIBITIGNITIONLDG LT

84.2 84.2

F.USEDVIB (N1)

1.6 1.6

VIB (N2)

0.3 0.3

1.6 1.6

0.3 0.3

ENGINE

14.4

166

207 215KG

OIL

QT

PSI

14.5

166

25

300 300

25

0

0

0

0

1401400C

+10TAT C+10SAT C

54700GW KG2503 H

A/C ID DATE GMT FLNT CITY PAIR.F-GFXU 20JUN 1045 AF2706 LFPO LFPNO

MAINTENANCE POST FLIGHT REPORT

A/C ID DATE GMT FLTN CITY PAIR.F-GFXU 20JUN 1842 AF2706 LFPO LFPN

ECAM WARNING MESSAGES

GMT PH ATA 0920 05 73-21 ENG 1 EGT OVERLIMITGMT PH ATA 0920 05 73-21 ENG 1 COMPRESSOR FAULTGMT PH ATA 1125 06 28-21 FUEL L TK PUMP 1 LO PRFAULT MESSAGES

GMT PH ATA SOURCE IDENT.0920 05 73-21-10 J8,HMU EIU1FAD VBVTM,ECU

GMT PH ATA1125 06 28-21-00 PUMP(1QW)







GMT PH ATA1125 06 28-21-00 PUMP(1QW)

MASTER

WARN

DINGDING

DINGDING

AFTER FlIGhT





Page May 0

ON-BOARD MAINTENANCE SYSTEM - FAUlT RETRIEVAl PRINCIPlE

The Post flight report may be available with the MCDU (Multipurpose Control and Display Unit) through different ways:

- By acars (aircraft Communication and reporting System) if one presses the Send MCDU key on the bottom left of the screen.

- It can be printed by pressing the opposite Print key on the bottom right of the screen.

- The last possibility to obtain the PFr is automatic, during landing 0 seconds after 0 kts, the PFr is printed.

The Post Flight report provides the eCaM warning and the maintenance message.

as shown on the opposite page, the eng OVerlIMIT provided to the the crew does not have any maintenance message. Indeed the engine related does not have any problem. Overtemperature is linked to the VBV for which there is a flight crew message : eng COMPreSSOr Vane and this message is related to the maintenance message J, HMU (VBV TM), eCU.

Sometimes, according to the flight phase and the FWC (Flight Warning Computer) logic, the crew is not always aware of a failure. This logic belongs to airbus and will not be described in this course. To conclude on that point, the airplane logic says dont bother the crew with that fault now, there are something else to do.





Page May 0

ON-BOARD MAINTENANCE SYSTEM - FAUlT RETRIEVAl PRINCIPlECTC--0-00

REPORT AUTOMATICAllY PRINTED AT 80 KT + 30 SECONDS.OR AVAIlABlE ON DEMAND BY PRESSING ThE PRINT KEY.OR AVAIlABlE BY ACARS SYSTEM BY PRESSING ThE SEND KEY.







GMT PH ATA1125 06 28-21-00 PUMP(1QW)







GMT PH ATA1125 06 28-21-00 PUMP(1QW)

DIRDIRDIR ROGPRORO PERFPERPER NITININI DATATA

-PLNF-PL-PL NAVNN REDPRRPRPR F-PLNF-F- DUMCDUDUMENUMM

PORTPOPO

AGEPAGEPAPAAPA

BRT

AA BB CC DD EE

FF GG HH II JJ

KK NN OO

PP SS TT

UU XX YY

ZZ ++-- LRCL

55

FYOVFYO

11 2222 33

44 66

77 88 99

00 //

FAIL

FMGC

MCDU

MENU

RAIRR

XTNEXTXT

LL MMM

QQ RR

VV WW

DRADD CSECCCCMMCCCCTCATCTC

OFFELEEELFUELELELEE

CFDS





Page May 0

ON-BOARD MAINTENANCE SYSTEM - FAUlT RETRIEVAl PROCEDURE

To get the correct procedure, the PFr data has to be entered on the airnav software as shown on the opposite page.

When the search button is pressed, the appropriate procedure is available.

an intermediate screen pops-up to select the appropriate aircraft type.





