MPC 75 System Definition MPC AIRCRAFT A AC ACARS ACMS ACP ACU ADF ADM ADS AFT Ah AHRS ALT AMM AOA APU ATC B BAT BFE BITE BMC BRT BSCU CAA GAB CAPT CAUT CDU CFDS CIDS CMC CMS COND CPM

MPC 75 MPCAIRCRAFT

SYSTEM DEFINITIO

MPCAIRCRAFT

MPC 75

SYSTEM DEFINITION

MPCAircraftGmbH ReleasedKreetslag 102103 Hamburg 95West-GermanyJET-E System EngineeringRef: EE-T-90/025September 1990

SYSTEM DEFINITION MPCAIRCRAFT

CONTENTS

21 -Airconditioning22-Autoflight23-Communication24 - Electrical power25 - Equipment/Furnishing26-Fireprotection27-Flight controls28 - Fuel29-Hydraulic power30 - Ice and rain protection

31 - Instrumentation/Recording32-Landinggear33-Lights34 - Navigation35 - Oxygen36 - Pneumatic38 - Water/Waste45 - On-board maintenance49-Airborne auxiliary power70-Engine controls

ABBREVIATIONS MPCAIRCRAFT

A

AC

ACARS

ACMS

ACP

ACU

ADF

ADMADS

AFT

Ah

AHRSALT

AMM

AOA

APU

ATC

B

BAT

BFE

BITEBMC

BRT

BSCU

CAA

GAB

CAPT

CAUTCDU

CFDSCIDS

CMC

CMS

CONDCPM

CSM/G

ASendantSeat

Altemaling CurrentARINCCommunicationAddressing

and Reporting SystemAircraft Condittonirtg Monitoring

System

Audio Control Panel

APU Control Unit

Automatic Direction Flndlng

Air Data Module

Air Data System

Afterward

Ampere-hour

Attitüde Headlng Reference System

Attitüde

Aircraft Maintenance ManualAngle of Attack

Auxiliary Power Unit

Air Trafflc Control

Blue

Battery

Buyer Furnlshed Equipment

Built-in Test Equipment

Bleed Air Monitoring Computer

Brighl

Brake and Steering Control Unit

Civil Aviation Authority

Cabin

Captain

CautionControl Display Unit

Centralized Fault Display System

Cabin Integrated Data System

Centralized Maintenance Computer

Centralized Maintenance SystemCondtöon

Computer Modules

Constant Speed Motor/Generator

CTL

CU

CuR

DC

DET

DEU

DFDR

DME

DMPDU

ECAM

ECS

ECU

EE(E/E)

EFCSEFIS

EISEIVMUEIUELACELEV

ELSE M ERG

ENG

ESS

E/WD

EXT

FAC

FADEC

FBW

FCGC

FCU

FDIU

FL

Control

Control Unit

Cubtc Feet

Dkect Current

Oatectlon

Decoder/Encoder

Digital Flight Data Recorder

Distance Measuring Equipment

Display Management Processor

Display Unit

Electronic Centralized Aircraft Monttoring

Environmental Control System

Engine Control Unit

Electconic/ElectricalElectronic Flight Control System

Electronic Flight Instrument System

Electronic Instrument System

Engine Interface Vibration Monitoring Unit

Engine Interface UnitElevator Alleren ComputerElevation

Electronic Library SystemEmergencyEngine

Essential

Engine/Warning Display

External

Flight Augmentation Computer

Füll Authority Digital Engine Control System

Fly-By-Wire

Ftight Control and Guidance Computer

Flight Control Unit

Flight Data Interface Unit

Flight

FMCFMGC

FMSF/0

FQCC

FOIC

FQMS

FSDUFWDFWS

GGGCUGLC

GPCUGPSGPWS

GRDGS

H

h

HF

HP

HS

HYDR

ICAO

DG

ILS

in

IOM

IPC

IRS

Flight Management ComputerFlight Management and

Guidance ComputerFlight Management SystemFirst OfffcerFuel Quantity ControlComputerFuel Quantity Indication ControlFuel Quantity MeasuringSystemFlre Source Detection UnitForwardFlight Warning System

GalleyGreenGenerator Control Unit

Generator Une ContactorGround Power Control UnitGlobal Positioning SystemGround Proximity WarningSystemGroundGlide Slope

Hfch

hour

High Frequertcy

High Pressure

High Speed

Hydraulic

International Civil Aviation

Organisation

Integrated Drive Generator

Instrument Landing System

inch

Input and Output Modulea

Illustrated Parts Catalog

Inertial Relerence System

Page:1

ABBREVIATIONS MPCAIRCRAFT

L Lavatofy

1

LA

b

LCD

LOG

LGCIU

LH

LOC

LP

LRU1 TLT

M

MAGN

MCDU

MCU

MDDU

MEL

MKR

MLI

MLS

MSU

MAI

NAV

ND

NCd

OMS

PCU

PFD

Ph

PSM

litre/liter

Linear Accelerometer

Pound

Liquid Crystal Display

Landing

Landing Gear Control Interface

Unit

LeftHand

Locallzer

Low Pressure

Une Replaceable Unit

Ught

Motor

Magnete

MuWpurpose Control Display Unit

Modular Concept Unit

Multipurpose Disk Drive Unit

Minimum Equipment List

Marker

Magnetic Level Indicator

Microwave Landing System

Mode Selector Unit

Nacelle Anti Ice

Navigation

Navigation Display

Nickel Cadmium

On-Board Maintenance System

Power Control Unit

Prima/v Flight Display

Phase

Power Supply Module

PSU

PWR

QAR

RA/RALT

RATRFRHRMCU

RMPRS-232

S

SATCOM

SD

SEC

SEL

SELCAL

SERV

SFCC

TAS

TAT

TDB

TCAS

TH3

TLA

TRU

US

USgal

VHP

Passenger Service Unit VOR VHF Omnidirectional Range

Power v/s Vertical Speed

Quick Access Recorder w Wardrobe

W WidthRadio Altimeter ,„ . . . ... _ . . .

WAG Waming Acquisition ConcentratorsRam Air TurbineRadio Frequency *AI Wing Anti Ice

Right Hand WXfl Weather Radar

Remote Magnetic Compensation

Unit Y YellowRadio Management Panel

Seviell Data Bus Standard

Stowage

Satellite Communicatkxi

System Display

Spoiler Elevator Computer

Setect

Selective Calling

Servo

Slat and Flap Control Computer

True Air Speed

Total Air Temperature

To-Be-Determtned

Traffic Collison Alert and

Avoidance System

Trimabto Horizontal Stabttlfzer

ThrotHe Lever Angle

Transformer Rectifier Unit

United States

United States gallon

Very High Frequency

Page :2

ENVIRONMENTAL CONTROL SYSTEM MPCAIRCRAFT

Contents

1- Air conditioning System

2 - Cabin pressure control System

21


Air conditioning pack schematic

MAIN HEAT

EXCHANGE

RAM AIR

OUTLET

NACA INLET

WATER

wirciim

WATER

EXIRACTOR

GROUND COOLING FAN + CHECK VALVE

21

Figure 21-2September 1990 Page: 21.03


Air distribution

MAIN SUPPLY DUCTS

RISER DUCTS

7 PER SIDE IN WING AREA

21INDIVIDUAL

AIR SUPPLY

COCKPIT SUPPLY

MAIN SUPPLY DUCTS

METERING RESTRICTOR

CABIN AMBIENT AIRTO UNDERFLOOR AREA

September 1990Figure21-3

Page:21.04

ENVIRONMENTAL CONTROL SYSTEM <§> MPCAIRCRAFT

21

2 - Cabin pressure control System

The cabin pressure control System controls the cabin pressure and the change rate to provide

maximum passenger comfort and safety.

The pressure control System consists of an automatic System and a manual standby System

to control the outflow valve.

Provision for an optional second pressure Controller isprovided (seefigure21-4). Excessive

positive or negative cabin overpressure is prevented by two pneumatically controlled safety valves.

Automatic prepressurisation to prevent sudden pressure fluctuation (bumps) on take off and

landing, is provided in AUTO mode.

September 1 990 Page :21 >05


Cabin pressurisation control schematic 21

AIR DATA

SYSTEM

PRESSURISATIONCONIROLLER 1

PRESSURISATIONCONTROLLER 2

(OPTION)

PRESSURISED AREA

GABMAN V/S CTL MODE SEL

lOU7FLOW VALVE

(BUT1EHFLY TYPE)

SAFET*VALVES

y HUIUU UUlVt(3RD MOfÜ« OPTIONAL)

0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0

UNPRESSURISED AREA

September 1990Figure 21-4

Page: 21.06

AUTO FLIGHT MPCAIRCRAFT

22

Contents

- General

- Flight management

- Auto flight

AUTO FLIGHT €> MPCAIRCRAFT

Auto Flight Modes - General 22

• AP/FD can either be selected at the flight control unit (FCU)

or provided by the FMS according to active flightplan

• Basic Flight Control mode is always in accordance with the electrical flight control System (EFCS)

and always engaged


Autopilot / Flight Director Modes

Selectable upper modes

1. Longitudinal Modes

22

- altitude hold- altitude acquire-flight levelchange- profile- climb- descent-flight path angle

ALTALT ACQFLCHPROFCLBDESFPA

Lateral Modes

- heading- track- navigation (incl. VOR)

HDGTRKNAV

3. Approach Modes

- approach- go around- take-off- localiser

APPRGATOLOG


Autothrust Modes

• A/THR permanently armed

• Autothrust modes in association to autopilot / flight director mode

AP/FD

ALT, ALT ACQ, FPA

DES, CLB, PROF

FLCH

APPR -FINAL-G/S-FLARE

TO/GA

A/THR modes

SPD/MACH

SPD/MACH, THR

THR

SPDSPDRETARD

A/THR functionoverridden

SPD / MACH speed/mach mode selectable

THR thrust mode, limit thrust according to selected thrust rating

Alpha Floor Protection overrides all modes with GA limit thrust if A/C flight envelope is exceeded

AUTO Fl IfiHT W MPCM U l \J rLIvan l " AIRCRAFT

Flight Management System (FMS) 22• The aircraft is equipped with a single Flight Management System,

with a second Flight Management Computer available äs an option.

