cfm 56-5b diff l3 (dec2003 cmp)

ATA 104 SPEC L3

Book No: CFM 56-5B DIFF L3

Lufthansa

Lufthansa BaseFor Training Purposes Only

Lufthansa 1995Technical Training GmbH

Training ManualA318

ATA 71-80Engine CFM 56-5B

Diff to CFM 56-5A

For training purposes and internal use only.

Copyright by Lufthansa Technical Training GmbH.

All rights reserved. No parts of this trainingmanual may be sold or reproduced in any formwithout permission of:

Lufthansa Technical Training GmbH

Lufthansa Base Frankfurt

D-60546 Frankfurt/Main

Tel. +49 69 / 696 41 78

Fax +49 69 / 696 63 84

Lufthansa Base Hamburg

Weg beim Jger 193

D-22335 Hamburg

Tel. +49 40 / 5070 24 13

Fax +49 40 / 5070 47 46

Lufth

ansa

Tec

hnic

al T

rain

ing

For

Tra

inin

g P

urpo

ses

Onl

y

POWER PLANTGENERAL

A318/319/A320/321CFM 56-5B

71-00

Page: 1FRA US/T Kh August 03 L1/L2

ATA 71 POWER PLANT

71-00 GENERALCFM 56 CONCEPT

The CFM 56 turbofan engine family is a product of CFMI (Comercial Fan MotorInternational). CFM International is a company jointly owned by General Elec-tric of the USA and Societe Nationale dEtude et de Construction de MoteursdAviation (SNECMA) of France.

- L.P. System - Accessory Drive System- Control & Accessories- Engine Installation- Thrust Reverser

- Core Engine- Fuel System Design- E.C.U. & H.M.U.

Page: 1

Lufth

ansa

Tec

hnic

al T

rain

ing

For

Tra

inin

g P

urpo

ses

Onl

y

POWER PLANTGENERAL

A318/319/A320/A321CFM 56-5B

71-00


CFM56-5B CONFIGURATIONS

Page: 2

Lufth

ansa

Tec

hnic

al T

rain

ing

For

Tra

inin

g P

urpo

ses

Onl

y

POWER PLANTGENERAL

A318/319/A320/A321CFM 56-5B

71-00


5 ST. LP COMPRESSOR(4 STAGE BOOSTER)

9 STAGE HPC ANNULAR COMBUSTOR(SAC OR DAC)

1 ST. HPT4 ST. LPT

FAN DIAMETER 68.3

Page: 3Figure 1 Engine Layout

Lufth

ansa

Tec

hnic

al T

rain

ing

For

Tra

inin

g P

urpo

ses

Onl

y

ENGINE FUEL AND CONTROLFADEC GENERAL

A319/A320/A321CFM56-5B

73-20

Page: 4FRA US/T Kh August 03 L3

FADEC FUNCTIONS

Full Authority Digital Engine Control ( FADEC )The FADEC consists of the Engine Control Unit ( ECU ), Hydromechanical Unit( HMU ) and its peripheral components and sensors used for control and moni-toring.

FADEC DefinitionEach engine is equipped with a duplicated FADEC system. The FADEC acts asa propulsion system data multiplexer making engine data available for conditionmonitoring.

FADEC ControlsThe FADEC provides the engine sytem regulation and scheduling to control thethrust and optimize the engine opration.The FADEC provides:

- Fuel control regulation- power management- gas generator control- Turbine active clearance control- engine limit protection- feedback- flight deck indication data- Engine maintenance data- Contitioning monitoring data- thrust reverse control- automatic engine starting- Fuel return control for IDG cooling

Power ManagementThe FADEC provides automatic engine thrust control and thrust parameter lim-its computation.The FADEC manages power according to two thrust modes:

- manual mode depending on thrust lever angle ( TLA )

- Autothrust mode depending on autothrust function generated by the autoflight system ( AFS ).

The FADEC also provides two idle mode selections:- Approach Idle: It is obtained when slats are extended in FLT.- Minimum Idle: It can be modulated up to approach idle depending on:

Air conditioning demand Engine anti ice demand Wing anti ice demand Temperature Engine Oil TEO (for IDG cooling ).

Engine Limit ProtectionThe FADEC provides overspeed protection for N1 and N2, in order to preventengine exceeding certified limits, and also monitors the EGT.

Engine Systems ControlThe FADEC provides optimal engine operation by controlling the:

- Fuel Flow- Compressor air flow and- Turbine clearence.

Thrust ReverseThe FADEC supervises entirely the thrust reverse operation.In case of a malfunction, the thrust reverser is stowed.

Start and Ignition ControlThe FADEC controls the engine start sequence. It monitors N1, N2 and EGTparameters and can recycle or abort an engine start.

Power SupplyThe FADEC system is self-powered by a dedicated permanent magnet alter-nator when N2 is above 12%, and is powered by the aircraft electrical systemfor starting, as a backup and for testing with engine not running.

Page: 4

Lufth

ansa

Tec

hnic

al T

rain

ing

For

Tra

inin

g P

urpo

ses

Onl

y

ENGINE FUEL AND CONTROLFADEC GENERAL

A319/A320/A321CFM56-5B

73-20


T5P25

FOR ENGINE TREND MONITORING

Ps13

THRUSTLEVER

TRUST CONTROLUNIT

RESOLVER

T/R REVERSER Stow / Deploy Feedback

T/R REVERSER Stow / Deploy Command

HCU

IgnitionBoxes

A B

ECU( CH: A & B )

28 V DC115 V

400 HZ

ECU ALTERNATOR

FUELFLOW

IGN A

IGN B

Thrust Reverser

FEEDBACK

FUEL PRESS

HMU

T-CASETEO

P 0 T25

N1 N2

T12 PS12 PS3T49.5 T3 FMV

FUEL RETURNVALVE

FEEDBACKReturn Fuel to AC Tank

IGNITORS

HYDRAULICPRESS

ANALOG &DISCRETESIGNALS

FUEL FLOWTO

BURNERS

CFM 56-5B

FEEDBACK(EGT)

TB

V

OPTIONRACC/SB

TBV/NAC

TBV

Page: 5Figure 2 FADEC Presentation CFM 56-5B

Lufth

ansa

Tec

hnic

al T

rain

ing

For

Tra

inin

g P

urpo

ses

Onl

y

ENGINEGENERAL

A319/A320/A321CFM56-5B

72-00

Page: 6FRA US-T Kh August 03 L2

ENGINE STATION DESIGNATIONStation numbers are assigned to identify specific positions along the aerody-namic flowpath of an engine.A station is a position at the engine, where thermodynamically changes(Pressure, temperature or airspeed) starts or ends.

Engine stations are labeled continuously from 1 to 5 along the aerodynamicflowpath.The station numbers are also used to identify instrumentation positions forpressure and temperature sensors. Temperature sensors are labeled with T,pressure sensors are labeled with a P, followed by a station number.

Page: 6

PRIMARY AIRFLOW

SECONDARY AIRFLOW

STATIONS OF

STATIONS OF

COMPRESSOR DISCHARGEPRESSURE AND TEMPERATURE

P3 / T3

LP TURBINE EXIT

TEMPERATURE

T5

HPC INLET PRESSURE

AND TEMPERATURE

P25 / T25

FAN INLET PRESSURE

AND TEMPERATURE

T12 / P12

FAN OGV OUTLET

PRESSUREP13

EXHAUST GAS

TEMPERATURET49.5

P0

AMBIENT PRESSURE

Lufth

ansa

Tec

hnic

al T

rain

ing

For

Tra

inin

g P

urpo

ses

Onl

y

ENGINEGENERAL

A319/A320/A321CFM56-5B

72-00


FM

V

VS

V/V

BV

HP

TC

C

LPT

CC

RA

CC

/SB

/TB

V

BS

V

STA

RT

/ I

GN

CO

N M

ON

EC

AM

01

02

04

03

05

06

07

08

09

10

11

12

T 12

T 25

T 3.0

T 4.9

T 5.0

PS 13

P 25

PS 3.0

PO

PS12

N1

N2

USEDFOR

A B C D E F G H ILEGENDE (HORIZONTAL)A FUEL METERING VALVEB VARIABLE STATOR VANES / VARIABLE BLEED VALVESC HIGH PRESSURE COMPRESSOR TURBINE CLEARANCE CONTROLD LOW PRESSURE COMPRESSOR TURBINE CLEARANCE CONTROL

E ROTOR ACTVE CLEARANCE CONTROL /START BLEED OR TRANSIENT BLEED VALVE CONTROLF BURNER STAGING VALVEG AUTO START & IGNITIONH CONDITION MONITORINGI ELECTRONIC CENTRALIZED AIRCRAFT MONITORING

01 ELECTRICAL FAN INLET TEMPERATURE SENSORS (2)02 HIGH PRESSURE COMPRESSOR INLET TEMPERATURE SENSOR03 HIGH PRESSURE COMPRESSOR DISCHARGE TEMPERATURE SENSOR04 EXHAUST GAS TEMPERATURE SENSOR (EGT)05 LOW PRESSURE TURBINE EXHAUST GAS TEMPERATURE SENSOR06 FAN EXIT PRESSURE SENSOR07 HIGH PRESSURE COMPRESSOR INLET PRESSURE SENSOR08 HIGH PRESSURE COMPRESSOR DISCHARGE PRESSURE SENSOR 09 AMBIENT PRESSURE (PO)10 ENGINE FAN INLET PRESSURE11 LOW PRESSURE COMPRESSOR ROTOR SPEED SENSOR (N1)12 HIGH PRESSURE COMPRESSOR ROTOR SPEED SENSOR (N2)13 FUEL FLOW TRANSMITTER

14 T-CASE TEMPERATURE SENSOR

LEGENDE (VERTICAL)

13 FF

14 TCASE

Page: 7Figure 3 Aerodynamic Stations

Lufth

ansa

Tec

hnic

al T

rain

ing

For

Tra

inin

g P

urpo

ses

Onl

y

ENGINESTATIONS

A318/319/A320/321CFM56-5B

72-00

Page: 8FRA US/T Kh August 03L2/3

STAGE NUMBERING CFM56-5B

STAGES : COMPONENT : STAGE NUMBER : NOTES :

1 FAN 1 Fan air used for ACC

123

4

LOW PRESSURE COMPRESSOR (BOOSTER)

123

4 VBV

123456789

HIGH PRESSURE COMPRESSOR

123456789

( IGV )VSV VSVVSV

HPT ACC

CUST. BLEED, Eng. Anti Ice (A/I),

CUST. BLEED, Muscle Press A/IStart Bleed, HPT ACC

COMBUSTION CHAMBER 20 Fuel Nozzles,2 Ignitor Plugs

1 HIGH PRESSURE TURBINE

1 ACTIVE CLEARANCE CONTROL

1234

LOW PRESSURE TURBINE

1234

ACTIVE CLEARANCE CONTROL

EXHAUST NOZZLE

Page: 8

Lufth

ansa

Tec

hnic

al T

rain

ing

For

Tra

inin

g P

urpo

ses

Onl

y

ENGINESTATIONS

A318/319/A320/321CFM56-5B

72-00

Page: 9FRA US/T Kh August 03L2/3

5 ST. LP COMPRESSOR(4 STAGE BOOSTER)

9 STAGE HPC ANNULAR COMBUSTOR(SAC OR DAC)

1 ST. HPT4 ST. LPT

FAN DIAMETER 68.3

Page: 9Figure 4 Engine Cross Section

Lufth

ansa

Tec

hnic

al T

rain

ing

For

Tra

inin

g P

urpo

ses

Onl

y

ENGINEFAN SECTION

A318/A319/A320/A321CFM 56-5B

72-20


SPINNER FRONT CONE

Interference fit and single annular mounting positions are characteristics of theinstallation of the front and rear cone onto the fan disk.The spinner front cone has an offset hole on its rear flange, identified by anindent mark, to ensure correct alignment for installation onto the rear cone frontflange.The rear flange has 6 mounting screw locations and 3 threaded inserts,located every 120, for installation of jackscrews used in removal procedures.