Page May 0

ON-BOARD MAINTENANCE SYSTEM - FAUlT RETRIEVAl PROCEDURECTC--0-00







GMT PH ATA1125 06 28-21-00 PUMP(1QW)







GMT PH ATA1125 06 28-21-00 PUMP(1QW)





Page 0May 0

ON-BOARD MAINTENANCE SYSTEM - CORRECTIVE PROCEDURE

The Maintenance message J, HMU (VBV TM), eCU provides the three most probable faulty components. In our case, the J harness, the HMU itself and in particular the VBV torque motor, or the eCU.

The procedure indicates the possible causes and as usual the related procedure to get to the components on the engine.





Page May 0

ON-BOARD MAINTENANCE SYSTEM - CORRECTIVE PROCEDURE (1)CTC--0-00





Page May 0

ON-BOARD MAINTENANCE SYSTEM - CORRECTIVE PROCEDURE

now that the procedure is available, the philosophy is the following: by disconnecting the indicated connectors at defined places and measuring the pin resistance, it is possible to isolate the cause of the defect.

In our case, the eCU connector is removed if the resistance measured on the harness and component is within the specified values, i.e. the eCU is certainly the cause of the defect. Otherwise, if the resistance is out of the limit, it might be either the harness or the component. By measuring the resistances on the component itself, it is possible to say which one is the cause of the failure.





Page May 0

ON-BOARD MAINTENANCE SYSTEM - CORRECTIVE PROCEDURE (2)CTC--0-00

DISCONNECT hARNESSFROM ECU CONNECTOR

ChECK RESISTANCEOF

hARNESS / COMPONENT

DISCONNECT hARNESSFROM COMPONENT

NO YES

YES

NO

ChECK RESISTANCEOF COMPONENT

REPlACE COMPONENT REPlACE hARNESS REPlACE ECU

IS RESISTANCEWIThIN lIMIT ?

IS RESISTANCEWIThIN lIMIT ?






Page May 0




eCU SIGnaLSFAULT DETECTION& ANNUNCIATION

Page 55Jul 08

ECU SIGNALS

EFFECTIVITYALL CFM56-5A/-5B FOR A318-A319-A320-321




Page 56Jul 08

ECU INpUTS

each eCU channel receives critical engine signal inputs from separate sources.

Dual inputs:LVDt/rVDt and resolver:

- VSV, VBV, tBV (-5B onLy), LPtC, HPtC, FMV.

Valve position switches:- SaV, FrV, BSV (-5a onLy), HPSoV.

PS12.PS3.P0.t25.t12.t3.teo.teCU.n1 and n2 signals.

Dual power:engine alternator.

When the signal is less critical, only one source sends a signal, which is connected to both channels.

Shared inputs:eGt.t Case.Fuel flowmeter.HMU oSG switch.ID plug inputs.

non-critical control inputs are only sent to one channel.

Single inputs:PS13 to channel a (PMUX option).t5 to channel a (PMUX option).P25 to channel B (PMUX option).





Page 57Jul 08

ENGINE INpUTS TO THE ECUCtC-232-009-00

DUAL

LVDT/RVDT/RESOLVERSOL pOSITION SW (e.g. SAV)pS12pS3p0T25T12T3TEOTECUN1N2

ENGINE ALTERNATOR

CHANNELA

CHANNELB

ECU

SINGLE

p25 (pMUX)

SINGLE

pS13 (pMUX)T5 (pMUX)

SHARED

EGTTCFUEL FLOWMETERHMU OSG SWID pLUG





Page 58Jul 08

ECU INpUTS

Identification plug parameters

the engine identification (ID) plug provides the eCU with engine configuration information. It is plugged into connector J14 on the eCU and attached to the fan case by a metal strap. It remains with the engine even after eCU replacement.

It consists of a plug connector and a coding circuit. the coding circuit is welded to the plug connector pins and is equipped with fuse and push-pull links, which are used to ensure, or prohibit connection between the different plug connector pins.