• The system is controlled via two Multipurpose Control and Display Units (MCDU). Many frequent

and complex tasks are performed by the Flight Management Computer to

increase safety and punctuality of the flights.

Flight Planning / Flight Plan Execution

Take-off and Landing Data/Cruise Speed Calculation

Radio tuning / Radio navigation

On-board Maintenance System (QMS) support

Data link function (ACARS)

In case of failure of the flight management Computer, the MCDUs have to perform

a back-up function for navigation tasks, such äs radio navigation.

September 1990 Page :22.01


Flight managementSystem architecture

,September 1990

FLIGHT MANAGEMENT

COMPUTER

ARINC 429 ARINC 429

basic

optional

Page :22.02


System architecturePilot controlsand other aircraft sensors

22

** Warnings, Indications,Maintenance, Recording

NoseWheelSteering,Auto Brake

Actuators

FADEC

FADEC

Actuators

NoseWheelSteering,Auto Brake

Wamings, Indications,Maintenance, Recording

Pilot controlsand other aircraft sensors

September 1990 Page:22.03

COMMUNICATION MPCAIRCRAFT

23

Contents

- RF communication

- Audio communication

- Cabin communication

- Antenna location

COMMUNICATION ^ MEß

1 - RF communication: 23Main components are two VHF tranceivers and two Radio Management Panels (RMP). A third VHF transceiverand an HF transceiver are options, just äs the Aircraft Communication Addressing and Reporting System (ACARS)and a Satcom System. The ACARS uses only the third VHF.The Radios will be tuned via the two RMPs. Also tuning through the Flight Management System (CDU) ispossible.Bussystem is to the Standard ARINC 429 norm.

2 - Audio communication

The audio management unit handles all audiosignals.The System is controlld by three Audio ControlPanels (ACP), where transceiver selection, volume control and other functions are possible. The Signalsfrom all microphones (except Cockpit Voice Recorder microphone) and from other sources are collectedin the Audio Management Unit and distributed to loudspeakers or, for cabin communication, to CIDS.

3 - Cabin communication

The cabin communication mainly is done through a Cabin Integrated Data System (CIDS). The signels aretransceived in a digitizform ed in a databus System between the Decoder/Encoder Units (DEU) and theCIDS director located in the avionic bay. The CIDS director also interfaces with the audio management unitsfor interphone functions.The System is controlled via the attendant panel. A second attendant panel is available äs an Option.

September 1990 Page: 23.01


RF communicationSystem architecture

MLS1

September 1990

ARINC 429Audio line

23

ADF1

DME1

VOR/MKR

LLt>t>t>-oo

OAudio Management Unit

RF control

RFcommunication

1

VOR2

1

DME2

1

ADF2

MLS2

Page: 23.02


Audio communicationSystem architecture

23

Handmicrophone^

LoudSpeaker

SideStick

Oxyganm äs k

Jackbox

CAPTAGP

Audio Management Unit

"

F/0ACP

rf

Handmicrophone

LoudSpeaker

SideStick

Oxygenmask

Cockpitcommunication

^ ^ JackO O box

Hand

DEUB

DEUB

— ( i

CIDS

...... . l __.FWD

ATTNDPANEL

2ndATTNDPANEL <

L J

i

1

>

NS || F S B ] | Sign

Loud-speaker

Cabincommunication

Lighta



Antenna location MPC 75-10023

LOC #MLS t #TCAS VHF i # GPS 1/2 3 *SATCOM # HF VOR 1/2

0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0

* SPACEPROVISION

wx

# STRUCTURALPROVISION

GS # MLS 2 DME 1/2 # TCAS ATC 1/2 VHF 2 * OMEGA * FL. PHONE MKR RALT l RALT 2

September 1990 Page: 23.04/1

ELECTRICAL POWER MPCAIRCRAFT

24

Contents

1 - General

2 - Normal flight configuration

3 - Abnormal flight configuration

4 - Aircraft ground supply

5 - Electrical power generation control

September 1990

ELECTRICAL POWER AIRCRAFT

1. General: 24The electrical generation network consists of the following units:

AC Generation:

2 main integrated drive generators (IDG), one per engine, nominal 60 kVA. The IDG provides a 115 / 200 V,400 Hz AC supply. It consists of an oil spray cooled three states AC generator driven at constant speedby a hydromechanical constant speed drive.1 APU generator, nominal 60 kVA, to provide a seif contained source of ground power.This generator may also be used in flight to compensate the loss of main generators.1 ground power connector, nominal 60 kVA.1 emergency hydraulic driven generator (CSM/G), nominal 5 kVA.1 static inverter, nominal 350 VA, single phase 115V/ 400 Hz to provide a source of AC power from the batteries.

DC Generation:

3 transformer rectifier units (TRU), nominal 100 A at 28 V DC. Two of them are used to feed a DC maindistribution System. A third TRU supplies via the AC essential power the DC essential network.2 NiCd - batteries, nominal capacity 40 Ah each, provided for autonomous power for APU starting,DC equipment during transient switching of electrical power.


J

ELECTRICAL POWER

2. Normal flight configuration: 24

In normal flight condition each IDG supplies an AG distribution network via generatorline contactors (GLC). These two networks are not parallel / isolated mode.

Under normal conditions:Distribution network No. 1 comprises:

- AC MAIN BUS 1- AC ESS BUS , which is supplied by AC MAIN BUS 1- GALLEY BUS 1-DC MAIN BUS 1 viaaTRU

Distribution network No. 2 comprises:- AC MAIN BUS 2- GALLEY BUS 2-DC MAIN BUS2viaaTRU

Two 40 Ah batteries are connected to the DC EMERGENCY BUS.

Each battery has its own busbar (BAT BUS 1 and BAT BUS 2) which are permanently connected.

September 1990 page. 24.02

ELECTRICAL POWER

3. Abnormal flight configuration: 24

Any one of the three generators is able to supply the total technical load and part of the commercialload. In case of total loss of all main AC power generation a hydraulic driven emergency generator( CSM/G ) with 5 kVA provides power for no limitation on usable time . ( CSM/G driven by the greenhydraulic circuit / engine driven pump or RAT ).

4. Aircraft ground supply:

On ground the complete network can be supplied either from the APU generator or fromthe ground power unit ( EXT - PWR ) up to a maximum of 60 kVAFor ground servicing Operation, it is possible to supply only the GND SERV AC / DC BUSdirectfrom the EXT - PWR connection.

5. Electrical power generation controls:

The electrical power generation and distribution control, indication and monitoringis operated through the Electronic Centralized Aircraft Monitoring System (ECAM) andthe overhead control panel.


ELECTRICAL POWER MPCAIRCRAFT

System Architecture

i L

HYDR. =i CSM/G

< FUEL PUMPS)

AG MAIN BUSl

ELGALLEY BUS 1

AG ESS BUS

Tr1AC SHED ESS BUS

AC EMERG BUS

AC GND SERV BUS

GALLEY BUS 2Ur

AC MAIN BUS 2

TRU1

TRU2

24

DC MAIN BUS 1

DC SHED ESS BUSlr

DC ESS BUS

BAT BUS 1Lr

.BAT.1

HI-

DC EMERG BUS

rBAT BUS 2

.BAT.2

HI-DC GND SERV BUS

DC MAIN BUS 2


Electrical Power MPCAIRCRAFT

Overhead-Panel

GCU1GEN1

September 1990

Emerg.Panel

ControlPanel

CMCMANAGEMENT

24System ManagementUttofabbravfation

CMC ConlraHzedMainleruuK« ComputerCSH/OCU ConMamSpMdMotor/GwieratorControlUnItECAM B«*onfc<^rrtraUzedAlrcnft MonitorOCU Generator Control UnitOPCU Qround Power Control UnitMCOU MuWpurpoMControl and Display Unlt

••M > Dtocroto Signal

AfllNC429databus

ECAM(Status Indication)

GCU2GEN2

GCU3APU

CSM/G CU

Page: 24.05

EQUIPMENT/FURNISHING MPCAIRCRAFT

Contents

1 - General

2 - Cockpit

3 - Cabin

4 - Underfloor cargo holds

25


EQUIPMENT/FURNISHING

1. General

The MPC75 interior equipment/furnishings have been designed to provide a comfortable environmentfor the flight crew and passengers.This has been achieved by giving special attention to three main areas äs

- passenger comfort and cabin Styling

- carry-on baggage capacity

- underfloor compartment capability.