SPINNER REAR CONE

The front flange of the spinner rear cone has 6 line replaceable, crimped, self-locking nuts.The inner rear flange has 12 mounting screw holes, for installation onto the fandisk, and there are a further 6 threaded holes for the installation of jackscrewsused in rear cone removal procedures.

Both front and rear flanges have an offset hole to ensure correct installationand they are identified by indent marks.The rear cone also has an integrated air seal that is glued to its inner rearflange.The rear cone prevents axial disengagement of spacers used in the fan bladeretention system.lt also supports a series of balancing screws that are installed on its outer di-ameter. There are two sets of balancing screws available and the screws ineach set are identified as either PO1 to P07 or, P08 to P14. The numbers,which are engraved on the screw heads, are equivalent to various weights.An indent mark is located in between two balancing screws for correct installa-tion of the rear cone onto the fan disk and for identification of fan blade No.1.

SPINNER REAR CONE

Page: 10

Lufth

ansa

Tec

hnic

al T

rain

ing

For

Tra

inin

g P

urpo

ses

Onl

y

ENGINEFAN SECTION

A318/A319/A320/A321CFM 56-5B

72-20


8 MOUNTING SREWS

LOCATION

OFFSET HOLE

SPACER

FAN BLADE

3 JACKSREW LOCATIONS

Page: 11Figure 5 SPINNER CONES

Lufth

ansa

Tec

hnic

al T

rain

ing

For

Tra

inin

g P

urpo

ses

Onl

y

ENGINEFAN FRAME MODULE

A318/A319/A320/A321CFM 56-5B

72-00


FAN INLET CASE

The inner surface of the upstream fan inlet case is lined with 6 forward acousti-cal panels, 6 mid acoustical panels and provides an abradabie shroud whichfaces the fan blade tips.The inner surface of the downstream fan inlet case is lined with 12 aft acousti-cal panels.The fan inlet case also houses the Outlet Guide Vane (OGV) assembly.

FAN FRAMEThe fan frame has 12 radial hollow struts that house various equipment andlines. Compartments formed between the adjacent struts house Variable BleedValve (VBV) actuators which, under certain conditions, redirect primary air intothe secondary airflow.The rear face of the fan frame mid section provides the front mounts for theengine and the front flange for the High Pressure Compressor (HPC) section.

THE OUTLET GUIDE VANE (OGV) ASSEMBLYThe fan OGV assembly consists of the inner shroud and 35 twin vanes.The inner shroud rear flange is bolted to the fan frame and has 35 apertures toallow passage of the vane inner platforms.The vane inner platforms are axially retained by the inner face of the fan OGVinner shroud.The vane outer platforms are bolted to the downstream fan inlet case.A splitter fairing, which separates the primary and secondary airflows, is boltedonto the fan OGV inner shroud forward flange.There are 2 unplugged holes on the inner shroud, between the 3 and 4 oclockpositions, to enable borescope inspection of the booster vane assemblies. Oneis located between the OGVs at the stage 3 vane assembly and the other atthe stage 5 vane assembly.

Page: 12

Lufth

ansa

Tec

hnic

al T

rain

ing

For

Tra

inin

g P

urpo

ses

Onl

y

ENGINEFAN FRAME MODULE

A318/A319/A320/A321CFM 56-5B

72-00


FANFORWARDACOUSTICALPANEL

ABRADABLESHROUD

FAN MIDACOUSTICALPANEL

FANFRAMECASING

FAN AFTACOUSTICALPANEL

STRUT

VBV ACTUATOR

Page: 13Figure 6 Fan Inlet Case / Fan OGV

Lufth

ansa

Tec

hnic

al T

rain

ing

For

Tra

inin

g P

urpo

ses

Onl

y

ENGINEGENERAL

A318/A319/A320/A321CFM56-5B

72-20


FAN BLADE / FRONT AND REAR CONE

Fan BladesThere are 36 titanium alloy, mid-span shrouded, fan blades approximately 25inches ( 630 mm ) long. Each blade has a dovetail base that engages in adovetail recess on the disk rim. A spacer limits the radial movement of eachblade. A retainer lug, machined in the rear end of blade root, engages the for-ward flange of the booster spool and limits the forward and rearward axialmovements.

Spinner Front ConeThe spinner front cone is made of composite material. Its design precludes theneed for an engine nose anti-icing system. The front cone is bolted to the rearcone.

Spinner Rear ConeThe spinner rear cone is made of aluminum alloy. Its rear flange is bolted to thefan disk and is part of the fan blades axial retention system. The outer rim ofrear flange is provided with tapped holes for trim balance screws. The frontflange provides attachment of the spinner front cone.

Fan DiskThe fan disk is a titanium alloy forging. Its inner rear flange provides attach-ment for the fan shaft and its outer rear flange is bolted to the booster rotor.The outer front flange provides attachment for the spinner rear cone. The diskouter rim has 36 recesses designed for fan blade retention.

Page: 14

Lufth

ansa

Tec

hnic

al T

rain

ing

For

Tra

inin

g P

urpo

ses

Onl

y

ENGINEGENERAL

A318/A319/A320/A321CFM56-5B

72-20


Spinner Rear Cone

Spinner Front Cone

Bolt

Bolt

Washer

Fan Blades

Fan Disk

Booster Spool

Booster SpoolFront Flange

Retention Slot

Fan DiskSpacerSpinner Rear ConeMounting Flange

Retention Lug

Balancing Screws

1ST

STEP

2ND

STEP

Page: 15Figure 7 Fan Blade Retention

Lufth

ansa

Tec

hnic

al T

rain

ing

For

Tra

inin

g P

urpo

ses

Onl

y

ENGINEFAN SECTION

A319/A320/A321CFM 56-5B

72-20


FAN DISK

The fan disk outer rim has 36 dovetail recesses for the installation of the fanblades.The inner front flange has an imprint to identify an offset hole for rear coneinstallation.There are also two identification marks engraved on either side of blade re-cesses No.1 and 5.

FAN BLADESThere are 36 titanium alloy, mid-span shrouded fan blades.Each blade has a dovetail base that slides into a recess on the fan disk outerrim.A retainer lug, machined at the rear end of the blade root, engages the forwardflange of the booster spool and limits axial movements.A spacer, installed underneath each blade, limits the radial movement.

Page: 16

Lufth

ansa

Tec

hnic

al T

rain

ing

For

Tra

inin

g P

urpo

ses

Onl

y

ENGINEFAN SECTION

A319/A320/A321CFM 56-5B

72-20


OFFSET HOLE

SPHERICAL

IMPRINT

Page: 17Figure 8 Fan Disk and Blades

Lufth

ansa

Tec

hnic

al T

rain

ing

For

Tra

inin

g P

urpo

ses

Onl

y

ENGINEFAN SECTION

A318/A319/A320/A321CFM 56-5B

72-20

Page: 18FRA US/TKh August 03 L3

FAN BLADES (CONTINUED)

Each blade has specific indications engraved on the bottom of the root:- Part number- Serial number- Momentum weight- Manufacturer code

The fan blade root pressure faces are sprayed with an anti-friction plasmacoating, and dry molybdenum base film is applied to the blade roots.The fan blade mid-span shroud contact surfaces are coated with tungsten-car-bide.

COLD FAN BALANCING

Engine Fan Trim BalanceTrim balance is a procedure used to reduce the engine vibration level. This pro-cedure must be applied every time the engine vibration level reaches 8.0 mils,which leads to rapid loss of the EGT margin, every time the engine vibrationresult in significant cabin noise, or after an engine check.This procedure is known as Cold Trim Balance because it consists in correctingthe imbalance on a cold engine without the need for successive ground runs todetermine the imbalance to be corrected.The imbalance is determined using the vibration parameters recorded in flightby the aircraft EVMU. This procedure allows the engine vibration level to bemaintained continuously, at minimal cost. The same calculation can be per-formed with the aircraft on the ground, after replacement of fan blades when astatic imbalance correction of 400 g.cm or more is necessary.For imbalance correction calculation, the cold trim balance procedure uses thein-flight recording of the vibration delivered by the engine No. 1 and TRF bear-ing vibration sensors.

Page: 18

Lufth

ansa

Tec

hnic

al T

rain

ing

For

Tra

inin

g P

urpo

ses

Onl

y

ENGINEFAN SECTION

A318/A319/A320/A321CFM 56-5B

72-20

Page: 19FRA US/TKh August 03 L3 Page: 19Figure 9 Fan Blade Root

Lufth

ansa

Tec

hnic

al T

rain

ing

For

Tra

inin

g P

urpo

ses

Onl

y

ENGINEGENERAL

A319/A320/A321CFM56-5B

72-00


HIGH PRESSURE COMPRESSOR

DescriptionThe major components of the High Pressure Compressor (HPC) are: one compressor rotor one compressor stator one compressor rear frame.