Page 59Jul 08

ID pLUG (J14) pARAMETERS (-5B)CtC-232-058-01

RATING

FUSE LINKS

29RATING 16RATING 15

RETURN 28NEW/pREVIOUS 1

N1 TRIM 5N1 TRIM 14RETURN 4N1 TRIM

pUSH-pULL LINKS

11pMUX 13

RATING 12SpARE 24

RETURN 255A/5B 20

N1 TRIM pARITY 21RETURN 8

SAC SpARE, DAC NACTB 31CONFIG. pARITY 30

ENGINE CONFIG. (5B) 19ENGINE CONFIG. (5B/p) 7

ACOUSTICAL pACKAGE INSTALLED 6RETURN 18

TCMA DISABLED 23SYSTEM COMBUSTOR CONFIG LSB 22

SYSTEM COMBUSTOR CONFIG 9CHANGE AIRCRAFT TYpE 3

SYSTEM COMBUSTOR CONFIG MSB 2RETURN 10

SpARE 17SpARE 26SpARE 27

- 5B

tHIS PaGe IntentIonaLLy LeFt BLanK





Page 60Jul 08





Page 61Jul 08

ID pLUG (J14) pARAMETERS (-5A)CtC-232-083-00

FUSE LINKS

- 5ARATING 29RATING 16RATING 15

RETURN 28pMUX INHIBIT 13

BUMp 12BUMp 24

RETURN 25OpEN 30OpEN 31OpEN 19OpEN 7

ENGINE CONFIG 6RETURN 18

SpARE 3SpARE 2

A3 TEST CELL 9pARITY 22SpARE 13

RETURN 10UNAVAILABLE 1UNAVAILABLE 20UNAVAILABLE 5UNAVAILABLE 11UNAVAILABLE 14UNAVAILABLE 21UNAVAILABLE 26UNAVAILABLE 17

SHIELD GROUND 27

NEW pLUG

RATING 29RATING 16RATING 15

RETURN 28pMUX INHIBIT 13

RATING 12RATING 24

RETURN 25ESN 30ESN 31ESN 19ESN 7ESN 6

RETURN 18ESN 3ESN 2ESN 9ESN 22ESN 13

RETURN 10

OLD pLUG





Page 62Jul 08

ECU OUTpUTS

each eCU channel has 2 independent arInC429 serial databuses, which interface with the aircraft. there are no differences in the bus outputs, but data which is specific to a channel, such as fault and maintenance data, may differ from channel to channel. Sensor values that are output by the two separate eCU channels will also be slightly different, but within signal tolerance requirements.

Cockpit indication data is output to the aircraft to keep the flight crew informed of the operational status of system components and FaDeC system controlled engine parameters.

Maintenance data is output, via the arInC429 buses to the aircraft maintenance computer. this data provides information to help the ground crew identify system faults and isolate the faults to the correct LrU, or system interface.

engine condition monitoring parameters are output to the aircraft, via the arInC buses, as digital equivalents of all sensor inputs to the eCU.

Channels a and B deliver constant outputs, irrespective of which channel is in control. Channel switch-over does not affect the output data of the eCU, except for the status indication for the channel in control, items specific to the channel in control and whatever faults caused the switch-over.

Both eCU channels are able to control torque motor and solenoid output loads, but only the active channel supplies control outputs during normal operation and the standby outputs are not used.

the eCU turns the two engine igniters on, or off, using relay-controlled switches, internal to the eCU.

each channel of the eCU also provides excitation voltages for the throttle control system resolvers.





Page 63Jul 08

ECU OUTpUTSCtC-232-011-01

ECU

ACTIVE / STD-BY

STD-BY / ACTIVE

STD-BY / ACTIVE

STD-BY / ACTIVE

ACTIVE / STD-BY

ACTIVE / STD-BY

ENGINE

FMVHpTCLpTCVSVVBV

TBV * BSV **

SAVFRV

IGNITERS

ENGINE SENSORS

RVDTLVDT

RESOLVER

CHANNELA

CHANNELB

AIRCRAFT

OUTpUT DATABUS 1B

OUTpUT DATABUS 2B

T/R pRESSURE VALVE

T/R DIRECT VALVE

TRA EXCITATION

OUTpUT DATABUS 1A

OUTpUT DATABUS 2A

T/R pRESSURE VALVE

T/R DIRECT VALVE

TRA EXCITATION

* -5B ONLY** -5A ONLY

tHIS PaGe IntentIonaLLy LeFt BLanK





Page 64Jul 08




FaUlT DeTeCTIOnFAUlT DETECTION& ANNUNCIATION

Page May 0

FAUlT DETECTION





Page May 0

FAUlT DETECTION

Signal processing

Within the eCU, the various inputs from sensors, switches and the aircraft pass through several stages of checks before the values received are finally selected to be used in the control law calculations.

Both eCU channels validate their inputs, process the data and check their outputs identically.

after they have been converted to a digital format, the parametric/discrete values and the arInC datawords must first pass through a signal and range check logic.