The major parts are provided within the pressurized section of the fuselage, äs follows:

- cockpit

- cabin

- underfloor cargo holds

25



2. Cockpit 25The cockpit has been designed for two-man Operation with facilities for a third occupant.The flight compartment arrangement is shown in Fig. 25-1.The captain and first officer seats are seat rail mounted and fuliy adjustable in all directions.Each seat is fitted with an inertia-type five-point harness that allows a pilot to reach allaircraft controls without discomfort.

An observer seat is provided which occupies a position aft of the pedestal when in usebut is normally stowed against the right-hand aft wall of the flight compartment.The seat is provided with a fixed-type three-point harness.

The planview of the cockpit is shown in Fig. 25-2.For detail of the Instrument panels refer to chapter 31.

September 1990 Page: 25.03-1 /2


Cockpit Design 25

The main features:

- Sidestick Controllers leave the main Instrument panel unobstructed

- EFIS/ECAM/FWS Integration- Improved f lexibility through switching of six Fiat Panel Display Units- Standby Instrument display with similar Images

Most features evolved directly from A 320 such äs:

- "Lights out" philosophy on overhead panels

- "Need to know" concept for Information presentation- Monitoring of Systems through the ECAM-system

- Monitoring of Computers through Centralized Maintenance System (CMS, CFDS in A 320)

i

All control panels are within easy reach of the two crew members

A 3rd crew member seat aft of the pedestal offers maximum visibility over all panels

September 1 990 Page : 25.03-2/2


Fig. 25-1 Forward view

September 1990

25

Page: 25.04


Fig. 25-2 Planview3rd seat Circuit breaker rack Galley

September 1990 Attendant seat Stowage

25

Page: 25.05


3. Cabin 25The aim in the design of the Standard cabin has been to offer a higher Standard of comfortrelative to existing single-aisle aircraft in that class.This target, combined with the requirement to carry Iower and upper deck cargo either in bulkform or in Containers, results in the choice of a blended-double bubble fuselage cross-section havinga width of 3.45 m (1 35.8 in) and a height of 3.67 m (1 44.5 in).

The passenger cabin itself has a maximum width of 3.23 m (1 27 in) and a height of 2.1 0 m (82. in).

The MPC75 cabin interior has been designed according to the latest industrial design conceptswhich makes the cabin particularly spacious.All functional and decorative items such äs Windows, side-walls and ceiling panels and overheadstowage compartments have been longitudinally divided into two frame pitch (40 in) units. The factthat the cabin is parallel over most of its length and with a constant frame pitch, except around thewing box area, results in a minimum number of different furnishing panels.


EQUIPMENT/FURNISHING AIRCRAFT

Standard Seat Layout 25

In its baseline configuration, the MPC75 provides single-class accommodation for 91 passengersat 32" pitch in a single-aisle, five-abreast, arrangement (see Fig. 25-3).

Other interior configurations are available for enhanced cabin flexibility.

The four seat rails running the length of the cabin allow 4 or 5 abreast seating (see Fig. 25-4).Standard and optional toilet and galley positions are provided at each end of the cabin either sideof the entrance areas (see Fig. 25-6 and Fig. 25-7).

The seat rails are spaced apart so äs to give approximately equal space under each seat, in5 abreast seating, for carry-on baggage (see Fig. 25-5).

Fixed galleys and lavatories are attached to the aircraft structure by bolts at hard points.

Lavatories are manufactured äs preassembled wet cells with an integrated floor panel designedäs a bowl for leakage protection.



Fig. 25-3 Standard seat layout 25

Baseline seat layout: 91 seats 32" pitch

A: Attendant seatG: GalleyL : LavatoryS: StowageW: Wardrobe



Fig. 25-4 Alternative seat layouts

Economy seat layout: 89seats 32"pitch

25

A: Anendant seatG: GalleyL : LavatoryS: StowageW: Wardrobe

September1990

Mixed class layout: 8 seats 36" pitch74 seats 32" pitch

Page: 25.09

EQUIPMENT/FURNISHING * MBQ

Cabin cross-section 25

Each seat is fitted with a two-point seat belt and gives a high degree of comfort with the followingfeatures:

- Lightweight aluminium structure- Reclineable back- Füll width table- Ashtray built into the armrest- Seat back pocket- Stowage for lifevest.

Passenger Service and information panels are provided above each seat row and are designed andinstalled with special consideration for passenger convenience.The panels are fully adjustable to match variations in seat pitch where optional cabin layouts arechosen. The features provided on each panel are:

- Readinglight- Individual air outlets- Attendant call buttons- Fasten seat belt/no smoking sign- Loudspeakers- Emergency oxygen.



Fig. 25-5 Cross-section

5 - abreast arrangement

all-economy class

63 " - 20 " - 42"

4 - abreast arrangement

first-class

50" - 23 " - 50"



Fig. 25-6 Location of galleys

The Standard aircraft is equipped with one galley in the forward cabin (G1).

Standard options are located in the forward cabin (G2) and in the aft cabin (G3, G4.).

25

September 1990

Standard - G1Option - G2, 63, G4

Page: 25.12


Fig. 25-7 Location of lavatories

The Standard aircraft is equipped with two lavatories, both in the aft cabin.

Standard options are located in the forward and rear cabin areas.

25

Standard - E,FOption - A,C



Fig. 25-8 Location of cabin attendant seats

The Standard aircraft is equipped with two cabin attendant seats, one forward cabin (A1) and onein the aft cabin (A2).

Standard options are located in the aft cabin areas (A3.A4).

25

Standard - A1, A2Option - A3, A4



Overhead stowage compartments 25

For passengers' carry-on luggage the cabin is equipped with lateral overhead stowage compartments runningalong each side of the cabin above the passenger seating area. Each compartment is approximately 80"(2032 mm) and can carry a load of TBD Ib (TBD kg).

In the Standard all-economy seat layout with 91 passengers the total stowage volume is approximately 184 cu.ft.(5.2 m3). Thus there ist approximately 2 cu. ft. and TBD Ib (TBD kg) per seat.

A handrail is incorporated into each stowage compartment.

The optional overhead stowage bin has in the Standard all-economy seat layout with 91 passengers a total stowagevolume of approximately 217 cu. ft. (6.14 m3). Thus there is approximately 2.4 cu.ft. and TBD Ib (TBD kg) per seat.

Fig. 25-9 Standard bins Fig. 25-10 Optional bin



Emergency equipment provisions 25The MPC 75 meets the JAR/FAR safety requirements for passenger emergency evacuation.The speed of evacuation is enhanced by:

- 20" wide aisle- 72" (H) x 34" (W) door fwd. left side and 55" (H) x 30" (W) door fwd. right side- 72" (H) x 30" (W) door aft left side and 55" (H) x 30" (W) door aft right side

The main doors are fitted with single lane slides. All slides are of inflatable type using cylinder compressed air.Escape slide illumination is achieved with lights integrated into the slides and powered from the aircraft emergencylighting System.

As an option, all passenger/service doors can be equipped with slide rafts instead of escape slides, which arefurnished with survival kits. The slide rafts can be disconnected from an inoperative door, moved to anotherdoor and mounted to the structural slide raft fittings.

In easily accessible positions in the passenger compartment, the following safety equipment is installed:- fire extinguishers,- protective breathing equipment,- portable oxygen units,- flashlights,- crash axe,- life vests,- megaphone,- first aid kits, etc.



Fig. 25-11 Escape slide arrangementType C door

30"x55"exitwithsingle lane slide orslide-raft.

Type C door

34"x72"exitwithsingie lane slide orslide-raft.September 1990

25Type C door

30"x55"exitwithsingle lane slide orslide-raft.

Type C door

3o"x72"exitwithsingle lane slide orslide-raft.

Page: 25.17


4. Underfloor cargo holds 25The forward and aft cargo holds are designed for bulk loading and to meet the requirements of"class D" regulations. Smoke warning System is available äs an Option, äs well äs a cargoVentilation System.

The basic doors are outward opening.

A semi automatic "sliding carpet" cargo loading System is available äs an Option.



Fig. 25-11 Location of cargo hold compartments 25

0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0

rrn


FIRE PROTECTION MPCAIRCRAFT

Contents

2

3

4

5

6

General

Fire Protection Engines and APU

Smoke Detection Electronic Bay

Cargo Hold Smoke Detection

Portable Fire Extinguishers

Lavatory Fire Protection

26

September 1990 Page 26.01

FIRE PROTECTION

1. General 26Aircraft fire and smoke protection Systems are provided for:

- Engine and APU, fire detection and extinguishing Systems- Electronic bay smoke detection System- Automatic waste bin fire extinguishing System and smoke detectors in lavatories

Portable fire extinguishers are installed in the flight deck and the passenger cabin.

2. Fire Protection Engines and APU

2.1 DetectionThe detection loops for the engines and the APU are of pneumatic type. All detection loops are duplicated(Loops A and B). The Fire Source Detection Units (FSDU 1/2) 'acts' äs the 'central manager' between theSystems and the EIS äs well äs the overhead panel control and indication devices.