A portion of the fan discharge airflow passes through the booster and compres-sor. The front of the compressor stator is supported by the fan frame, and thecompressor rotor is supported by the No. 3 bearing in the fan frame. The rearof the compressor stator is supported by the combustion case forward flange,and the rear of the compressor rotor is supported by the No. 4 bearing in theturbine rear frame.Air, taken in through the fan and booster sections passes through successivestages of rotor blades and stator vanes, being compressed as it passes fromstage to stage. After passing through the 9 HPC stages, the air is fully com-pressed.The inlet guide vanes and the first 3 stages of the stator are variable, andchange their angular position as a function of compressor inlet temperature andengine speed. The purpose of this variability is to optimize efficiency and pro-vide improved stall margin.

The High Pressure Compressor (HPC) rotor.The High Pressure Compressor ( HPC ) rotor is a 9-stages, high-speed,spool-disk structure.The HPC rotor consists of 5 major parts: front shaft stage 1-2 spool stage 3 disk stage 4-9 spool rear Compressor Discharge Pressure ( CDP ) rotating air seal.

Spools are assembled by inertia welding. The front shaft, disk, and spools arejoined at a single bolted joint to form a rigid unit. Interfering rabbeted diametersare used for proper positioning of parts providing rotor balance stability.

Front ShaftThe front shaft which is bolted between the stage 1-2 spool and stage 3 disk,is the forward support for the rotor. The shaft is splined and secured to the InletGearbox ( IGB ) horizontal bevel gear by a coupling nut. The IGB contains thethrust anti-orbiting bearings for the core engine. This shaft is made of a tita-nium alloy.

Disk and SpoolsThe stage 1-2 spool and stage 3 disks are made of titanium alloy forgings andretain the blades in axial slots.The stage 4-9 spool is made of a nickel alloyand retains all blades in circumferential grooves.The rotor internal temperature is maintained below pressure compressor( booster ) discharge air which enters through holes in the front shaft and 5thstage air, provided by the Rotor Active Clearance Start Bleed ( RACSB ) valvethrough the fan frame. Labyrinth seals between the rotor blade stages improvecompressor performance.

BladesBlades in stages 1, 2 and 3 are made of titanium alloy. Blades in stages 4through 9 are made of nickel alloy. The first 3 stages of blades are secured inthe disk and spool axial slots by retaining rings; blades in the stages 4-9 spoolare secured in the circumferential grooves by locking lugs. All blades are re-placeable without disassembling the rotor.

Rear SealThe HPC rotor rear rotating ( CDP ) air seal is a one-piece nickel alloy forgedpart with abrasive, protective-coated labyrinth seals. The seal is mounted tothe aft flange of the stages 4-9 spool by a tight fitting rabbet diameter and isaxially clamped by the bolts and nuts securing the forward flange of the highpressure turbine rotor to the compressor.

Page: 20

Lufth

ansa

Tec

hnic

al T

rain

ing

For

Tra

inin

g P

urpo

ses

Onl

y

ENGINEGENERAL

A319/A320/A321CFM56-5B

72-00


IGV !. 2. 3. 4. 5.

6. 7. 8.

9.

Stage Number of Vanes

COMBUSTION CASE

4x 4th Stage Air for LPT Nozzle Guide Vane Cooling1x 4th Stage Air for HPT Clearance Control ( HPTCC )

1x 5th Air for HPC Clearance Con-trol ( RACC )2x 5th Stage Air for Inlet Anti Ice,Pneumatic System

4x 9th Stage Air for Pneumatic System1x 9th Stage Air for HPTCC1x 9th Stage Air for Start Bleed ( SB )

HPC ROTOR CDP SEAL DISK

8.

HPC FRONT SHAFT

1.

Page: 21Figure 10 High Pressure Compressor

Lufth

ansa

Tec

hnic

al T

rain

ing

For

Tra

inin

g P

urpo

ses

Onl

y

ENGINETURBINE SECTION

A318/A319 /A320 /A321CFM56-5B

72-50

Page: 22FRA US-T Kh August 03L3

HIGH PRESSURE TURBINE

GeneralThe turbine section provides the necessary power to drive the compressor ro-tors. lt consists of the

- High Pressure Turbine (HPT)- Low Pressure Turbine (LPT)

High Pressure Turbine (HPT)The HPT module is housed in the combustion case and consists of a singlestage nozzle that directs the gas flow from the combustion chamber to the HPTrotor blades that drive the HPC rotor.The LPT stage 1 nozzle is also housed in the combustion case.

Engine coolingAir from the 4th stage of the HPC is ducted through 4 pipes, to cool down the 1st stage of the LPT and the front cavity.Air from the 4th and 9th stage of the HPC goes through the HPTCC valve andis ducted through 2 pipes, to cool down the cavity that surrounds the HPTshroud.

Page: 22

Lufth

ansa

Tec

hnic

al T

rain

ing

For

Tra

inin

g P

urpo

ses

Onl

y

ENGINETURBINE SECTION

A318/A319 /A320 /A321CFM56-5B

72-50

Page: 23FRA US-T Kh August 03L3 Page: 23Figure 11 High Pressure Turbine

Lufth

ansa

Tec

hnic

al T

rain

ing

For

Tra

inin

g P

urpo

ses

Onl

y

ENGINEGENERAL

A318/A319/A320/A321CFM 56-5B DAC

72-00

Page: 24FRA US/T Kh August 03L2

ENGINE GENERAL CONCEPTThe main hardware components are located in the core section of the engine.Compressor Discharge Pressure (CDP) air is delivered to the combustion sec-tion which has a Double Annular Combustor.Controlled release of the combustion energy is used to drive the turbine rotors.The residual energy is converted into thrust.The combustion case is a welded structure that encloses the combustionchamber and the following High Pressure Turbine (HPT) components:- CDP seal- High pressure Turbine (HPT) nozzles- HPT shroud- Low Pressure Turbine (LPT) stage 1 nozzle assembly

Page: 24

Lufth

ansa

Tec

hnic

al T

rain

ing

For

Tra

inin

g P

urpo

ses

Onl

y

ENGINEGENERAL

A318/A319/A320/A321CFM 56-5B DAC

72-00

Page: 25FRA US/T Kh August 03L2 Page: 25Figure 12 CFM56-5B DAC Core Engine

Lufth

ansa

Tec

hnic

al T

rain

ing

For

Tra

inin

g P

urpo

ses

Onl

y

ENGINEGENERAL

A318/A319/A320/A321CFM 56-5B DAC

72-00

Page: 26FRA US/T KH August 03 L3

COMBUSTION SECTIONThe fuel/air mixture is ignited in the combustion section which consists of:- the combustion case- the combustion chamberThe combustion case provides the structural interface between the High Pres-sure Compressor section and the Low Pressure Turbine section.The front face of the combustor is attached to the rear of the compressor mod-ule.Its rear face is bolted onto the LPT module front flange. The rear section of thecombustor houses the HPT module.The -5B Double Annular Combustor contains an outer dome, known as pilot,and an inner dome, known as main.The case has 20 double-tip fuel nozzle mounting pads and accommodates thefollowing fuel supply manifolds: - Pilot - Main 1 - Main 2

Pilot domeThe pilot dome is the low power region of the combustor and is designed toachieve high combustion efficiency. lt minimizes the production of carbon mon-oxide and unburned hydrocarbons and acts as a pilot source of heat for themain dome.

Main domeThe main dome is a high velocity/high air flow region and is designed to reducethe reaction temperature and residence time of combustion to minimize sootand nitric oxide formation.

Page: 26

Lufth

ansa

Tec

hnic

al T

rain

ing

For

Tra

inin

g P

urpo

ses

Onl

y

ENGINEGENERAL

A318/A319/A320/A321CFM 56-5B DAC

72-00

Page: 27FRA US/T KH August 03 L3

LINER

CENTERBODY

SWIRL CUPS

COWL

Page: 27Figure 13 Combustion Chamber

Lufth

ansa

Tec

hnic

al T

rain

ing

For

Tra

inin

g P

urpo

ses

Onl

y

ENGINEGENERAL

A318/A319/A320/A321CFM56-5B DAC

72-00


TURBINE FRAME ASSEMBLY

GENERALThe turbine frame module is the major structural assembly at the rear of theengine. It supports the Low Pressure Turbine ( LPT ) rotor rear section andprovides for engine rear mounting on the airframe. The turbine frame includesthe following major parts: turbine frame No. 5 bearing support

Turbine FrameThe turbine frame assembly is a nickel-alloy weldment. It is made of a hub anda polygonal outer casing structurally connected by 16 airfoil-shaped struts,approximately tangent to the hub. The outer casing front outer flange bolts tothe LPT case rear flange. The outer casing rear outer flange supports the ehaust mixer.At the periphery of the outer casing, there are 3 clevis for aft engine mountingto the aircraft pylon and 2 brackets for handling/lifting purposes. The hub for-ward side provides support and attachment for: the No.5 bearing support ( hub inner front flange ) the oil collector of the aft oil sump assembly ( hub outer front flange )

On its aft side, the hub provides support and attachment for: the flange assembly supports for the flame arrestor the center body, or exhaust plug

From its front to aft sides, the hub provides passage for 8 tubular conduits.These conduits interconnect the rear rotating air/oil seals enclosure with thecenter body enclosure. They vent the oil vapors and drain the oil leakage ( pastthe seals ) into the center body enclosure from which they discharge outboard.

No. 5 Bearing SupportThe No.5 bearing support is made of steel alloy. Its outer flange mounts to theturbine frame hub inner front flange. The front bore of the No.5 bearing supportcontains the No.5 bearing outer race and a concentricity-adjusting sleeve. Therear face of the support carries the oil inlet cover which supports the oil supplytube for the No.4 and 5 bearings, and provides for pressurization of the dualair/oil seal installed at the rear end of the center vent tube.

Page: 28

Lufth

ansa

Tec

hnic

al T

rain

ing

For

Tra

inin

g P

urpo

ses

Onl

y

ENGINEGENERAL

A318/A319/A320/A321CFM56-5B DAC

72-00


View Aft Looking Forward

CLEVIS MOUNTS

HUB STRUCTURE

NO.5 BEARING SUPPORT

OUTER CASING

AFT SUMP ASSY

OIL LINES TO/FROM

AFT OIL SUMP

TURBINESTRUTS

Page: 29Figure 14 Turbine Frame Assembly

Lufth

ansa

Tec

hnic

al T

rain

ing

For

Tra

inin

g P

urpo

ses

Onl

y

ENGINEOIL SYSTEM

A319 / A320 / A321CFM 56-5A

79-00

Page: Page: 30FRA US/T KoA Nov 03 L2

ATA 79 OIL

79 - 00 GENERAL

OIL SYSTEM PRESENTATION

General Description a Supply Circuit a Scavenge Circuit a Vent Circuit

It Lubricates and cools the Bearings of the Forward and Aft Sumps.It also lubricates Bearings and Gears in the Transfer and Accessory GearBoxes.The Major Components of the Oil System are: The Oil Tank The Lubrication Unit The Servo Fuel Heater The Main Fuel Oil Heat Exchangers.