The values are then compared across the CCDl before being selected for control low calculations. The control laws are entirely managed by the eCU software and will not be described here as they have no impact on fault detection.

after the values have been calculated and processed in the control law logic, they pass through to the output stage for transmission to engine, or aircraft systems.





Page May 0

SIGNAl PROCESSINGCTC--0-0

lOCAl ChANNEl

PARAMETERSElECTION CONTROllAW

CAlCUlATIONS

OUTPUTCONTROl

ANAlOG

DISCRETE

ARINCS/TM

ARINC

CROSS ChANNEl

ASABOVE

ASABOVE AS

ABOVE

ASABOVE

ANAlOG

DISCRETE

ARINCS/TM

ARINC

CCDl

SIGNAlChECK

RANGEChECK





Page May 0

FAUlT DETECTION

Channel selection determination

each eCU channel determines whether to be in the active state, or the standby state, based on a comparison of its health with that of the cross channel.

The channel with the better health status becomes the active channel. When both channels are of equal health, the channel selection state remains as the previous selection state.

a hierarchy is assigned to the list of possible faults that could lead to a channel switch.

When a single fault occurs, the channel with lower priority faults (if any) becomes active. If the same equal priority fault(s) exist on both channels, no switching occurs.

The internal logic of the eCU ensures that each channel achieves an active status on an alternating basis. an nVM flag is assigned to identify the channel that is presently active.

The last-active flag is only set when n becomes less than % speed.

The nVM last-active flag becomes the lowest priority status in channel health determination.

The occurrence of any higher priority faults overrides the last-active flag to ensure the healthiest channel is made the active channel.





Page May 0

FAUlT hIERARChYCTC--0-00

No hEAlTh STATUS DISCRETE

DESCRIPTION

1 GROUP 1 FAUlT (SERIOUS ECU INTERNAl FAUlT) 2 FMV lOOP FAUlT 3 VSV lOOP FAUlT 4 VBV lOOP FAUlT 5 lATChED ChANNEl CCDl SERIAl FAUlT 6 SAV WRAPAROUND FAUlT 7 TRPV WRAPAROUND FAUlT 8 TR INDETERMINATE STATE FAUlT 9 TRDV WRAPAROUND FAUlT10 BSV WRAPAROUND FAUlT **11 SPARE12 FRV 1 / FRV 2 WRAPAROUND FAUlT13 hPTC WRAPAROUND FAUlT14 TBV WRAPAROUND FAUlT *15 lPTC WRAPAROUND FAUlT16 NVM FAUlT17 ARINC OUTPUT WRAPAROUND FAUlT18 AlTERNATOR WINDING FAUlT19 lAST ACTIVE ChANNEl FlAG

* -5B SAC ONlY** -5A ONlY





Page 0May 0

ChANNEl SElECTION

Output driver disconnect

Once the active channel is determined, each channel executes the output driver disconnect logic to assign the respective active status for the disconnect relays.

The standby channel disconnects all its torque motor and solenoid output drivers from the external loads.

With a normal healthy status (no faults), all the assignments are connected in the active channel. Some driver output assignments are switched through paired disconnect relays.

The respective assignments are:- BSV & SaV (K) (BSV On -a OnlY)- TrPV & FrV (K)- VBV (K)- TrDV & lPTC (K)- VSV (K)- FMV (K)- HPTC & TBV (K) (TBV On -B OnlY)

If there is a failure on the active channel, the disconnect relays of the functions that are faulty are opened to prevent damage to the engine.





Page May 0

OUTPUT DRIVER - NO CONTROl - FAIlSAFE POSITIONCTC--00-00

NO CONTROl

BSV ** OPEN

SAV ClOSED

TRPV ClOSED

FRV1 ClOSED

VBV OPEN

TRDV OPEN

lPTC ClOSED

VSV ClOSED

FMV ClOSED

hPTC ClOSED

TBV * ClOSED

NO CURRENTFAIlSAFE POSITION

OUTPUTDATABUS

ACTIVEChANNEl

OUTPUTDATABUS

STAND-BYChANNEl

K1 BSV ** / SPARESOlENOID DRIVERK1 SAVSOlENOID DRIVERK2 TRPVSOlENOID DRIVERK2 FRV1SOlENOID DRIVERK3 VBVTORQUE MOTOR DRIVERK4 TRDVSOlENOID DRIVERK4 lPTCTORQUE MOTOR DRIVERK5 VSVTORQUE MOTOR DRIVERK6 FMVTORQUE MOTOR DRIVERK7 hPTCTORQUE MOTOR DRIVERK7 TBV * / SPARE