September1990 Page 26.02

FIRE PROTECTION

2.2 Extinguishing

The extinguishing System for each engine consists of two extinguisher bottles (APU one bottle), instaliedin the pylon area (engines). The APU extinguishing bottle Installation is close to the APU compartment.The bottles can be discharged by means of wired cärtridges.The APU fire extinguishing System is provided with an automatic sequence for fire extinguishing in anemergency case on ground if qualified personnel is not present in the cockpit.Extinguishing agent: HALON 1 301

3. Smpke Detection Electronic BaySmoke in the electronic bay is detected by Ionisation type smoke detectors, instalied in the air extractionduct. The smoke detector is connected to the FSDU and warnings are given on the Smoke Detection Panel(Overhead Panel) and associated Information will appear on the EIS.

4. Cargo Hold Smoke DetectionCan be provided äs an Option.

September1990 Page 26.03

FIRE PROTECTION • MEG

5. Portable Fire Extinguishers

Portable fire extinguishers are instalied at appropriate location in the flight and passenger compartment.Extinguishing agentwill be HALON 1211.

6. Lavatory Fire Protection

A smoke detector is instalied in the air extraction duct from each lavatory. The detectors are connectedto the FSDU. Audio and visual warnings of lavatory smoke are provided in flight compartment and at theattendants panel.Fire extinguishers will discharge automatically.


FIRE PROTECTION MPCAIRCRAFT

(OPTION)

ENGINE 1FIRE/OVERHEAT

DETECTION

ENGINE 2FIRE/OVERHEAT

DETECTION

APUFIRE/OVERHEAT

DETECTION

E.E.BAYSMOKEDETECTION

LAVATORYSMOKEDETECTION

CARGO HOLDSMOKEDETECTION

26

FIRE SOURCE DETECTION UNIT

(FSDU1)

l i iDISCRETE FIRE SIGNALS FIRE SOURCE DETECTION UNIT

( FSDU 2)

7\G

SYSTEM

CENTR.MAINTENANCESYSTEM

OVERHEAD PANELFIRE CONTROL PANELENGINE 1 APU ENGINE Z

FIREEXTINGUISHINGCONT.MONITORING

PRESSURE/SQUIBS

ENGINESAPU

CABINSMOKEWARNINGFACILITIES

September 1990 FIG. 26-00 Fire Protection Page 26.05

FLIGHT CONTROLS MPCAIRCRAFT

27

Contents

*

General

Architecture

Flight control and guidance Computers

Computer volume comparison

September 1990

L

FLIGHT CONTROLS

General 27M PC flight control is achieved by conventional surfaces which are f ully powered by hydraulic actuators. The flightcontrol System is designed äs a Fly-By-Wire (FBW) System based on two different types of digital Computers.

The aircraft' s primary flight control Systems - ailerons, roll spoilers, elevators, Trimmable Horizontal Stabilizer(THS) and rudder - control pitch, yaw, and roll flight attitudes. The secondary System comprises leading edgeslats and trailing edge flaps for high lift configuration, speed brakes for deceleration, and lift dumpers to improvebraking efficiency.

Some setected features of MPC flight control are:

• 3-axes FBW in pitch, roll, and yaw.

• Mechanical linkage to one rudder actuator and the THS

• Modern cockpit design with side stick Controller.

• Modern control laws for improved handling and safety.• Aircraft dispatchable after any single electrical flight control Computer failure or any single failure affecting

aileron, Spoiler or THS control.• No operational reduction after a single hydraulic failure in flight.

• Easy Identification of System failures with ECAM / Flight Warning System• High concentration of Computer tasks in each Computer module reduces EFCS-complexity. Flight control

Computer capacity can be reconfigured.

September 1 990 Pa9e : 27-01


Architecture

B

A2-—B1

GND-Spoiler

Roll"Speedbrake PCU

B B

r Flaps

A1 A2

PCU

A1 B1 B2 A2 B1 B2

•GND-Spoiler

•RollSpeedbrake

Flaps~I

A2 B2 B1 A1

27

B

B2

FCGCA

IOM1

l

IOM2

1

l 3 l

N

B

KflL_d_l

September 1990

Pitch trim handwheels —

B2

B1

A2

;>A1

Motor 1

Motor 2

Motor 3 i

B G

B

A1

B1

THS ActuatorG

A1A2

B1B2

B1B2

A1A2

A2

B2

r

B

G

RudderPedals Page: 27.02


Sidestick positionsRudder pedals positionPitch trim controlRudder trim controlSpeedbrake control lever positionSlat / flap control lever position

Air data moduls 1,2,3AHRS 1,2,3RA 1,2LGCIU 1,2Wheelspeed tachometersHydraulic pressure sensorsAccelerometersYaw rate gyrosEIUFCUFMCAutopilot sensors (ILS/VOR/DME ...)•Thrust lever position

Rudder pedalsPitch trim handwheels

ünks_

Flight ControlGuidance Computers

27. Nosewheel steering,

Auto brake. Warnings, Indications,

Maintenance, Recording

Elevators

Ailerons

r Rudder

rTrimmable Horizontal

Stabilizer

l| Tl ll L

Spoiler

l lJ l

l

Flaps/Slats

September 1990 Page: 27.02a


Flight Control and Guidance Computer (FCGC) 27

IOM

1

PSM

1

CPM

1

IOM

3

PSM

3

CPM

2

IOM

2

PSM

2

15MCU



Flight control and guidance Computers

Flight control, envelope protectionand flight guidance cpmbutation isachieved by two dissimilarComputer Systems: FCGC A andFCGCB.

• Each FCGC comprises twoComputer Modules (CPM)each of them similar in hardwarebut dissiuiilar in Software, and threel/O-Modules (IOM) which connectthe FCGC to the pilot1 s controlsthe actuators, and the otheraircraft Systems and sensors.

Each CPM consists of twoComputing lanes which check eachother.

FCGC A

27

FCGCB



FCGC failure analysis

Normal Operation

Ist Failure

2nd Failure

3rd Failure

Normal Operation

4th Failure

5th Failure

6th Failure

Flight control Flight guidance

Flight controli

Flight control

TTTIFCGC A

27Flight control

Flight control

Flight guidance

Flight guidancei

Flight control

T r TTFCGC B

September 1990Page: 27.05


Computer volume comparison

MCDU

FCU

32

Autoflight

ComputerVnli impMCU

FlightControl

September 1990 50

A320

2

1

FMGC-

2x8

FAC/V/Q«o

SFCC2x5

ELAC2x6

SEC3x8

FCDC2x2

27MPC75

2

1

12i FMC2

6FMC1

6

FCGCA/B2x15

30

Page : 27.06

FLIGHT CONTROLS

Roll control

MPCAIRCRAFT

27

FCGC A FCGC B

IOM| IOM ! lOtol IOM| IlOM ! ION/I1 1 o ' t ' 1 l l 9 *}1 2 1 J 1 ' 1 L _J— > h—rl l-r-rl LT — 1 L,— pl L,— rl h- — '

S\r ffAII f S S

Spoiler

Aileron

1B

t

G

1

4 3A A

t tB G

) )

2 LUt -

Y

t

Spoiler

[J] 2A

iY

J

3A

t

G

)

4

/B

)

Aileron

t tB GA i

) )



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FCGCA

IOM IOM

f f_

I

FCGCB

[iöiüjl n \3_l

THS actuator

' S X

Electricmotors

Hydraulicmotors

THS

Elevator Elevator

fT TTHydraulicactuators



Yaw control

Trim actuator Artificialfeel

C

27



Ground spoiler control 27

SpeedBrakeLever

FCGCA

IOMI1

IOM2

L,-!

FCGCB

IOM1

X X

IOM2

X /

[iÖM|

Throttle

September 1990

Spoiler4 3 2 1

Weight On Wheels

Spoiler1 2 3 4

iiil IM lB

T T T T T T T T

Page: 27.10

FLIGHT CONTROLS

Speed brake control

MPCAIRCRAFT

27

Lever

Spoiler

a mm mt±l

Spoiler

["DOGGEttrt

FCGCj

IOM1

X

IOM |l<

2 !(

s X

A

DM!0 1AJ

FCGC

IOM1

IOM| !l<

2 1!

B

DM|

r

J J J J J T



Slat / f lap control

FCGCB

27

PCU : Power Control Unit

1 : Flap/slat lever2: Command sensors3 : Wing tip brake4 : Asymmetry position

pick - up

5: Gearbox6 : Rotary actuator7: Flap carriage8: Feedback position

pick - up

9 : Torque shaft

10: Instrumentation positionpick - up

to ECAM and flight warning system


FUEL MPCAIRCRAFT

28

Contents

1 - Storage

2 - Vent System

3 - Distribution

4 - Refuel / Defuel

5 - Quantity Indication


FUEL• VJE-L AIRCRAFT

281. Storage

Fuel is stored in two tanks, one in each outer wing spar box. The tanks are an integral part of the wing structure.2% expansion space is provided in each tank.The engines are fed from collector tanks between rib 1 and rib 2, which are protected by flapper valves.Water drain valves are installed at the Iowest points of each tank.A vent surge tank is located near each wing tip, which collects fuel spilled into the vent System.

Total wing fuel: 10,000 l (8,000 kg; density: 0.8 kg/l)

The centre wing box is free for a later capacity enlargement.