Indicating and Monitoring is provided by the Detectors and Sensors shown onthe Schematic.

Oil Supply CircuitThe Oil from the Tank passes through the Supply Pump and Supply Filter tolubricate the forward and aft Sumps, and also the Accessorys and Gearboxes.On the Oil Supply Line a Visual Filter Clogging Indicator, an Oil TemperaturSensor, an Oil Low Pressure Switch and an Oil Pressure Transmitter are pro-vided for Indication and Monitoring.Also an Oil Quantity Transmitter is provided on the Oil Tank.Note the Installation of the ECU Oil Temperatur Sensor for the Fuel ReturnValve.

Oil Scavenge CircuitThe Oil from Bearings, Transfer Gearbox and Accessory Gearbox returns tothe Tank by means of four Scavenge Pumps protected upstream by Strainersand Chip Detectors.

To keep Oil Temperatur within Limits, the Oil is cooled through the Servo FuelHeater and the Fuel/Oil Heat Exchanger.In Case of Scavenge Filter Clogging, an Oil Differential Pressure (Delta P)Switch signals it to the Cockpit and its Clogging Indicator shows it on the En-gine system page with a message on E/WD accompanied by a single chime

Oil Vent CircuitSome Air entrained in the Scavenge Oil is separated in the Tank by a Dearatorand is vented to the Forward Sump through the Transfer Gearbox and RadialDrive Shaft.The Sumps are vented Overboard through the Low Pressure Turbine Shaft toprevet Overpressure in the Sump.Air entrapped in the Scavenge Oil Pressurizes the Tank and provides adequateOil Pressure to the Supply Pump.

System Monitoring and LimitationsThe operation of the engine oil system may be monitored by the following flightdeck indications. engine oil pressure engine oil temperature

- MIN.PRIOR EXCEEDING IDLE : -100C- MAX CONTINIOUS: 1400C- MAX TRANSIENT: 1550C

oil tank contents 24 US quartsIn addition warnings may be given for the following non normal conditions: low oil pressure

- RED LINE LIMIT: 13 PSI high oil pressure

- ADVISORY: 90 PSI scavenge filter clogged.

Page: 30

Lufth

ansa

Tec

hnic

al T

rain

ing

For

Tra

inin

g P

urpo

ses

Onl

y

ENGINEOIL SYSTEM

A319 / A320 / A321CFM 56-5A

79-00

Page: Page: 31FRA US/T KoA Nov 03 L2

PUMP SUPPLY

PRESSURE OIL SCAVENGE OIL VENT PRESSURE

RDS HOUSING

SUPPLY FILTER

SUPPLYPUMP

SCAVENGEFILTER

BY PASS VALVE& CLOGGING IND.

CLOGGING SWITCH

OIL TEMP.SENSOR

ECU OIL TEMP.SENSOR

ANTI SIPHONDEVICE

OIL PRESSURE TRANSMITTER

LOW OIL PRESSURESWITCH

COLD STARTPRESSURE RELIEF VLV

MAIN

Page: 31Figure 15 OIL SYSTEM SCHEMATIC

Lufth

ansa

Tec

hnic

al T

rain

ing

For

Tra

inin

g P

urpo

ses

Onl

y

ENGINEOIL SYSTEM

A319 / A320 / A321CFM56-5A

79-00

Page: 32FRA US/T KoA Nov 03 L3

LUBRICATION UNIT

GeneralThe lubrication unit provides oil under the required pressure for lubrication and for scavenge of the oil after lubrication and circulation to the oil/fuel heatexchanger and oil tank. The lubrication unit its mounted on the AGB front face.DescriptionThe lubrication unit has a single housing containing the following items : Five positive displacement pumps( Gear Type, one oil supply and 4 scav-

enge pumps). Six filters (one oil supply filter, 4 chip detectors and scavenge pump filters). One relief valve (305 psi, on oil supply pump discharge side). Two clogging indicators (one for the oil supply filter and one for the main

scavenge filter). Two bypass valves (one for the oil supply filter and one for the main scav-

enge filter).

Anti siphon SystemThe supply lines from the oil tank to supply the pump has an antisiphon deviceto prevent the drainage of the lube tank into the gearboxes and sumps whenthe engine is shut down for extended periods.

Lube pump supply filter Downstream of the supply pump, the oil flows through the supply filter assembly. The filter has the following components. One filter (15 microns) One clogging indicator subjected to the upstream and downstream pres-

sures of the supply filter.The indicator has a red warning indicator and isrearmed manually (2 bars to 2.3 bars) (29 PSID to 33 PSID).

One bypass valve which opens if the supply filter clogs (2.50 bars to 2.70bars) (36 PSID to 39 PSID).

Two capped provisions for a pressure gage upstream of the filter,and a tem-perature sensor.

Scavenge filter The flows from the 4 scavenge pumps are mixed together at the scavengecommon filter inlet. This filter assembly consists of the following : One 25 micron filter One clogging indicator, similar to the one on the supply filter (2 bars to 2.3

bars) (29 PSID to 33 PSID). An upstream and a downstream provision for measurement of filter pres-

sure loss as a function of clogging.Filter clogging is indicated on the ECAMsystem.

One bypass valve which opens if the filter clogs.(2.5 bars to 2.7 bars) (36PSID to 39PSID)

Page: 32

Lufth

ansa

Tec

hnic

al T

rain

ing

For

Tra

inin

g P

urpo

ses

Onl

y

ENGINEOIL SYSTEM

A319 / A320 / A321CFM56-5A

79-00

Page: 33FRA US/T KoA Nov 03 L3

A

LUBRICATION UNIT

TEMPERATUR SENSORPOSITION

Page: 33Figure 16 Lubrication Unit

Lufth

ansa

Tec

hnic

al T

rain

ing

For

Tra

inin

g P

urpo

ses

Onl

y

ENGINEOIL SYSTEM

A318/A319/A320/A321CFM 56-5B

79-00


ATA 79 OIL SYSTEM79-00 GENERALOIL SYSTEM PRESENTATION

GENERALThe engine oil system includes : a supply circuit, a scavenge circuit, a vent circuit.

It lubricates and cools the bearings of the forward and aft sumps.It also lubricates bearings and gears in the transfer and accessory gear boxes.The major components of the oil system are : the oil tank, the lubrication unit,the servo fuel heater and the fuel/oil heat exchangers.Indicating and monitoring is provided by the detectors and sensors shown onthe schematic.

OIL SUPPLYThe oil from the tank flows through the supply pump and the main filter, orthrough the back up filter in case of main filter clogging.It then flows to the forward and aft sumps, and to the accessory and transfergearboxes.The pump delivery pressure is not controlled, but the oil output flow is, by de-sign, always in excess of the lubrication requirements.A pressure relief valve bypasses part of the output flow to protect the supplypump against abnormal output pressure build-up.If the main filter becomes clogged, a bypass valve opens and the oil flowsthrough the backup filter.A clogging switch sends a signal to the Engine Interface Unit ( EIU ) and aclogging indicator pops out on the filter housing.The anti-siphon device prevents oil from draining by gravity from the tankthrough the pump into the gearbox after engine shutdown. It uses air from theforward sump.

OIL SCAVENGEThe oil scavenge from the forward and aft sumps, and the transfer and acces-sory gearboxes is sucked by four scavenge pumps.Each pump is protected by a strainer and a electrical Master Magnetic ChipDetector.The scavenge oil then flows through a master chip detector and a scavengefilter screen ( inside the servo fuel heater ), then it is cooled through the servofuel heater and the fuel/oil heat exchanger before returning to the oil tank.

OIL VENTThe air mixed with the scavenge oil is separated in the tank by a deaerator andis vented to the forward sump through the transfer gearbox and radial driveshaft.The sumps are connected together by the center vent tube, which vents themto the outside air by the engine exhaust plug, through a flame arrestor.

System Monitoring and LimitationsThe operation of the engine oil system may be monitored by the following flightdeck indications. engine oil pressure engine oil temperature

- MIN.PRIOR EXCEEDING IDLE : -100C- MAX CONTINIOUS: 1400C- MAX TRANSIENT: 1550C

oil tank contents 24 US quartsIn addition warnings may be given for the following non normal conditions: low oil pressure

- RED LINE LIMIT: 13 PSI high oil pressure

- ADVISORY: 90 PSI filter clogged.

Page: 34

Lufth

ansa

Tec

hnic

al T

rain

ing

For

Tra

inin

g P

urpo

ses

Onl

y

ENGINEOIL SYSTEM

A318/A319/A320/A321CFM 56-5B

79-00


ANTI

SIPHON

BYPASS

VALVE

MAIN

FILTER

OIL TEMP

SENSOR

SUPPLY

PUMPSCAVENGE

SCREENS

BACK UP FILTER

MAG PLUG

SCREEN

FUEL/OILHEATEXCH

SERVOFUELHEATER

ELEC CHIP DETECTOR

SCAVENGEPUMPS

INDIVDUALCHIP

DETECTORS

OIL PRESS

X-MITTER

LOW OIL

PRESS SW

FWD

SUMP

AFT

SUMPEIU

TGB

A

G

B

OIL TANK

LUBRICATION UNIT

OIL OTY

SENSOR

CLOGGING

SWITCH

RE

FE

R T

O A

3 PA

GE

E

VISUALPOPOUTINDICATOR

SDAC

TOIL

SENSOR

(OPTION)

ECU

Page: 35Figure 17 Oil System

Lufth

ansa

Tec

hnic

al T

rain

ing

For

Tra

inin

g P

urpo

ses

Onl

y

ENGINEOIL SYSTEM

A318/A319 / A320 / A321CFM 56-5B

79-00

Page: Page: 36FRA US/T Kh August 03 L3

LUBRICATION UNIT

The lubrication unit has two purposes- it pressurizes and filters the supply oil for lubrication of the engine bear-ings and gears- it pumps in scavenge oil to return it to the tank.

lt is installed on the left hand side of the AGB front face.Externally, the lubrication unit has :

- a suction port (from the oil tank).- four scavenge ports (aft & fwd sumps, TGB, AGB).- four scavenge screen plugs.- an oil out port (to master chip detector).- a main oil supply filter.- a back-up filter.- pads for the oil temperature sensor and the oil differential pressure switch.