BSV ** / SPARESAV (FAUlT DETECTED)TRPV

FRV1

VBV

TRDV

lPTCVSVFMVhPTCTBV * / SPARE

TORQUE MOTOR DRIVER

K1SOlENOID DRIVERK1SOlENOID DRIVERK2SOlENOID DRIVERK2SOlENOID DRIVERK3TORQUE MOTOR DRIVERK4SOlENOID DRIVERK4TORQUE MOTOR DRIVERK5TORQUE MOTOR DRIVERK6TORQUE MOTOR DRIVERK7TORQUE MOTOR DRIVERK7TORQUE MOTOR DRIVER

CCDl

* -5B ONlY** -5A ONlY






Page May 0




engIne InDICaTIng SYSTeM

FAUlT DETECTION& ANNUNCIATION

Page May 0

ENGINE INDICATING SYSTEM





engIne InDICaTIng SYSTeM


Page May 0




WarnIngInDICaTIOnS


Page May 0

WARNING INDICATIONS




WarnIngInDICaTIOnS


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WARNING INDICATIONS

Depending on the data transmitted from the engine, messages are generated on the:

- Upper eCaM: engine Warning Display (eWD).- lower eCaM: Systems Display (SD).- Master caution, or warning.- audible chimes and oral warnings.- Status box ( STS) displayed after landing- advisory box (aDV) may appear to draw pilots

attention during flight.

These messages are used to run the engine under normal conditions throughout the operating range, or to provide warning messages to the crew and maintenance personnel.

The aircraft computers that impact the engine are:- System Data acquisition Concentrators (SDaC).- Display Monitoring Computers (DMC).- Flight Warning Computers (FWC).- engine Interface Units (eIU).- engine Vibration Monitoring Unit (eVMU).




WarnIngInDICaTIOnS


Page May 0

ENGINE INDICATING SYSTEMCTC--0-0

ECU

RED INDICATION

AMBER INDICATION

INDICATION PUlSINGGREEN OR AMBER

DIGITS PARTIAllYDAShED

All DIGITSDAShED

STATUSES

ADVISORIES

AMBER MESSAGES

RED MESSAGES

ACCElEROMETER 1ACCElEROMETER 2N1 SPEEDN2 SPEED

OIl QUANTITYOIl PRESSUREFUEl FIlTER ClOGOIl FIlTER ClOG

N1 SPEEDN2 SPEEDEGTFUEl FlOWT/R SWITChESSAV SWITChESIGNITION

NACEllE TEMP.OIl TEMP.OIl QUANTITYOIl PRESSURE SWITCh

UPPERECAM

lOWERECAM

DMC 3DMC 2

DMC 1

EIU 2

EIU 1

FWC 2

FWC 1

SDAC 2

SDAC 1

EVMU

104.1

105

726726

105

15

100

0

105

101.44

1410

STS

aDVENG 1 N1 OVERLIMITTHR LEVER 1 BELOW LIMITENG 1 N1 OVERLIMITTHR LEVER 1 BELOW LIMIT

ENG 1 OIL LOW PR THROTTLE 1 IDLEENG 1 OIL LOW PR THROTTLE 1 IDLE

MASTERWARN

MASTERCAUT DONG

DINGDING

DINGDING

SINGlEChIME

CONTINUOUSChIME




WarnIngInDICaTIOnS


Page May 0

WARNING INDICATIONS

Advisory status information

STS and aDV message indications may appear on the eCaM upper display to draw the pilots attention.

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 System page is displayed automatically and the affected parameter pulses. an advisory may, or may not, lead to a failure.

The aDV parameters are:

- egT above MCT limit - Oil high pressure > upper limit - Oil low pressure limit < lower limit - Oil quantity low limit < lower limit - nacelle temp > upper limit - Vib n above limit - Vib n above limit.

Status informations

The status box comes on when the aircraft is back on ground with class failures. The box is displayed after 0 knots plus 0 seconds.