2. Vent System

Open vent System with non-icing NACA air inlet in outboard vent surge tank.Flame arrestors protects surge tanks against ground fire.Fuel spilled into the surge tanks is discharged back into the main tanks.

Optional centre tank vented via left hand vent surge tank.


AIRCRAFT

3. Distribution 28

Each engine is fed from two electrical bopster pumps (1 1 5 V AC 3-Ph.) located in the adjacent collector tank.Pumps are canister housed for removal without draining the tanks.One pump can supply one engine under all conditions.Pumps are controlled by pushbuttons on the overhead fuel control panel and monitored by pressure Switches,sensing at the outlet of each pump.Fuel shutoff valves with double electrical actuators are mounted at the front spars. The isolation of an engineis controlled by the engine master switch and the fire pushbutton.

A crossfeed valve with a double electrical actuator allows fuel supply from both tanks to one engine or fromone tank to both engines. The crossfeed valve is controlled by a pushbutton on the overhead fuel control panel.

Optional center tank with two pumps. Pumps identical to the wing tank pumps.

The APU is supplied from the LH supply System. On ground, a battery powered APU Starter pump can be used.The pump is canister housed. Pump control automatically by pressure switch.The APU fuel line is shrouded in pressurized areas of the fuselage and drained overboard.An APU shut-off valve with double actuator is mounted at the rear spar. Valve control is by the APU masterswitch and the APU fire pushbutton.

System control on overhead fuel control panel.System Status and failure indication on Electronic Instrument System (EIS) displays.


FUEL• UCL AIRCRAFT

4. Refuel / Defuel 28

Single pressure refuel coupling (2.5" Standard) under RH wing leading edge.Max. refuel rate 1200 l/min with 50 psig nominal pressure at the coupling.Refuel valves are solenoid actuated with manual override. Refuel valves canister housed.No fuel remaining in refuel pipe due to air inlet valve.

Automatic refuel distribution by use of Fuel Quantity Measuring System.Independent refuel shut-off by 'High Level' sensors.Refuelling possible by using battery power only.

Refuel panel located in RH wing fairing. It contains refuel and defuel valve control Switches, fuel preselector,digital indicators for preselected fuel, total fuel and fuel in each tank and 'High level' indication lights.

Cockpit refuel panel optional, allows preselection of required fuel quantity.

Defuel valve electrically actuated, manually controlled from refuel panel.Defuelling and fuel transfer from one tank to another possible by using the booster pumps.

Gravity filling point provided for each wing tank.


FUEL• i-lt- L AIRCRAFT

5. Quantity Indication 28

Fuel Quantity Indication

Capacitance type measuring System with linear probes.Computation by dual channel Fuel Quantity Control Computer (FQCC).Densitometer in each wing tank.Individual connection of all probes with FQCC for BITE function.ARINC 429 connection to Refuel Panel and aircraft Systems.Digital indication of fuel quantity on EIS and Refuel PanelDigital indication of fuel temperature on EIS, one temperature sensor located in each wing tank.

Level Sensing

'Low level' warning and 'high level' refuel shut-off controlled by independent level sensing stage, integratedin FQCC. One 'Iow level' sensor and one 'high level' sensor in each tank.Independent power supply.Sensor BITE control via FQCC.

Secondary Fuel Quantity Indication

Magnetic Level Indicators (MLI) can be used on ground in connection with an attitude monitor andcorrection charts.

Provision for centre tank equipment similar to wing tanks (except temperature measurement).


FUEL MPCAIRCRAFT

Fuel System Schematic

/ ENG \8

FLOATVENTVALVE

September 1990

SINGLE POINTPRESS. REFUELING

SURGETANK

NACA FLUSHAIR INTAKE

Page 28.06

HYDRAULIC SYSTEM MPCAIRCRAFT

29

Contents

2

3

4

5

General

Hydraulic System description

Installation and distribution

Indication and control

Ground handling

October1990 Page 29.01

HYDRAULIC SYSTEM AIRCRAFT

1 - General

The hydraulic System powers the hydraulic actuation for the flight control System, the landing gearand the braking system. It consists of three fully independent hydraulic circuits, äs there is no manualmode for the flight control.

The system failures including dormant failures shall not prevent continued safe flight andlanding unless they are extremely improbable.

3000 psi supply pressure is delivered from piston pumps with variable displacements. All pumps areof conventional types.

The fire-resistant hydraulic fluid is of phosphate ester type.

29

October1990 Page: 29.02

HYDRAULIC SYSTEM

OQ2- Hydraulic System description ^^

Three independet hydraulic circuits - called No.1 / green System, No.2 / blue System and No.3 /yellowSystem are continuously operating and supply power to their respective sub Systems äs shown on thepage 29.08.

No. 1 System / green derives primary power from one engine driven pump ( EDP ) mounted on theleft engine . Additional power is provided by an electric driven pump .

No. 3 System / yellow is similar, except that the EDP is mounted on the right engine . The Systemincludes also an electric driven pump .

No. 2 system / blue is powered from a continuously operating electric driven pump .

Each hydraulic System has an air pressurized reservoir .

In case of double engine failure,emergency hydraulic power will be provided via a ram air turbine ( RAT )to the blue system .


HYDRAULIC SYSTEM ^ME£AIRCRAFT

29The primary flight control servos are protected against high flow consumers by priority valves .

High pressure filters are installed in each pump delivery line providing a continuous cleaningof the hydraulic fluid The return and the case drain lines are equipped with Iow pressure filters .

The high pressure lines of all hydraulic circuits are protected against overpressure by relief valves

3- Installation and distribution

In order to meet the segregation requirements the three hydraulic circuits are separated insuch a manner that at least one System remains undamaged at any one failure .This consideration includes all components and the pipe routing located in the areasexposed to damage i.e. due to engine desintegration, tyre burst or accumulator burst.

The hydraulic consumers are distributed to the three hydraulic circuits in away to provide a well balanced and efficient System with adequate redundancy .



29Provision is made for the Installation of manifolds in each hydraulic circuit . These manifoldsinclude hydraulic components which help to minimize the number of pipes and unions andprovide easy maintenance.

Different materials for the pipes are considered :- Titanium alloy for all high pressure pipes .- Aluminium alloy , corrosion protected , for all Iow pressure pipes such äs return , suctionand drain pipes.

- Stainless steel for all high and Iow pressure pipes installed in the fire exposed areasand for coiled pipes.



4- Indication and control 29

Indication is provided on the ECAM displays and on the overhead panel for the followingParameters:

- Reservoir Iow pressure level- Reservoir quantity- Reservoir Iow air pressure- High System pump pressure- Low System pressure- Fluid temperature

The two engine driven pumps, the electric driven pumps , the deployment of the RATand the EDP-depressurizing valves are controlled by push buttons .The fire shut-off valves are controlled by the engine fire handles .



5- Ground handling 29

Reservoir refilling is provided by means of a hand-pump and refilling selector valve which allowsto select the reservoir to be refilled .

External ground connections for the electric and hydraulic power supply are available .

The three hydraulic reservoirs can be externally air pressurized by means of a separatecentralized connection.


HYDRAULIC SYSTEM MPCAIRCRAFT

MAINSYSTEM MAINACC SYSTEM

ACC

PRIORITYVALVE

29

October 1990 Page: 29.08

IGE AND RAIN PROTECTION MPCAIRCRAFT

30

Contents

2

3

4

5

6

7

8

General

Wing Anti Ice (WAI) System

Nacelle Anti Ice (NAI) System

Ice Detection System

Cockpit Windows

Probe Hearing

Waste Water Drain MastHeating

Rain Removal System

SEPTEMBER 1990 Page 30.01

ICE AND RAIN PROTECTION AIRCRAFT

1. General 30Critical areas of the aircraft are protected against ice by hot air Systems or electrical heating.The ice protection System permits unrestricted aircraft Operation in ice and rain conditions.

1.1 Hot Air Systems- The slats outboard the engines of each wing- The engine air intakes

1.2 Electrical Heating Systems- The fligr^t compartment Windows- The air data probes and sensors- The waste water drain mast

1.3 Rain Removal- The front windshields by wipers- Rain repellent fluid System (Option)


IGE AND RAIN PROTECTION

2. Wing Anti Ice (WAI) System 30

2.1 DescriptionThe slats outboard the engines are protected by means of hot air. The air is taken from the pneumatic Systemand the supply to each wing is controlled by shut-off/pressure reducing valves. A flow limiter (orifice) is fitteddownstream of the valve.A telescope duct feeds the piccolo tube in the slats. The hot air passes through slots and exhausts via holesalong the slat rear skin and discharges through gaps between the slat trailing edge and the fixed wing.

2.2 System OperationThe System will be controlled automatically by the IGE DETECTION CONTROL LOGIC, which is connected to twoice detectors. In case of one ice detector failure indication on the EIS is shown and manual System controlis possible (Pushbuttons on the overhead panel).

3. Nacelle Anti Ice (NAI) SystemDescription:The engine nacelles are anti-iced by hot air. The control valve (ON/OFF) is operated automatically from theIce Detection Control Logic or manually by the use of the pushbuttons on the overhead panel.


ICE AND RAIN PROTECTION • MEß-

4. Ice Detection System 30The ice detection System is a primary System with 'AUTO'-control for WAI and NAI. Therefore two ice detectorsare installed and connected to the Ice Detection Control Logic.In case of one ice detector failure Information will be given on the EIS and System Operation control forWAI and/or NAI is manually possible.