Internally, it has 5 pumps driven by the AGB, through a single shaft. The lubeunit is lubricated with supply pump outlet oil, which flows within the drive shaft.The AGB mounting pad has no carbon seal and the lube unit has an 0-ring forsealing purposes.NOTE: Individual Chip Detectors are used for Trouble Shooting !

MAIN SUPPLY AND BACKUP FILTER

Description and OperationIn the supply circuit, downstream from the pressure pump, oil flows through thesupply system which includes, first, the main oil supply filter.A sensor, installed in between the upstream and downstream pressures of thesupply filter, senses any rise in differential pressure due to filter clogging.lf the filter clogs, an electrical signal is sent to the aircraft systems for cockpitindication.A by-pass valve, installed in parallel with the filter, opens when the differentialpressure across the valve is greater than the spring load.The oil then flows through the back-up filter and goes to the pump outlet.The back-up filter is a metallic, washable filter.

During normal operation, the oil flow, tapped at the main supply filter outlet,washes the back-up filter and goes back to the supply pump inlet, through arestrictor.The main filter is discardable and secured on the lube unit cover by a drainplug.To prevent the filter element from rotating when torquing the drain plug, a pininstalled on the filter element engages between two ribs cast in the lube unitcover.

DescriptionDownstream of the supply pump, the oil flows through the supply filter assem-bly.This filter assembly comprises: one filter one clogging indicator transmitter subjected to the upstream and down

stream pressures of the supply filter one bypass valve which opens if the supply filter clogs one back up filter operating if the supply filter clogs two capped provisions for a pressure gage upstream of the filter, and a tem-

perature sensor.Downstream of supply filter, the oil flows through three outlets to the forwardsump, aft sump and the AGB / TGB.

Page: 36

Lufth

ansa

Tec

hnic

al T

rain

ing

For

Tra

inin

g P

urpo

ses

Onl

y

ENGINEOIL SYSTEM

A318/A319 / A320 / A321CFM 56-5B

79-00


O-RING

BACK-UP

SEAL

BACK-UP FILTER ELEMENT

BACK-UP FILTER

O-RING

O-RINGBACK-UP

SEAL

LUBE FILTER

ELEMENT

COVER

DRAIN

PLUG

MAIN OIL SUPPLY FILTER

SCAVENGE SCREEN

ASSY

OPTIONAL

MAGNETIC BAR

SCAVENGE SCREEN

PLUG ASSY

Page: 37Figure 18 Lubrication Unit Interface

Lufth

ansa

Tec

hnic

al T

rain

ing

For

Tra

inin

g P

urpo

ses

Onl

y

ENGINEOIL SYSTEM

A319 / A320 / A321CFM 56-5B

79-00

Page: Page: 38FRA US/T Kh August 03 L2/3

MASTER CHIP DETECTORThe Master Chip Detector (MCD) collects magnetic particles suspended in theoil that flows from the common outlet of the four scavenge pumps, by means oftwo magnets on a probe immersed in the oil flow.lt is installed on the lubrication unit and is connected to an oil contaminationpop-out indicator, through the DPM wiring harness.The probe is locked in position through a bayonet system.When a sufficient amount of particles are caught, the gap between the 2 mag-nets is bridged and the resistance between them drops. This electrical signal isthen sent to the contamination pop-out indicator.The MCD assembly consists of :

- a housing which has two flanges for attachment.- a check valve, built in the housing, that prevents oil spillage when the probe is removed and also provides a passage for the oil flow, in case of chip detector disengagement.- a hand removable probe, which has a back-up seal, an 0-ring seal, and two magnets.- a two-wire, shielded electrical cable and an interface connector.

Page: 38

Lufth

ansa

Tec

hnic

al T

rain

ing

For

Tra

inin

g P

urpo

ses

Onl

y

ENGINEOIL SYSTEM

A319 / A320 / A321CFM 56-5B

79-00

Page: Page: 39FRA US/T Kh August 03 L2/3 Page: 39Figure 19 Master Chip Detector

Lufth

ansa

Tec

hnic

al T

rain

ing

For

Tra

inin

g P

urpo

ses

Onl

y

ENGINEOIL SYSTEM

A318/A319 / A320 / A321CFM 56-5B

79-00


MAGNETIC CONTAMINATION INDICATORThe magnetic contamination indicator works in conjunction with the MCD andits purpose is to provide maintenance personnel with a visual indication of oilcircuit contamination.The indicator is an electro-mechanical device, located on the right hand side ofthe downstream fan case, just above the oil tank.When magnetic contamination in the oil occurs, an electronic circuit in the indi-cator detects a drop in resistance between the two magnets on the MCDprobe.The electronic circuit then energizes a solenoid which triggers a red pop-outbutton, thus providing a visual indication.After maintenance action, the pop-out button must be manually reset.lt has 2 electrical connectors

- One for the wiring harness connected to the MCD- One for the harness connecting the indicator to the EIU. .

Page: 40

Lufth

ansa

Tec

hnic

al T

rain

ing

For

Tra

inin

g P

urpo

ses

Onl

y

ENGINEOIL SYSTEM

A318/A319 / A320 / A321CFM 56-5B

79-00

Page: Page: 41FRA US/T Kh August 03 L2 Page: 41Figure 20 Oil Contamination Pop-Out Indicator

Lufth

ansa

Tec

hnic

al T

rain

ing

For

Tra

inin

g P

urpo

ses

Onl

y

ENGINEOIL SYSTEM

A318/A319 / A320 / A321CFM 56-5B

79-30

Page:Page: 42FRA US/T Kh August 03 L3

OIL FILTER DIFFERENTIAL PRESSURE SWITCHThe oil differential pressure switch is located on a bracket on the engine abovethe scavenge filter. Lines are routed to the switch from bosses on the scavengefilter.Actuation of the differential pressure switch is at :25.5 plus or minus 1 PSID increasing pressure 22 PSID decreasing pressure.

OIL TEMPERATURE SENSORThe oil temperature sensor is located on the oil pressure filter downstream ofthe pressure pump.The oil temperature is sensed by a dual resistor unit.

OIL TEMPSENSOR

OIL FILTERDIFFPRESSURESWITCH

Page: 42

Lufth

ansa

Tec

hnic

al T

rain

ing

For

Tra

inin

g P

urpo

ses

Onl

y

ENGINEOIL SYSTEM

A318/A319 / A320 / A321CFM 56-5B

79-30

Page:Page: 43FRA US/T Kh August 03 L3

0.8

1.2

0.8

1.2

EIU

SDAC1

OIL FILTER

CLOG CLOG

Page: 43Figure 21 Oil Temperature and Diff. Press. Indication

Lufth

ansa

Tec

hnic

al T

rain

ing

For

Tra

inin

g P

urpo

ses

Onl

y

Engine Fuel and ControlDistribution

A319 / A320 / A321CFM 56-5B DAC

73-10

Page: Page: 44FRA US/T bu July 02 L3

ATA 73 FUEL SYSTEMFUEL DISTRIBUTION

Fuel from the aircraft tank enters the engine fuel pump through a fuel supplyline.The pressurized fuel then goes to the main oil/fuel heat exchanger to cooldown the engine supply oil. It goes back to the fuel pump, where it is filtered,pressurized again and split into two fuel flows.The main fuel flow goes to the Hydromechanical Unit (HMU) metering systembefore passing through a fuel flow transmitter.Before going to the fuel nozzles, the fuel is directed to a Fuel Manifold Modulat-ing Valve (FMMV). In some manuals it is also sometimes called a BurnerSelection Valve (BSV). The FMMV splits the metered fuel into three fuel flows,and delivers them to the nozzles through three different supply manifolds.The other fuel flow goes to a servo fuel heater, to prevent any ice particles en-tering sensitive servo systems. The heated fuel enters the HMU servo mecha-nism area and is then directed to different fuel actuated components.A bypass line returns unused fuel from the HMU to the inlet of the main oil/fuelheat exchanger.An Integrated Drive Generator (IDG) oil cooler and a Fuel Return Valve (FRV)are also an this line. The FRV may re-direct some returning fuel back to theaircraft tank. Before returning to the aircraft tank, the hot fuel is mixed with coldfuel from the outlet of the first stage of the fuel pump.

Page: 44

Lufth

ansa

Tec

hnic

al T

rain

ing

For

Tra

inin

g P

urpo

ses

Onl

y


A319 / A320 / A321CFM 56-5B DAC

73-10

Page: Page: 45FRA US/T bu July 02 L3 Page: 45Figure 22 DAC FUEL DISTRIBUTION

Lufth

ansa

Tec

hnic

al T

rain

ing

For

Tra

inin

g P

urpo

ses

Onl

y


A318/A319 / A320 / A321CFM 56-5B DAC

73-10


FUEL DISTRIBUTION

Fuel Nozzle SupplyThe FMMV delivers fuel to the nozzles through three different supply man-ifolds: -The Pilot manifold supplies the 20 outer domes of the combustion cham -

ber (pilot burner). -The Main 1 manifold supplies 10 of the 20 inner domes of the combustion

chamber (main burner 1). -The Main 2 manifold supplies the remaining 10 inner domes (main burner

2).The fuel flow sent to the Pilot and the Main 1 manifolds is controlled by theDouble Annular Modulating Valve (DAMV).

The fuel flow sent to the Main 2 manifold is controlled by the Main Burner Stag-ing Valve (MBSV).The fuel is delivered to 20 double-tip fuel nozzles. Each nozzle has two inlets,one connected to the Pilot manifold, and the other connected to either the Main1 or the Main 2 manifold.Each manifold is made up of two halves. The halves are connected to the threeFMMV outlets through a connection box. The six halves are secured to thecore engine.Tubes with connecting nuts are brazed onto the manifolds, to supply fuel to thecorresponding nozzles

Page: 46

Lufth

ansa

Tec

hnic

al T

rain

ing

For

Tra

inin

g P

urpo

ses

Onl

y


A318/A319 / A320 / A321CFM 56-5B DAC

73-10


LINER

CENTERBODY

SWIRL CUPS

COWL

Page: 47Figure 23 DAC NOZZLE SUPPLY

Lufth

ansa

Tec

hnic

al T

rain

ing

For

Tra

inin

g P

urpo

ses

Onl

y


A319 / A320 / A321CFM 56-5B DAC

73-10


FUEL MANIFOLD MODULATING VALVE (FMMV)The FMMV is installed at 6 oclock an the core engine. To access the valve,both thrust reverser C ducts must be opened.The FFMV is controlled by the Electronic Control Unit (ECU) and is fuel oper-ated by the Hydro Mechanical Unit (HMU).lt consists of a Double Annular Modulating Valve (DAMV), a Main Burner Stag-ing Valve (MBSV), two override valves and a manifold cooling orifice.The DAMV controls the fuel split between the Pilot manifold and Main 1 man-ifold. Its linear motion is commanded by two Electro Hydraulic Servo Valves(ESHV). Its position is controlled by the ECU through feedback from a LinearVariable Differential Transducer (LVDT).The MBSV acts as an on/off valve controlling the fuel circuit to the Main 2 man-ifold. lt has dual redundant signal switches.The override valves discharge excessive pressure in the Pilot circuit into theMain 1 and Main 2 manifolds.The manifold cooling orifice allows passage of cooling fuel into the Main 1 andMain 2 manifolds and nozzles during staged operation.