WarnIngInDICaTIOnS


Page May 0

ADVISORY STATUS INFORMATIONCTC--0-00

2100 2100

EGT0C

F.FKG/HKG/H

95.00 95.0N2%%

105N1%%

FOB:16300KG16300KG57510

5

575

105

105

STS

SEAT BELTSNO SMOKINGSEAT BELTSNO SMOKING

ENG A ICEENG A ICE

89.0%0%CL

2100 2100

EGT0C

F.FKG/HKG/H

95.00 95.0N2%%

105N1%%

FOB:16300KG16300KG57510

5

575

105

105

aDVSEAT BELTSNO SMOKINGSEAT BELTSNO SMOKING

ENG A ICEENG A ICE

89.0%0%CL

STATUS REMINDERCOMES AFTER lANDINGAT 80 kt + 30 SECONDS

ADVISORY ADV REMINDERCOMES IN REAl TIME DURINGFlIGhT AND RElEVANT ECAMPAGE COMES INTO VIEW.ADVISORY INDICATION ISPUlSING TO DRAW CREWSATTENTION.

EGTOIl hIGh PRESSURE ABOVE MCT lIMITOIl lOW PRESSURE lIMIT < lOWER lIMITOIl QUANTITY lOW lIMIT < lOWER lIMITNACEllE TEMP > UPPER lIMITVIB N1 ABOVE lIMITVIB N2 ABOVE lIMIT

89.0 89.0 89.0 89.0




WarnIngInDICaTIOnS


Page 0May 0

SYSTEM AND STATUS DISPlAY

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 displayed at the bottom. The right part indicates the inoperative systems and the maintenance status. On the lower part of the status and system page, some data are displayed.

nOTe:When the STaTUS page disappears, a white STS mes-sage appears on the engIne and WarnIng display to indicate that the STaTUS page is no longer empty.

When pressed the STaTUS page is displayed. If no STaTUS message is present the nOrMal message is displayed for seconds.

Overflow low arrow: only concerns the warning messages and indicates that the messages exceed the capacity of the display on the left Memo area. In this case, the hea-ding titles of the warning messages are displayed on the right Memo area.

Automatic display

The status page contain an operational summary of the aircraft condition. This page is automatically called when slats > (in approach).




WarnIngInDICaTIOnS


Page May 0

STATUS PAGECTC--0-00

66000GW KG5623 H

+19TAT C+18SAT C

APPR PROC DUAL HYD LO PR

-IF BLUE DVHT OUT=-BLUE ELEC PUMP...AUTO-IF GREEN OVHT OUT: -GREEN ENG 1 PUMP..ON -PTU.............AUTO

-L/G........GRVTY EXTN-LDG SPD INCREN...10 KT-LDG DIST........X 1.2

CAT 1 ONLYFLAP SLOW

CANCELLED CAUTION

NAV IR 2 FAULT

STATUSINOP SYS

G+B HYDCAT 2B ELEC PUMPG RSVRL+R AILSPLR 1+3+5L ELEVFLATSAP 1+2ENG 1 REVNORM BRKNV STEER

MAINTENANCE

APUAIR COND

INOP SYS

G+B HYDCAT 2B ELEC PUMPG RSVRL+R AILSPLR 1+3+5L ELEVFLATSAP 1+2ENG 1 REVNORM BRKNV STEER

MAINTENANCE

APUAIR COND

APPROACH PROCEDURES

PROCEDURES

LIMITATIONS

CANCELLED CAUTION

STATUS

INOPERATIVE SYSTEMS

MAINTENANCE

STATUS PAGE CONTENT

STATUS PAGE SAMPlE: ENGINE ThRUST REVERSER INOPERATIVE

SYSTEMS DISPlAY BUTTOM ON ThE CENTRAl CONSOlE

OT OT OT OT OT OT OT OT OT OT OT OT OT OT OT OT OT OT OT OT OT OT OT OT OT OT OT OT OT OT OT OT OT OT OT OT OT OT OT OT OT OT OT OT OT.OT.OT.OT.OT.OT.OT.OT.OT.OT.OT.OT.OT.OT.OCO GCONFIGCONFIGCONFIGCONFIGCONFIGCONFIGCONFIGCONFIGCONFIGCONFIGCONFIGCONFIGCONFIGCONFIGCONFIGCONFIGCONFIGCONFIGCONFIGCONFIGCONFIGCONFIGCONFIGCONFIGCONFIGCONFIGCONFIGCONFIGCONFIGCONFIGCONFIGCONFIGCONFIGCONFIGCONFIGCONFIGCONFIGCONFIGCONFIGCONFIGCONFIGCONFIGCONFIGCONFIGCONFIGCONFIGCONFIGCONFIGCONFIGCONFIGCONFIGCONFIGCONFIGCONFIGCONFIGCONFIGCONFIGCO G