5. Cockpit WindowsThe windshields and side Windows are protected against fogging and icing by electrical heating whereasthe windshields are additionally demisted by conditioned air.Control and monitoring units on each side monitor, control and regulate the System. On each side temperaturecontrol of front and lateral Windows are independent.Engine starting automatically initiates the System functioning.

6. Probe HeatingAir data probes and sensors (TAT) are electrically heated.In order to allow correct Operation of the aircraft after a major electrical power supply failure, one System issupplied from the emergency electrical supply.


IGE AND RAIN PROTECTION AIRCRAFT

307. Waste Water Drain Mast Heating

The lavatory/galley waste water drain masts are electrically heated to prevent ice formation. The heatinglevel is automatically reduced when the main landing gear is extended. Heating is operative äs long äspower is avaiiable.

8. Rain Removal SystemEach front windshield is equipped with a two speed electric wiper. A rain repeilent System can be instaliedoptionally.



Wing Anti Ice

ENGINEBLEEDAIR

ECS1

APUBLEEDAIR

Control Volve

Restrictor

'Crossfeed-valve

11

ECS2

Supplyduct

Telescopic DuctFlexible InterconnectionPiccolo Tube

30



Nacelle Anti Ice and Ice Detection 30

WINQ

Engine 1

Ice Detector 1

ANTI IGE

ENG.2

ICE DETECTION

CONTROL LOGIC

l

l

Electr.Power l

TOEIS

Antiiced area

Engine 2

Valve, ON/OFF

Valve position indicating

Ice detector Signal ICE

Hot Airsupply duct

Valve positioncontrol

/ce Detector 2


INDICATING / RECORDING MPCAIRCRAFT

Contents

1 - Cockpit panels

2 - Electronic instrument System

3 - Standby instrument

4 - Recording

I

31

INDICATING AND RECORDING

31

1 - Cockpit panels

iSeptember 1990 Page: 31.01

INDICATING AND RECORDING MPCAIRCRAFT

Main Instrument panelMaster warning/caution l

Flight control unit

EFIS control panel

t~~Side panel

EFISdisplays-

September 1990

Standby

Instrument

panel

ECAM displays Landinggear

L—and brake

panel

31

STANDBYINSTRUMENTDISPLAY

Side panel •

EFISdisplays

Page: 31.02


Pedestal

In addition to the thrust levers and engine controlfunctions the main features are:

- Pitch trim wheels- Flap/slat control lever- Airbrake control lever- Rudder trim panel- Multipurpose Control and Display Units (MCDU)for flight management, maintenance.data link, etc.

- Radio Management Panels (RMP) for communica-tion and navigation control function

- Audio Control Panel (ACP) for intercommunication,transmitting/receiving function

- Lighting Control Panel- Gravity landing gear control handle for emergencyfunction

SEPTEMBER 1990

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Page:31.03


Overhead panel

The overhead panel has a

three-row arrangement with

the main engine related

Systems in the centre line.

The Iower section is for

frequently used Systems,

where the upper section is

reserved to handle any ab-normal situations.

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SEPTEMBER 1990 Page: 31.04

INDICATING AND RECORDING

EIS components

Six identical flat panel Display Units (DU)Size 7.25" x 7.25", füll colour, high resolutionIntegrated graphic generatorInternal source switching (DMP)

Three Display Management Processors (DMP)Digital High Speed (HS) data link to display unitsNo.3 DMP may replace either No.1 or No. 2

- Flight Warning processing integrated in the DMP'sFully redundant warning/caution generation

Two Warning Acquisition Concentrators (WAG)Fully redundant System data acquisition

2 - Electronic Instrument System (EIS) 31



EIS general arrangement

EFIScontrol panel

NavigationDisplay

Primary FlightDisplay ~~

SEPTEMBER 1990ECAM Transfercontrol panel switching panel

31

EFIScontrol panel

NavigationDisplay

Primary FlightDisplay

Page:31.06


Electronic Instrument and Flight Warning System

j_

WXR

SEPTEMBER 1990

fEFIS l FCU EFIS 2l

31

ECAMCONTROLPANEL

WXR

Page:31.07

INDICATING AND RECORDING ® MIESAIRCRAFT

3 - Standby Instrument 31

Air Data

Modern design with smart probes, i.e. transducer and electronic modules are integrated in one airdata module

and assembled directly at the aerodynamically compensated probes. Through serveral wholes Ptot, Ptat and

AOA can be measured. With an input from a TAT probe the module calculates TAS, MACH and ALT.

Advantage: - No more tubing- TAS indication, also MACH is possible- Indication in a symbology similar to PFD



StandbyLCD-Display

SEPTEMBER 1990

31

Page:31.09

rINDICATING AND RECORDING MPC

AIRCRAFT

Standby instruments

RADIAL LOC/GS BEARING

GCO

3S?

u.Q

31

MAGN.NORTH MAGN.NORTH TAS.MACHSTANGE ATTITÜDE ATTITÜDE ALTTTÜDE

W

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< 5

PROBEHEATER

SEPTEMBER 1990

BATTERY POWER BUS 24VDC

Page:31.10

INDICATING AND RECORDINGMPCAIRCRAFT

4 - Recording System 31Mandatory recorder System to fullfill all applicableairworthiness authorities requirements:- ICAOAnnexö- CAASpecNo.lOA- EUROCAEED-55

Use of basic available aircraft avionic System datacollection and validation capacities.

Equipment

- Flight Data Interface Unit (FDIU)

- Linear Accelerometer (LA)

- Digital Flight Data Recorder (DFDR)(optional äs Solid State Flight Data Recorder)

- optional Quick Access Recorder (QAR) with samedata set äs DFDR or free programmable

DATASOURCES

FDIU

\

QAR

September 1990 Page:31.11

LANDING GEAR MPCAIRCRAFT

32

Contents

General

Landing gear Operation

Braking System

Steering System

Landing gear arrangement

Main landing gear

Nose landing gear

September 1990

LANDING GEAR AIRCRAFT

General 32

The MPC 75 landing gear is of the conventional retractable tricycle type with direct action shock absorbers. The mainlanding gear is wingmounted and retracts sideways into the fuselage. The nose landing gear retracts forward into thefuselage.

All tires shall be of the radial type. The main gear tire size is H40x14-R19 and the nose gear tire size is 24x7.7-R10.Bias tires shall be also available in the same size äs an Option.

Landing gear Operation

The nose and main landing gears are operated by hydraulic actuating jacks. The landing gears are mechanically lockedin the fully extended and retracted positions.The sequence of the gears is electrically controlled by duplicated Systems incorporated in a Landing Gear Control andInterface Unit (LGCIU) which utilises proximity Switches to detect the various gear positions. An emergency extensionSystem shall be installed.There are conventional lighted annunciators on the centre Instrument panel and a second display on ECAM to providevisual monitoring of the gear Operation and position.


LANDING GEAR

Braking System 32

The braking System incorporates four braking modes, besides an antiskid and an automatic braking System (autobrake). The main landing gears are equipped with carbon brakes. Steel brakes shall be optional available. The braketemperature is indicated on ECAM. Provisions are provided for Installation of optional brake cooling fans.

The normal braking mode is supplied by the green hydraulic System. The controi computation is done in a fully digitalBrake and Steering Controi Unit (BSCU) which Signals a servovalve for each wheel. For redundancy the BSCU hastwo identical channels with two separate electrical power supplies.

The alternate braking mode is supplied by the yellow hydraulic System. The pressure is controlled by a hydraulicmetering valve and antiskid function is provided through one alternate servovalve for each wheel (signalled by BSCU).The autobrake System is not available in this mode.

The alternate braking without antiskid mode is similar to the alternate mode with the exception that the antiskid functionis not available.

The parking brake mode is activated by an electrical switch. Switching on the parking brake deactivates all otherbraking modes. Hydraulic pressure is supplied from the yellow System or from the brake accumulator.



Steering System 32The nose wheel steering is a hydraulic servo System, electrically controlled from the flight deck via the BSCU. Hydraulicpressure is provided by the green System.In case of loss of hydraulic pressure steering control can be achieved by differential braking or by differential enginethrust. Steering is controlled by two handwheels (±75deg.) or by the rudder pedals (±5deg.).An infernal cam mechanism returns the wheels to the center position (during shock absorber extension) after take-off.If a rollout guidance System is needed, the BSCU can be signalled to provide automatic steering.

Landing gear arrangement

Doors driven bythe gear

Nose gear

Hydraulically driven doors Doors driven by the gear



Main landing gear 32

VIEW X Pintle axis

Side sfay

fuüy extended



Nose landing gear 32

Refracfion actuator

Steering aduator

\Torque links

Tire size 2^7.7-RW

Statte ground line


LIGHTS MPCAIRCRAFT

33

Contents

1 - General

2- Cockpit

3- Cabin

4- Cargo and Service area

5- Exterior

6- Emergency lighting

September 1990

•••••• h l III i ^CH^ . TT"* /""1T"* A T~v-r<AIRCRAFT

l- General 33

- Sufficient illumination is provided to passengers, crew and ground Service.

- Lights System is designed to comply with requirements of JAR relevant items.