Page: 48

Lufth

ansa

Tec

hnic

al T

rain

ing

For

Tra

inin

g P

urpo

ses

Onl

y


A319 / A320 / A321CFM 56-5B DAC

73-10

Page: Page: 49FRA US/T bu July 02 L3 Page: 49Figure 24 DAC FUEL MANIFOLD MODULATING VALVE

Lufth

ansa

Tec

hnic

al T

rain

ing

For

Tra

inin

g P

urpo

ses

Onl

y


A319 / A320 / A321CFM 56-5B DAC

73-10


FUEL MANIFOLD MODULATION VALVE MODES (FMMV)

In order to operate the Double Annular Combustor (DAC) at all flight conditions,the DAC control turns the fuel manifolds ON, or OFF, to maintain combustorstability, while adhering to operability, turbine temperature limitations and pol-luant emission limits.Three modes can be set by the DAC control: - Mode 20/0: 20 Pilot burners are fuel supplied. Main 1 and Main 2 are

OFF. - Mode 20/10: 20 Pilot burners and 10 Main 1 burners are fuel supplied.

The 10 Main 2 burners are OFF. - Mode 20/20: 20 Pilot burners, 10 Main 1 burners and 10 Main 2 burners

are supplied.

Page: 50

Lufth

ansa

Tec

hnic

al T

rain

ing

For

Tra

inin

g P

urpo

ses

Onl

y


A319 / A320 / A321CFM 56-5B DAC

73-10

Page: Page: 51FRA US/T bu July 02 L3 Page: 51Figure 25 DAC FMMV OPERATING MODES

Lufth

ansa

Tec

hnic

al T

rain

ing

For

Tra

inin

g P

urpo

ses

Onl

y

Engine Fuel and ControlControlling General

A319 / A320 / A321CFM 56-5B DAC

73-20


FUEL FLOW SPLITS

Fuel flow splits between the Pilot and Main 1 manifolds are hydraulically con-trolled by the fuel nozzle Dp characteristics in 20/10 mode which are, approxi-mately, 52% Pilot and 48% Main 1 total flow.In 20/20 mode, the DAMV is commanded to allow fuel flow to all three fuelmanifolds (Pilot, Main 1, Main 2). During transient operation, fuel flow splitsbetween the Pilot and Main manifolds are hydraulically controlled by the fuelnozzle Dp characteristics which are, approximately, 40% Pilot and 60% Main(Main 1 + Main 2).During steady state operation, fuel flow splits between the Pilot and Main man-ifolds are controlled to a lower Pilot percentage.

ENGINE FUEL AND CONTROL - DISTRIBUTIONThe DAC engine reduces exhaust emissions.The DAC system increases the size of the flame, while the fuel flow is almostthe same as that of the Single Annular Combustor (SAC) engine. This reducesthe flame core temperature which also reduces the amount of nitric-oxide(NOx) gasses that the engine exhausts. Of the 20 double tip fuel nozzles, tenare cooling fuel nozzles that supply fuel to the Pilot and Main 1 zones of thecombustor.The other ten nozzles are the Bleed nozzles that supply fuel to the Pilot andMain 2 zones of the combustor. The nozzles are in pairs of two Cooling nozzlesand Bleed nozzles around the combustor outer case.

The fuel nozzles have three thrust modes: - 20/0 (low power) mode:

20 Pilot fuel nozzle tips of the cooling and bleed nozzles supply fuel. -20/10 (medium power) mode:

20 Pilot fuel nozzle tips of the cooling and bleed nozzles supply fuel. The ten Main 1 zone tips of the cooling nozzles also supply fuel. -20/20 (high power) mode:

20 Pilot fuel nozzle tips of the cooling and bleed nozzles supply fuel. The ten Main 2 zone tips of the bleed nozzles also supply fuel. The ten Main 1 zone tips of the cooling nozzles also supply fuel. All 40 tips are in operation in the 0/ 20 mode.

FUEL MANIFOLD MODULATION VALVE MODES (FMMV)In order to operate the Double Annular Combustor (DAC) at all flight conditions,the DAC control turns the fuel manifolds ON, or OFF, to maintain combustorstability, while adhering to operability, turbine temperature limitations and pol-luant emission limits.Three modes can be set by the DAC control: - Mode 20/0: 20 Pilot burners are fuel supplied. Main 1 and Main 2 are

OFF. - Mode 20/10: 20 Pilot burners and 10 Main 1 burners are fuel supplied.

The 10 Main 2 burners are OFF. - Mode 20/20: 20 Pilot burners, 10 Main 1 burners and 10 Main 2 burners

are supplied.

Page: 52

Lufth

ansa

Tec

hnic

al T

rain

ing

For

Tra

inin

g P

urpo

ses

Onl

y


A319 / A320 / A321CFM 56-5B DAC

73-20

Page: Page: 53FRA US/T bu July 02 L3 Page: 53Figure 26 FUEL FLOW SPLIT

Lufth

ansa

Tec

hnic

al T

rain

ing

For

Tra

inin

g P

urpo

ses

Onl

y


A319 / A320 / A321CFM 56-5B DAC

73-10


FUEL NOZZLESThe fuel nozzles are welded assemblies that deliver carefully calibrated andprecisely patterned fuel spray for combustion. Each nozzle is essentially twonozzles in one.They ensure a good light-off capability, efficient burning at high power enginesettings and a reduction of polluant emission at low power engine settings.They are installed on the combustion case assembly and connected to the fuelmanifold assembly.There are 20 double-tip fuel nozzles which deliver fuel into a double-headedcombustion chamber through a Pilot burner and a Main burner.The double-tip fuel nozzles deliver fuel through 4 independent flows.The nozzles consist of: - 2 independent fuel inlet connectors - 2 independent metering valve assemblies - A support to secure the fuel nozzle an the combustion case - 2 metering sets to calibrate Primary and Secondary fuel flow sprays for

Pilot and MainTwo types of nozzle are installed an the engine: - 10 Cooling nozzles - 10 Bleed nozzles

Though there are differences between the cooling and bleed nozzles in the in-ternal cooling circuits, they are more similar in function and fuel management.An additional Bleed Valve body is manufactured onto the Bleed Nozzle body.To facilitate nozzle type identification, a color band is wrapped around the valveassembly: - A blue band at the Cooling nozzle - A yellow band at the Bleed nozzle

FUEL NOZZLE

Page: 54

Lufth

ansa

Tec

hnic

al T

rain

ing

For

Tra

inin

g P

urpo

ses

Onl

y


A319 / A320 / A321CFM 56-5B DAC

73-10

Page: Page: 55FRA US/T bu July 02 L3 Page: 55Figure 27 DAC FUEL NOZZLES

Lufth

ansa

Tec

hnic

al T

rain

ing

For

Tra

inin

g P

urpo

ses

Onl

y


A319 / A320 / A321CFM 56-5B DAC

73-10


DISTRIBUTION 20/20 MODEIn the 20/20 mode, the Pilot zone tips, Main 1 zone tips and Main 2 zone tipssupply fuel to the combustor.In this mode, all tips in both the cooling and bleed nozzles operate.In this condition, the engine operates in the high power range.Fuel goes through the open Pilot and Main 1 manifold ports of the Double An-nular Modulating Valve (DAMV). The DAMV decreases the amount of fuel thatgoes to the Pilots to distribute the fuel better in all three manifolds.The Main 2 staging valve (MSV) is also open and metered fuel from this openvalve goes to the Main 2 manifold.The Main 2 manifold supplies fuel to the Main 2 zone tips in the bleed nozzles.

COOLING NOZZLE - PILOTThe fuel nozzle assembly receives fuel through the inlets.The fuel passesthrough inlet strainers and flows into the portion of the support that houses thevalves.Primary Pilot flow travels through a Manifold Leakage Control Valve (MLCV)and follows a cooling circuit down to the Main tip.The Primary Pilot flow keeps the Main tip cool during staged operation.From this point, the fuel is routed back up to cool the Pilot tip before exiting intothe combustion chamber.Secondary Pilot flow is scheduled by a Pressure Flow Divider Valve Cartridge.Fuel is routed by this valve through a trimming restrictor orifice, then ported toa Secondary Metering Set and sprayed into the combustor. The Secondaryspray profile is wider than the Primary spray profile, thus it surrounds the Pri-mary spray.

COOLING NOZZLE - MAIN 1

Primary Main flow travels through the MLCV and is routed into the Main Pri-mary Metering Set to be sprayed into the combustion chamber.Secondary Main flow is scheduled by a Pressure Flow Divider Valve cartridge.This valve routes fuel through a trimming restrictor orifice and fuel is thenrouted to the Main Secondary Metering Set.

COOLING NOZZLE - COOLING CIRCUITSDuring staged operation, a restrictor routes fuel from the Pilot inlet to a pas-sage around the Main flow valves to cool them down (main circuit shut off).Then, the fuel exits into the Main manifold.Another purpose of this restrictor is to provide flow to the Main manifold andBleed nozzles to cool them down during staged operation.