EREREMERERERCANCCACACACA

ALLALLALLALLALLALLALLALLALLALLALLALLALLALLALLALLALLALLALLALLALLALLALLALLALLALLALLALLALLALLALLALLALLALLALLALLALLALLALLALLALLALLALLALLALLALLALLALLALLALLALLALLALLALLALLALLALLALL

STS CRCLRCLRCLRCLRCLRCLRCLRCLRCLRCLRCLRCLRCLRCLRCLRCLRCLRCLRCLRCLRCLRCLRCLRCLRCLRCLRCLRCLRCLRCLRCLRCLRCLRCLRCLRCLRCLRCLRCLRCLRCLRCLRCLRCLRCLRCLRCLRCLRCLRCLRCLRCLRCLRCLRCLRCLRCLC




WarnIngInDICaTIOnS


Page May 0

ENGINE INDICATING FAUlTS - N1

Upper display N1 indications

The n indication is displayed on the upper display unit of the eCaM system:

- In analog form, by a pointer deflecting in front of a dial,

- In digital form, in the lower section of the dial.

The indication is normally green.

Warning indication

When n lIM < n < red line: - The indication becomes amber.

When n = red line: - The indication becomes red. - The MaSTer CaUT light comes on accompanied

by a single chime.

The following messages appear on the eCaM:

- eng () n OVerlIMIT - THr leVer ()...BelOW lIMIT

If fan speed reaches the red line n, a small red line remains positioned on the analog scale at that value (max pointer). The max pointer is reset at next Master lever On.




WarnIngInDICaTIOnS


Page May 0

ENGINE INDICATING FAUlTS - N1CTC--0-00

84.6% 3566%%%% 353500CC

955 977

EGT0C

F.FKG/HKG/H

99.88 99..9N2%%

105N1%% FLX

FFOB:163001630016300KGKG

FLAPS F

2

670

105

670

105

104.0

105

ENG 1 N1 OVERLIMITTHR LEVER 1 BELOW LIMITENG 1 N1 OVERLIMITTHR LEVER 1 BELOW LIMIT

101.4

A/C ID DATE GMT FLNT CITY PAIR.F-GFXU 20JUN 1842 AF2706 LFPO LFPN




GMT PH ATA0820 05 73-21 ENG 1 N1 OVERLIMITGMT PH ATA0920 06 73-21 THR LEVER 1 BELOW LIMIT





GMT PH ATA0820 05 73-21 ENG 1 N1 OVERLIMITGMT PH ATA0920 06 73-21 THR LEVER 1 BELOW LIMIT

IF N1 RED lINE INDICATION BECOMES RED>

DONGMASTER

CAUT




WarnIngInDICaTIOnS


Page May 0

ENGINE INDICATING FAUlTS - N1


Warning indication

Upper display n indications: in case of failure of the normal n measurement system from the FWC, the eCU computes a theoretical value taking into account the other engine parameters. In this case the last digit of n is dashed with two horizontal amber lines.

In the event of a total failure of parameters to the Display Monitoring Computer (DMC), the eCaM indicates amber crosses to show that no value is available. This display is also used when the aircraft system is powered up and eCU power is off. In this case it is normal.