-White lighting is used for cockpit illumination. The cockpit lighting is designed to comply with the dark cockpitprinciple.

-The state-of-the-art technology is applied to lights System.

-The lights panel is located on the overhead panel in the cockpit.

2-Cockpit

-Two dorne lights with dimming control are installed to provide cockpit general illumination.

- Side consoles and its briefcase stowage and map holder lights are installed on each side console.

-Floor lights are located beside crew seats.

- Instrument and panel integral lighting with remote stepless dimming control are provided.

- Flood lights provide illumination for instrument panel and pedestal.

-Adjustable reading lights for the captain and first officer are installed in the ceiling. The pedestal floodlight isalso used for thethird occupant.

- Annunciator light dimming and test function are provided.


LIGHTS MPCAIRCRAFT

Lights panel

FUXJO LT INTEO LT FUXD LT•MMPIt. »»m.«PlD

33

September 1990

EXT LTBTROBC BEACON HXNO NAV t LOGOON ON ON OH

TAXIoj

ON

M^RETRACf*

SIGNSSEAT KLTS K] 9WXIN9

ON ON

OMOWTOLTINT LT

Page: 33.02

LIGHTS ® MPCl—• ^ • • • ^^ A TT> rf-«r> A erAIRCRAFT

3-Cabin 33

- Four Strips of fluorescent tubes with dimming control are used for passenger compartment general Illumination.

- Two fluorescent tubes are installed in each entrance area ceiling.

- Call lights are prepared for passengers and crew.

- Fluorescent light is provided in each lavatory to illuminate the wash basin and mirror.

-A reading light is provided for every passenger seat onthePSUs.

- Lighted NO SMOKING, FASTEN SEAT BELT and EXIT signs are provided in the passenger cabin.A RETURN TO SEAT sign is provided in each lavatory. The TOILET OCCUPIED sign is located close tolavatories on the cabin visible wall.

- Special galley work lights form part of the fixed BFE galley equipment. The Illumination is provided at theattendant's seat.

- Passenger stairway lights are integrated in the stairway.

4-Cargo and Service area-A separate lighting System is installed in cargo hold. The lights are controlled by each cargo door.

- Loading area light is sufficient to permit reading of labels on the ground in front of the cargo door.

- Service lights and electrical outlets are provided in all Service bays.


LIGHTS MPCAIRCRAFT

CABIN GENERAL FLUORESCENT LIGHTING

PASSENGERREADINGLIGHTS

PASSENGERREADINGLIGHTS

AISLE EMERGENCY LIGHTING

~7 ' ','',//'//'//////

33


LIGHTS|_IV2I • • VJ AIRCRAFT

5-Exterior 33-Twoforward facing navigation Hghts are installed on both wing tips and a rearward facing navigation light is

installed in the tan cone.-Two retractable landing lights are installed on the fixed underwing panel.

-Two fixed taxi lights are installed on the nose landing gear to provide wide beam illumination for taxiing.

-Two anti-collision/strobe lights are installed on top and bottom of the fuselage along the centre line.

-Two wing and engine scan lights on each side of the fuselage are provided to illuminate the wing leading edgeand the engine intake.

- Logo lights are provided on both sides of the horizontal tail äs an Option to illuminate the airline's logo on bothsides of the vertical tail whenever the landing gear is extended.

-Strobe lights at wing tips and tail cone äs an Option are used äs a duplicated navigation lighting System.

6-Emergency lighting-Some of the cockpit lights are used for emergency lighting powered by essential busbars.

-Cabin emergency lights are provided to illuminate the aisle area.

- A dome light is provided in each lavatory powered from essential busbar.

- Evacuation path lights are installed in the stairway and slides.

- Floor proximity escape path marking consists of pathway lights and EXIT signs above floor.

-The Emergency Power Supply Units equipped with internal Ni/Cd batteries will provide sufficient power to allemergency lights and signs for at least 12 minutes.


LIGHTS MPCAIRCRAFT

ANTI-COLLISION LIGHT

33

NAVIGATION LIGHTSTROBE LIGHT(OPTION)

LOADING AREA LIGHT

LOGO LIGHT(OPTION)WING AND ENGINE SCAN LIGHT

LANDING LIGHT


NAVIGATION MPCAIRCRAFT

34

Contents

- General

- Radio navigation aids

- Air data / attitude / heading

NAVIGATION MPCAIRCRAFT

Air data and attitude SystemH WXR hf 1

-« .- r r n WARN «--, —i——*—3rdPitot.StaticAOA

TAS, ALTMach

CPT'sPitot.StaticAOA,TAS, ALTMach

TAT 1 =L

34

September 1990

f] F/O'S/ Pitot «Statics AOA,/ TAS, ALT

Mach

TAT 2

DATA BUSSESTO USER SYSTEMS

Page: 34.01

NAVIGATION ^ MEß

2 - Radio Navigation Aids 34

The aircraft is equipped with conventional navigation aids. They are automatic tuned through the Flight ManagementSystem (FMS), for DME, VOR and ADF, or manual via the Radio Management Panels (RMP).

ILS and MLS will be integrated in one LRU if ever MLS becomes operational.Global Positioning System( GPS) is a Standard Option, and will be integrated in the AHRS or IRS.The Traffic Collision Alert and Avoidance System (TCAS) is an optional Installation (for US operator mandantory1993)

3 - Air Data and Attitüde System

The aircraft offers a triple Installation of air data Systems and Attitüde and Heading Reference System (AHRS).All installed air data probes are smart probes, which are aerodynamically compensated. They supply theSystem with pitot, static and alpha air data. The data are converted to ARINC 429 format in the Air Data Modules(ADM) that are mounted directly to the probes. With an input of Total Air Temperature (TAT) the ADMscalculate the miscellaneous air data TAS, ALT and Mach.

Basically the aircraft is equipped with conventional strapdown AHRS updated by two manetic sense units.The installation of an optical gyro Inertial Reference System (IRS) is provided äs a Standard Option.Both sytems are of the same size (4 MCU) an can be upgraded by an optional GPS Integration.


Navigation MPCAIRCRAFT

Radio Navigation

To DMPs

FCGC 1September 1990

FCGC 2Page: 34.03

OXYGEN SYSTEM MPCAIRCRAFT

35

Contents

1 - General

2 - Schematic

September 1990

OXYGEN SYSTEM

351 - General Description

Flight Crew Oxygen

The crew gaseous oxygen System is schematically shown in figure 35-1 . Each crew Station has a quick-donning mask with a demand regulator installed. The oxygen is supplied from a high pressure oxygencylinder to masks through a pressure regulator / transmitter assembly and distribution circuit.

The crew oxygen System is capable of supplying oxygen whenever the cabin pressurization fails at anyaltitude

The high-pressure portable oxygen cylinder assembly is used by crew members in case of emergencywhen fighting a fire , against the emission of smoke and noxious gas with füll mobility if there is a failureofthefixed crew oxygen System.

Passenger Oxygen Unit

The solid state chemical oxygen dispensing units are mounted above the passenger seats and in lava-tories (fig.35-2). Each mask isprovided with continuous flow of oxygen lasting at least 12 minutesafterflow Initiation.

Attendant Oxygen SystemAbove the attendants Station a two-mask chemical oxygen unit is installed to supply emergency oxygenin case of cabin decompression.Each attendant is equipped with a portable oxygen unit which can be used to provide emergency oxygenand first-aid treatment for passengers.

September 1 990 Pa9e: 35-01


2 - SchematicFLIGHT CREW MASK WITH

HP/LPOVERPRESSURE SAFETY

D&CHARGE INOICATOR

ATORS

H_FIRST

OFFICER'S

MASK

OXYGEN

INDICATINGANO

SW ITCHING

LP INOICATOR

,

p

\L

3RD LOCCUPANT'S CAPTAIN'S

MASK MASK._-!

HP INOICATOR

ri SUPPLY MANIFOLO fcr-»„

OXYGEN

STORAGE CYLINOER

LP TEST POINT

SOLENOID SUPPLY VALVE

PRESSURE CONTROL ASSEMBLY

(PRESSURE REOUCER PRESSURE TRANSMITTER)

'PRESSURE GAUGE

Figure 35-1 Crew Gaseous Oxygen System Schematic

35

September 1990 Page: 35.0Z


35

— key— 2mask units(15)3mask units(20)4mask units(5)

September 1990

Flgure35-2/1 Passenger Oxygen System Distribution

Page ;:35.03/1

OXYGEM SYSTEM MPCAIRCRAFT

- KEY - * = 2MASK UNITSC22)

O = 3MASK UNITS(24)

A = 4MASK UNITS(6)

35

September 1990

FJgure35-2/2 Crew Gaseous Oxygen System Schematic

Page: 35.03/2

PNEUMATIC SYSTEM MPCAIRCRAFT

36

Contents

- General

- Bleed air System

- Engine bleed air supply

- Additional sources of bleed air

September 1990

PNEUMATIC SYSTEM AIRCRAFT

General description 36

Bleed air System

The pneumatic (bleed air) System provides hot compressed air for the followingSystems / functions:

- Air conditioning- De-icing- Engine starting- Hydraulic and water Systems pressurisation

Engine bleed air supply (see figure 36.10)

Each engine has its own bleed air supply control System, normally isolatedfrom each other by the crössfeed valve.Bleed air is drawn from either an intermediate pressure (IP) stage or a highpressure (HP) stage of the engine compressor, depending on flight mode(power setting). In the climb and cruise modes, IP bleed air pressure is high enoughto satisfy system requirements and the HP control valve remains closed.