Page: 56

Lufth

ansa

Tec

hnic

al T

rain

ing

For

Tra

inin

g P

urpo

ses

Onl

y


A319 / A320 / A321CFM 56-5B DAC

73-10

Page: Page: 57FRA US/T bu July 02 L3 Page: 57Figure 28 DAC PILOT + MAIN1 + MAIN2 OPERATION

Lufth

ansa

Tec

hnic

al T

rain

ing

For

Tra

inin

g P

urpo

ses

Onl

y


A319 / A320 / A321CFM 56-5B DAC

73-10


DISTRIBUTION 20/10 MODEIn the 20/10 mode, the Pilot zone tips and the Main 1 zone tips supply fuel tothe combustor.In this condition the engine operates in the medium power range.Fuel goes through the open Pilot and Main 1 ports of the DMV. The MSV is inthe closed position and no metered fuel goes to the Main 2 manifold.Fuel for the Main 2 tip cooling circuit comes from the metered fuel supply forthe Pilot and Main 1 manifolds.The metered fuel goes through cooling orifices inside the FMMV.The cooling orifices enable sufficient flow to move fuel through the Main 2 tipcooling circuit, at a pressure low enough not to open the primary or secondaryvalves of the Main 2 tips. Fuel from the cooling circuit goes from the Main 2 tipbleed valves and mixes with the secondary orifice circuit of the Pilot zone tips.There is a constant positive flow through the cooling circuit for the Main 2 man-ifolds and tips.

BLEED NOZZLE - PILOT

Primary Pilot flow travels through the MLCV and follows a cooling circuit downto the Main tip. The primary Pilot flow keeps the Main tip cool during stagedoperation.Secondary Pilot flow is scheduled by a Pressure Flow Divider Valve cartridge.This valve routes fuel through a trimming restrictor orifice and fuel is thenrouted to the Pilot Secondary Metering Set.From this point, the fuel is routed back up to cool the Pilot tip before exiting intothe combustion chamber.

BLEED NOZZLE - MAIN 2

Primary Main Flow travels through the MLCV and is routed into the Main Pri-mary Metering Set to be sprayed into the combustion chamber.Secondary main flow is scheduled by a Pressure Flow Divider Valve Cartridge.This valve routes fuel through a trimming restrictor orifice and fuel is thenrouted to the Main Secondary metering Set.

BLEED NOZZLE - COOLINGThe bleed valve provides controlled cooling flow and helps to balance thenozzle/manifold system.

This is accomplished by an additional flow divider valve (Bleed Valve) whichschedules fuel from the Main Bleed fuel nozzle inlet to the secondary Pilot flowcircuit.The cooling fuel comes from the cooling nozzles. lt is routed through an orificelocated in the FMMV. The orifice allows passage of the cooling fuel from theMain 1 to the Main 2 manifolds.

Page: 58

Lufth

ansa

Tec

hnic

al T

rain

ing

For

Tra

inin

g P

urpo

ses

Onl

y


A319 / A320 / A321CFM 56-5B DAC

73-10

Page: Page: 59FRA US/T bu July 02 L3 Page: 59Figure 29 DAC PILOT + MAIN 1 OPERATION

Lufth

ansa

Tec

hnic

al T

rain

ing

For

Tra

inin

g P

urpo

ses

Onl

y


A319 / A320 / A321CFM 56-5B DAC

73-10


DISTRIBUTION 20/0 MODEOnly the pilot tips of both the cooling and bleed nozzles supply fuel to the com-bustor in this mode.In this condition, the engine operates at low power.Fuel goes through the open Pilot manifold port of the DMV in the FMMV. ThePilot manifold supplies metered fuel to the Pilot tips in the cooling and bleednozzles.The Main 1 manifold port of the DMV and the Main 2 staging valve (MSV),which controls fuel flow to the Main manifolds, are closed. The Main 1 and 2zone tips do not supply fuel to the combustor.Cooling fuel flows through the Main 1 and 2 manifolds and the Main 1 and 2tips of the cooling and bleed nozzles. This cooling fuel prevents coking at theMain 1 and 2 tips. It also prevents a delay in engine acceleration when one, orboth of the Main nozzles begin to flow fuel.Cooling fuel for the Main 1 manifold and nozzle tips comes from a tapping inthe Pilot manifold. The fuel goes through the cooling circuit of the Main 1 zonetips and goes back into the FMMV.In the FMMV, metered fuel supply from the Pilot manifold goes through one oftwo cooling orifices. Fuel from the Main 1 manifold goes through the other cool-ing orifice. Fuel from both of these cooling orifices goes to the Main 2 manifold.Fuel that passes through the Main 2 tip cooling circuit goes through a bleedvalve to mix with the secondary circuit of the Pilot zone tips of the Bleednozzles.

Page: 60

Lufth

ansa

Tec

hnic

al T

rain

ing

For

Tra

inin

g P

urpo

ses

Onl

y


A319 / A320 / A321CFM 56-5B DAC

73-10

Page: Page: 61FRA US/T bu July 02 L3 Page: 61Figure 30 DAC PILOT OPERATIONAL MODE

Lufth

ansa

Tec

hnic

al T

rain

ing

For

Tra

inin

g P

urpo

ses

Onl

y


A319 / A320 / A321CFM 56-5B

73-20


FRONT PANEL ELECTRICAL CONNECTORS

There are 15 threaded electrical connectors located on the front panel, identi-fied through numbers 11 to J15 marked on the panel.Each connector features a unique key pattern which only accepts the correctcorresponding cable plug.All engine input and command output signals are routed to and from channelsA and B, through separate cables and connectors.

ENGINE RATING / IDENTIFICATION PLUGThe engine rating/identification plug provides the ECU with engine configura-tion information for proper engine operation.lt is plugged into connector J 14 and attached to the fan case by a metal strap.It remains with the engine even after ECU replacement.The plug includes a coding circuit, equipped with pushpull links which eitherensure, or prohibit connections between different plug connector pins.The push-pull links consist of switch mechanisms located between 2 contactsand can be manually opened, or closed, according to customer requests.They include:

- 5B and 5B/P differentiation- engine type (SAC or DAC)- an optional PMUX engine condition monitoring kit- optional full EGT monitoring- tool, which enables the engine serial number to be loaded into the ECUs Non-Volatile Memory (NVM)- N1 trim level, to correct thrust differences between engines operating at the same N1 speedThe ECU stores schedules in its NVM, for all available engine configura-tions. During initialization, it reads the plug and selects a specific schedule.In the case of a missing, or invalid ID plug, the ECU uses the value storedin the NVM for the previous plug configuration.

Page: 62

Lufth

ansa

Tec

hnic

al T

rain

ing

For

Tra

inin

g P

urpo

ses

Onl

y


A319 / A320 / A321CFM 56-5B

73-20


SAFTY WIRE

O-RING

SHEATEDCABLE

Page: 63Figure 31 FADEC Electrical Connectors

Lufth

ansa

Tec

hnic

al T

rain

ing

For

Tra

inin

g P

urpo

ses

Onl

y

Engine IndicatingAnalyzers

A319 / A320 / A321CFM 56-5B DAC

77-30

Page: Page: 64FRA US/T Kh August 03 L2/3

MAINTENANCE TASKSAn LRU screen is available through the MCDU and is specific for DAC engines.This screen enables the operator to know the engine type and also a lot ofother useful information:

- Bump level- N1 trim level- Engine configuration- PMUX option- DAC configuration- SOV status*- Engine serial number*

* The last two items can be accessed and modified through the keypad of theMCDU to be in accordance with the aircraft configuration.The other parameters can only be read and are automatically updated when anew software version is programmed into the ECU, and/or, when the engineidentification plug is installed an the ECU.

Page: 64

Lufth

ansa

Tec

hnic

al T

rain

ing

For

Tra

inin

g P

urpo

ses

Onl

y

Engine IndicatingAnalyzers

A319 / A320 / A321CFM 56-5B DAC

77-30

Page: Page: 65FRA US/T Kh August 03 L2/3 Page: 65Figure 32 DAC LRU IDENT &SERIAL NR. ENTRY

Lufth

ansa

Tec

hnic

al T

rain

ing

For

Tra

inin

g P

urpo

ses

Onl

y


A319 / A320 / A321CFM 56-5B DAC

73-20


73-20 CONTROLLING(DAC )FADEC

SYSTEM OVERVIEWThe 3 different operating modes 20/0, 20/10, 20/20 are primarily commandedby the CFM56-5B DAC FADEC control system.The operating modes are controlled by the Electronic Control Unit (ECU),which sends electrical signals to position the DAMV and MBSV.The system switches from the 20/0 to 20/10 and to the 20/20 mode, dependingan two parameters:

- The corrected N 1 speed (N 1 K12)- The fuel/air ratio (FAR)

20/0 is used at low power conditions (30% thrust SLS)20/10 mode is an intermediate mode that can be thought of as a transition stepbetween 20/0 and 20/20 modes

Page: 66

Lufth

ansa

Tec

hnic

al T

rain

ing

For

Tra

inin

g P

urpo

ses

Onl

y


A319 / A320 / A321CFM 56-5B DAC

73-20

Page: Page: 67FRA US/T Kh August 03 L3 Page: 67Figure 33 ECU/DAC PARTICULARS

Lufth

ansa

Tec

hnic

al T

rain

ing

For

Tra

inin

g P

urpo

ses

Onl

y


A319 / A320 / A321CFM 56-5B DAC

73-20

Page: Page: 68FRA US/T Kh August 03L3

NORMAL OPERATION

The DAC 2 engine switch point from 20/0 to 20/10 occurs at 2630rpm N1 K12.The ECU will command the DAC to switch from 20/0 to 20/10 modes if bothcorrected N1 and combustor FAR are above their respective switch thresholds.This switch point is always active an ground and may also be active in flight, ifthe FAR is above a certain limit. The corrected N1 switch threshold is only ac-tive an ground.

APPROACHDuring approach operation, the ECU commands the DAC to switch from 20/0to 20/10 mode when:

- Approach idle is set- Flight condition- Symmetric ECS bleed is ON- FAR > specific threshold

The system switches back to 20/0 mode if one of the above conditions are nolonger met, or if a hung decel* is detected.* A hung decel is detected if the N2 decel rate is slower than 1/4 of the normaldecel schedule and when the engine is controlied an the FAR decel schedule.

NAC COOLINGIf the NAC system fails to move to the cooling position,or if both feedback sen-sors fail,the ECU forces the DAC to switch from 20/0 mode when EGT exceeds550 C on the ground.If EGT exceeds 700 C,the LPTCC valve will be commanded fully open.Once DAC mode has switched from 20/0 to 20/10 based on EGT,it cannotswitch back to 20/0 mode unless any of the following conditions are met:

One hour elapses20/20 mode is selectedAny engine transient (accel or decel) is madeEngine is operating on FAR decel control (to avoid decel hang-up)

Page: 68

Lufth

ansa

Tec

hnic

al T

rain

ing

For

Tra

inin

g P

urpo

ses

Onl

y


A319 / A320 / A321CFM 56-5B DAC

73-20

Page: Page: 69FRA US/T Kh August 03L3

NORMAL OPERATION APPROACH

NAC COOLING

Page: 69Figure 34 DAC FUEL FLOW SPLITS

Lufth

ansa

Tec

hnic

al T

rain

ing

For

Tra

inin

g P

urpo

ses

Onl

y


A319 / A320 / A321CFM 56-5B DAC

73-20


SIMPLEX HPTCC AND LPTCC

To support the added functionality of the FMMV, the HPTC and LPTC controlloops have been converted to ECU single channel control (simplex).Channel A of the ECU is dedicated to controlling the HPTC valve and channelB is dedicated to controlling the LPTCC valve.Loss of either channel will result in the respective valve moving to the fail-safeposition. However, the demand and selected positions of the valves are sentover ARINC to both channels.