WarnIngInDICaTIOnS


Page May 0




A/C ID DATE GMT FLTN CITY PAIR.F-GFXU 20JUN 1842 AF2706 LKFB LFML


GMT PH ATA0920 06 73-21 ENG 1 N1 SENSOR FAULT

FAILURE MESSAGES

GMT PH ATA SOURCE IDENT.0920 06 73-21-10 N1 SNSR, ECU1B J9, ECUGMT PH ATA0920 06 73-21-10 N1 SNSR, ECU1B J10, ECU



A/C ID DATE GMT FLTN CITY PAIR.F-GFXU 20JUN 1842 AF2706 LKFB LFML


GMT PH ATA0920 06 73-21 ENG 1 N1 SENSOR FAULT

FAILURE MESSAGES

GMT PH ATA SOURCE IDENT.0920 06 73-21-10 N1 SNSR, ECU1B J9, ECUGMT PH ATA0920 06 73-21-10 N1 SNSR, ECU1B J10, ECU

84.6% 3566%%%% 353500CC

955 977

EGT0C

F.FKG/HKG/H

99.88 99..9N2%%

N1%% FLX

FFOB:163001630016300KGKG

FLAPS F

2

670

105

670

105

ENG 1 N1 SENSOR FAULTENG 1 N1 SENSOR FAULT

lAST DIGIT AMBER & DAShEDECU MAThEMATICAl MODElDISPlAYED

DONGMASTER

CAUT

AMBER CROSSES WhEN NO VAlUEAVAIlABlE MIGhT BE NORMAl, I.E.ECU POWER IS OFF

105

101.44

105

XX XX




WarnIngInDICaTIOnS


Page May 0

WARNING INDICATION


If n has exceeded the red line: the maximum value reached is memorized with the corresponding time in overcondition. These values are available on the MCDU page Power setting Max values.

above the overspeed limit, the message displayed on the eCaM becomes:

- eng () n OVerlIMIT - THr leVer ()...IDle - eng MaSTer ()OFF




WarnIngInDICaTIOnS


Page May 0


84.6% 3566%%%% 353500CC

955 977

EGT0C

F.FKG/HKG/H

99.88 99..9N2%%

N1%% FLX

FFOB:163001630016300KGKG

FLAPS F

2

670

105

670

105

ENG 2 N1 OVERLIMITTHR LEVER 2...IDLEENG MASTER 2...OFF

ENG 2 N1 OVERLIMITTHR LEVER 2...IDLEENG MASTER 2...OFF





GMT PH ATA0820 05 73-21 ENG 2 N1 OVERLIMITGMT PH ATA0820 06 73-21 THR LEVER 2...IDLEGMT PH ATA0820 06 73-21 ENG MASTER 2...OFF





GMT PH ATA0820 05 73-21 ENG 2 N1 OVERLIMITGMT PH ATA0820 06 73-21 THR LEVER 2...IDLEGMT PH ATA0820 06 73-21 ENG MASTER 2...OFF

DONGMASTER

CAUT

105

104.1

105

105.9




WarnIngInDICaTIOnS


Page May 0

WARNING INDICATIONS

Upper display - EGT indications

The actual egT indications are normally green.

The index pulses amber above egT MCT max value (or above start limit during the start sequence).




WarnIngInDICaTIOnS


Page May 0

ENGINE INDICATING FAUlTS - EGT AMBER WARNINGSCTC--0-00

aDV

84.6% 3566%%%% 353500CC

955 977

EGT0C

F.FKG/HKG/H

99.88 99..9N2%%

N1%% FLX

FFOB:163001630016300KGKG

FLAPS F

2

670

105

726726

105

ENG 1 START FAULT-EGT OVERLIMITENG 1 START FAULT-EGT OVERLIMIT





GMT PH ATA0720 02 73-21 ENG 1 START FAULT-EGT OVERLIMIT






DONGMASTER

CAUT105 0

101.4 70.4

105

EGT0C 670

105

915915

105

- 5B

EGT lIMIT IN STARTING SEQUENCE

EGT ADVISORY VAlUE





WarnIngInDICaTIOnS


Page 0May 0




WarnIngInDICaTIOnS


Page May 0

ENGINE INDICATING FAUlTS - EGT AMBER WARNINGSCTC--0-00

aDV

84.6% 3566%%%% 353500CC

955 977

EGT0C

F.FKG/HKG/H

99.88 99..9N2%%

N1%% FLX

FFOB:163001630016300KGKG

FLAPS F

2

670

105

726726

105

ENG 1 START FAULT-EGT OVERLIMITENG 1 START FAULT-EGT OVERLIMIT

A/C ID DATE GMT FLNT CITY PAIR.F-GFXV 20JUN 1842 AF2706 LFPO LFPN


A/C ID DATE GMT FLTN CITY PAIR.F-GFXV 20JUN 1842 AF2706 LFPO LFPN



A/C ID DATE GMT FLNT CITY PAIR.F-GFXV 20JUN 1842 AF2706 LFPO LFPN


A/C ID DATE GMT FLTN CITY PAIR.F-GFXV 20JUN 1842 AF2706 LFPO LFPN



DONGMASTER

CAUT105 0

101.4 70.4

105

EGT0C 670

105

855855

105

- 5A

EGT lIMIT IN STARTING SEQUENCE

EGT ADVISORY VAlUE




WarnIngInDICaTIOnS


Page May 0

WARNING INDICATIONS

Upper display - EGT indications

The index and the numerical value pulse red above the egT red line.

Documents

CTC-232_FDA