Page: 70

Lufth

ansa

Tec

hnic

al T

rain

ing

For

Tra

inin

g P

urpo

ses

Onl

y


A319 / A320 / A321CFM 56-5B DAC

73-20


HPTC VALVE

LPTC VALVE

Page: 71Figure 35 Simplex HPTCC and LPTCC

Lufth

ansa

Tec

hnic

al T

rain

ing

For

Tra

inin

g P

urpo

ses

Onl

y

ENGINEAIR

A319/A320/A321CFM56-5B

75-30

Page: 72FRA US-T Kh Augustn 03 L2/3

ATA 75 ENGINE AIR75-00 GENERAL

GeneralThe engine air system is divided into three main functions: cooling compressor control indicating

The air system covers primary, secondary (bypass) and parasitic (cooling andpressurizing) airflows and the systems used to control airflow. It is composed of2 major sections: engine and nacelle

Engine section The airstream flowing through the CFM56 turbofan engine supplies 2 majorsystems: the internal the external air systems

The internal air system consists of the following sub-systems: propulsion airflow (secondary and primary flows) forward and aft bearing sump pressurizing air cooling air (HPTACC, RACSB) internal thrust balancing air

The external air system consists of the following sub-systems: fuel control system air (CDP) Low Pressure Turbine Active Clearance Control (LPTACC) high-energy igniter harness cooling air engine customer bleed air ECU cooling

Nacelle sectionThe nacelle installation is designed to provide cooling and ventilation air for en-gine accessories installed on the fan and core casing. The distribution and cir-culation of the air in the compartments is such that the temperature limit forspecific components is not exceeded. These limits ensure long life and providegood fire safety margins.

Compressor ControlThe Variable Stator Vane (VSV) and the Variable Bleed Valve (VBV) control thecompressor through the ECU.

IndicatingThe nacelle temperature only is indicated in the cockpit.

Page: 72

Lufth

ansa

Tec

hnic

al T

rain

ing

For

Tra

inin

g P

urpo

ses

Onl

y

ENGINEAIR

A319/A320/A321CFM56-5B

75-30

Page: 73FRA US-T Kh Augustn 03 L2/3

CORE COMPARTMENT

CUSTOMER BLEED

5TH STAGE

CUSTOMER BLEED

9TH STAGE

FAN AIR ( FROM SCOOP )

LPT COOLING

STG 9

STG 4

STG 5

STG 9

SB

FAN AIR FLOW

PRIMARY AIR FLOW

AIR INLET

ECU COOLING

LPTCCVALVE

RACSBVALVE

HPTCCVALVE

Page: 73Figure 36 Engine Air System Schematic

Lufth

ansa

Tec

hnic

al T

rain

ing

For

Tra

inin

g P

urpo

ses

Onl

y

Engine A319 / A320 / A321CFM56-5B

75-20Air

Seite: Page: 74FRA US/T Kh August 03 L2 /3

75-20 ENGINE COOLINGHIGH PRESSURE TURBINE CLEARANCE CONTROLThe HPTCC system optimizes HPT efficiency through active clearance controlbetween the turbine rotor and shroud and reduces compressor load duringstarting and transient engine conditions.The HPTCC system uses bleed air from the 4th and 9th stages to cool downthe HPT shroud support structure in order to:

- maximize turbine efficiency during cruise.- minimize the peak EGT during throttle burst.

The HPTCC valve is located on the engine core section at the 3 oclock posi-tion.A thermocouple, located on the right hand side of the HPT shroud supportstructure, provides the ECU with temperature information.The ECU uses various engine and aircraft sensor information to take into ac-count the engine operating range and establish a schedule.To control the temperature of the shroud at the desired level, the ECU calcu-lates a valve position.This valve position is then sent by the ECU to the HMU, which modulates thefuel pressure sent to command the HPTCC valve.Two sensors (LVDT), connected to the actuator, provide the ECU with positionfeedback signals and the ECU changes the valve position until the feedbackmatches the schedule demand.

HPTCC VALVEThe HPTCC valve has integrated dual butterfly valves, driven by a single ac-tuator which receives the fuel pressure from the HMU servo valve.Each butterfly valve controls its own dedicated compressor stage air pick-up.One butterfly valve controls the flow from 4th stage compressor bleed while theother butterfly valve controls the flow from the 9th stage compressor bleed.The 4th stage air is mixed with the 9th stage air downstream of the valve.The two airflows are mixed downstream of the valve and sent through a ther-mally insulated manifold to the HPT shroud support, at the 6 and 12 oclockpositions.The actuator position is sensed by a dual LVDT and sent to both channels ofthe ECU.A drain port on the valve directs any fuel leaks towards the draining system.

Component DescriptionThe HPTACC valve has integrated dual butterfly valves driven by a single fuelpowered actuator. Position feedback to the ECU is provided by a dual channelLVDT installed on the actuator. One butterfly valve controls the flow from 4thstage compressor bleed while the other butterfly valve controls the flow fromthe 9th stage compressor bleed. The 4th stage air is mixed with the 9th stageair downstream of the valve.

Page: 74

Lufth

ansa

Tec

hnic

al T

rain

ing

For

Tra

inin

g P

urpo

ses

Onl

y

Engine A319 / A320 / A321CFM56-5B

75-20Air

Seite: Page: 75FRA US/T Kh August 03 L2 /3

DISCHARGEMANIFOLD

4TH/9TH STAGEOUTLET

Page: 75Figure 37 HPTCC System Location

Lufth

ansa

Tec

hnic

al T

rain

ing ENGINE

AIRA318/A319/A320/A321

CFM56-5B

75-21


HPT ACTIVE CLEARANCE CONTROL SYSTEM

System DescriptionThe HPTACC system regulates the HP turbine shroud support structure tem-perature by means of valves controlled by the Hydromechanical Unit (HMU).The HPTACC Valve has integrated dual butterfly valves driven by a single fuelpowered actuator.Fuel pressure is distributed in accordance with electrical signals sent by theElectronic Control Unit (ECU).One butterfly valve controls the flow from the fourth stage compressor bleedwhile the other butterfly valve controls the ninth stage.The fourth stage air is mixed with the ninth stage air downstream of the valve.

GENERALThe Hight Pressure Turbine Active Clearance Control (HPTACC) system usesbleed air from the fourth and ninth stages to cool the high pressure turbineshroud support structure.The purpose of the system is to : maximize turbine efficiency during cruise, minimize the peak Exhaust Gas Temperature (EGT) during throttle burst.

The HPTACC valve also discharges ninth stage air to unload the Hight Pres-sure (HP) compressor on engine start.Two Linear Variable Differential Transducers connected to the actuator and onethermocouple located on the right hand side of the HP turbine shroud supportstructure provide feedback signals to the ECU.

CONTROLIn accordance with various parameters such as N2 and T3, the ECU sendselectrical signals to the torque motor within the HMU to move the HPTACCvalves.When the engine is shut down the valves move to a failsafeclosed position.On engine start the HPTACC valve moves to the ninth stage bleed positionwhich unloads the compressor to improve engine acceleration.Above ground idle the position of the valves is determined by the closed loopshroud temperature control.The HPTACC actuator drives the butterfly valves to different positions.

The HPTACC Actuator drives the Butterfly Valves to different positions as fol-low :

ACTUATOR

STROKEMODE

4TH STAGE

BUTTERFLY9TH STAGE

BUTTERFLY

0%

FAILSAFE

37%

37 - 100%

100%

NO AIR

FULL 9TH

MIXED

FULL 4TH

CLOSED

CLOSED

INTERMEDIATE

FULLY OPEN

CLOSED

FULLY OPEN

INTERMEDIATE

CLOSED

T caseThe T case sensor measures the High Pressure Turbine (HPT) shroud supporttemperature.The temperature value is used by the ECU in the HPT Clearance Control sys-tem logic.lt is installed on the combustion case at the 3 oclock position, and consists of :

- a housing, which provides a mounting flange and an electrical connector.- a sensing element, fitted inside the housing and in contact with the shroudsupport.

NOTE: THE PROBE IS SPRING-LOADED TO ENSURE PERMANENTCONTACT WITH THE SHROUD SUPPORT.

Page: 76

Lufth

ansa

Tec

hnic

al T

rain

ing ENGINE

AIRA318/A319/A320/A321

CFM56-5B

75-21


HPTACCHPTACC DISCHARGE AIR

9TH STAGE AIR4TH STAGE AIR

HMU HYDRAULIC

POWER T CASE THERMOCOUPLE

FEEDBACK

ECU

T3N2

Page: 77Figure 38 HPTACC Airflow Schematic

Lufth

ansa

Tec

hnic

al T

rain

ing

For

Tra

inin

g P

urpo

ses

Onl

y

ENGINEAIR

A318/A319/A320/A321CFM56-5B

75-20


LOW PRESSURE TURBINE CLEARANCE CONTROL SYSTEM

GeneralTo ensure the best performance of the LPT at all engine ratings, the LPTCCsystem uses fan discharge air to cool the LPT case during engine operation, inorder to control the LPT rotor to stator clearances.lt also protects the turbine case from over-temperature by monitoring the EGT.The LPTCC system is a closed loop system, which regulates the cooling air-flow sent to the LPT case, through a valve and a manifold.The LPTCC valve is located on the engine core section between the 4 and 5oclock positions.The LPTCC system consists of:

-an air scoop-the LPTCC valve.-an air distribution manifold.-six LPT case cooling tubes.

The purpose of the LPT Active Clearance Control System (LPTACC) is to con-trol the thermal expansion of the low pressure turbine case during engine op-eration in order to: optimize the LPT rotor-to-stator (blade tip) radial clearances and thus get the best performance from the LPT at all engine ratings

To achieve these requirements the system: scoops a controlled amount of fan discharge air, flows it th

cfm 56-5b diff l3 (dec2003 cmp)

Documents