September 1990 Page : 36.01

PNEUMATIC SYSTEM

In descent mode ( Iow engine power setting ) the l P bleed air pressure is

inadequate for system requirements. The drop in pressure allows the HP

control valve to open. HP stage bleed air then supplies the bleed system,

while isolating the l P stage by ciosing the l P stage check valve.

IP or HP bleed air is consequently passed through the pressure regulator,

overpressure valve and finally the precooler heat exchanger to provide

user Systems with pressure and temperature controlled bleed air.

Additional sources of bleed air are :

- Ground supply unit - via HP ground connector

-TheAPU

Under normal flight conditions the bleed air system functions fully

automatically. A dedicated bleed air monitoring Computer (BMC)

monitors each engine bleed system, and provides indication for crew

action when system faults occur.

September 1990 Page : 36.02

PNEUMATIC SYSTEM MPCAIRCRAFT

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FIGURE 36.1ft BLEED AIR SUPPLY-SCHE.MATIC

September 1990

2. PRESSURE REGULATING AND SHUT OFF VALVE

3- OVERPRESSURE VALVE

4. FAN AIR CONTROL VALVE

5. HEAT EXCHANGER (PRECOOLER)

6. CROSSFEED VALVE

Page : 36.03

WATER / WASTE MPCAIRCRAFT

38r

i

Contents

1

2

3

4

5

- General

- Potable Water

- Waste Water

- Waste Disposal

- Air Supply



1. GeneralThe aircraft is equipped with water and waste facilities. The System supplies potablewater for the galleys and lavatories, and the waste disposal Service allows wastewater from the galley and teilet wash basins to drain from the aircraft, via drainmastsfitted at the underside of the fuselage.The teilet waste contents are collected in tanks directiy connected to the teilet flushSystem.

38

2 . Potable WaterThe potable water is stored in a 60 liter (15.5 USgal) reservoir which is installed inthe pressurized underfloor area forward of the centre wing box.The tank is manufactured from glass fibre composite and pressurized from thebleed air System.


WATER / WASTE

383 . Waste WaterWaste water from the basins in the lavatories and galleys is drained out of theaircraft via two electrical heated drain masts in the forward and aft fuselage.Additional provisions are provided for an optional lavatory and galley.

4. Waste DisposalThe toilet System collects the waste disposal in the waste tanksof the toilet units. The waste tanks are manufactured from glass fibrecompound and are equipped with a filter/pump and a drain valve assembly.During ground Service the tanks are emptied, cleaned and filled with aprescribed quantity of flush-fluid.A Ventilation System for the toilet units prevents odours from coming out ofthe units.

5. Air SupplyThe potable water supply tank is pressurized with pressure regulated andtemperature controlled air from the bleed air System. The air supply linesare equipped with non-return valves to protect against reverse fluid or airflow induction.

September 1990 Pagef 38.03


Fig. 38-01

September 1990

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S 22553 53C ä Pa|e: 38.04

ON-BOARD MAINTENANCE SYSTEM MPCAIRCRAFT

Contents

- General

- Centralized maintenance

- Aircraft condition monitoring

45

ON-BOARD MAINTENANCE SYSTEM

General 45

The on-line maintenance of the electronic Systems is based on the use of the CMS(Centralized Maintenance System), comprising of one CMC (Centralized Maintenance Computer)and an optional printer.

The QMS provides for line- and shop-maintenance a central means to:

Display and interrogate the BITEs of the various Systems

andInitiale tests from the MCDUs (Multipurpose Control Display Units)

located on the centre pedestal (cockpit).

The intelligence required for:

detecting the failures

processing the corresponding maintenance data

formatting messages to be displayed on the MCDUs

is included in each BITE of the avionics Systems.

September 1990 PagöyJ5.01

ON-BOARD MAINTENANCE SYSTEM MPCAIRCRAFT

Electronic Library System (ELS) 45

InputInterface

Dataprocessing

OutputInterface

Storagedevice

5.25" Optical Disc (AMM,MEL,IPC,NAV...)


XIXILIARY POWER UNIT MPCAIRCRAFT

49

Contents

2

3

4

5

6

7

8

General

APU operating envelope

APU Installation

APU Subsystems

APU compartment cooling,Ventilation and drain System

Air intake

Exhaust

Control and monitoring

11990 Page :49.01

AUXILIARY POWER UNIT

1 General 49

The Auxiliary Power Unit (APU) instalied on the MPC 75 is designed to provide bleed air to the aircraftpneumatic System and to drive an oil spray cooled AC generator during ground and in flight Operation,thus increasing aircraft operational flexibility.

Independence of the MPC 75 from external sources on ground (up to 8000 ft) is assured by power availablefrom the APU to drive a 60 kVA generator and additionally to supply bleed air for main engine start or theair conditioning System. During flight Operation the APU is capable of providing electrical powerwithin the whole flight envelope.

This permits dispatch of the aircraft under certain MEL conditions, where together with the in flight restartcapability (up to TBD ft) a considerable increase in redundancy of electrical power is obtained.

Maximum Operation limit of the APU is aircraft ceiling altitude.

SEPTEMBER 1990 Page :49.02

APU MPCAIRCRAFT

APU Performance Comparison for MPC 75 (Sea Level Standard Day)MPC 75 electrical requirement 60 kVA

49

GTCP 36-150

(ATR 72)

(BAe 146-200)

(Fokker 100)

GTCP 36-250

GTCP 36-300(A320)

APIC APS2000(B 737-500)

KHDT218

150(AT) (16 kW)

150M (18.6kW)

150R(35.1 kW)

MPC 75 Bleed Air Requirements

2

(45kW)

(90 kW)

(60kW) (45 kW)

(60 kW)

0.1 0.2 0.3 0.4 0.5 0.6

Bleedairflow

0.7 0.8 0.9

[kg/s]SEPTEMBER 1990

Flight O

peration

O

(O8CD

X30T3Om

AUXILIARY POWER UNIT AIRCRAFT

3 APU Installation 49

The APU is located in the aft fuselage in the tail cone.The housing is arranged principally äs a support and aerodynamic fairing for the APU, air intake Systemand exhaust System.

In accordance with the construction regulations, the APU compartment is arranged äs a fireproof box and isrigidly attached to the fairing structure. The firewalls and the access doors are manufactured from titanium sheet.

There is a pressure relief door to limit the pressure in the APU compartment to a value acceptable for thestructure.

The Installation of the APU is designed for a rapid change, the time for 'quick engine change unit' removaland Installation, ready to start, will not exceed 1 h 15 min.

The APU support is designed according to fail safe principles covering the event of fire.

SEPTEMBER 1990 . Page: 49.06


4 APU Subsystems 49Fuel supply and controlThe APU is supplied with fuel through a separate fuel line, drawing fuel from the LH main engine fuel System.A dedicated APU pump is installed in that line to permit fuel supply to the APU when the engine fuel System is notoperating. Opening of the cross feed valve permits fuel supply from the RH tank to the APU.

An LP valve installed on the tank boundary permits to isolate the APU fuel line when the APU is cut off, thusavoiding pressure on the fuel line across the pressurised zone.

The LP valve is controlled by the APU master switch and closes automatically in case of emergency shut downof the APU.

The fuel control System operates fully automatically controlled and monitored by the APU Control Unit (ACU).

The System provides metered fuel to the APU combustor according to the inputs received from the ACU fortimed acceleration and constant speed control (+ 0.5 per cent N).


AUXILIARY POWER UNIT MPCAIRCRAFT

Bleed air System

The bleed air flow is controlled according to aircraft System demand.

Sufficient APU bleed air power is available, while providing electrical power for equipment required for groundOperation, to meet the specified ground air conditioning requirement.

The System is protected against reverse bleed air flow.

A compressor surge control System prevents compressor surge for all operating modes. The surge air isconducted into the APU exhaust.

49

Lubrication and generator cooling

The integral lubrication System for engine, generator and gear box includes an internal gear box oil reservoir,an oil pump assembly and an oil cooler.

By design the oil cooler capacity is sufficient to permit cooling of the oil spray cooled generator with the samecircuit.

The APU operates with the same oil types äs approved for the main engines and their accessories.



4Q5 APU compartment cooling/ventilation and drain System

The APU cooling air and Ventilation System provides air for cooling of the APU, equipment on the APU andVentilation of the APU compartment. Cooling air supply is the APU plenum chamber. Cooling air is deliveredto the oil cooler and the APU compartment. Air out of the oil cooler is led to the outside of the aircraft.

After APU shut down Ventilation of the APU compartment is achieved by a louvre in the upper side of theAPU compartment. Free convection provides efficient Ventilation of the APU compartment.

Drainage and leaks from various places are collected in a drain tank and discharged outboard via a drain mast.

SEPTEMBER1990 Page:49.12


6 Air intake

The common air intake for power section, bleed air generation and cooling is designed äs a silencer providinghigh intake pressure recovery.

The air intake is located considering

noise restrictionsicing conditionsdust ingestionre-ingestion of APU exhaust gaseshigh Installation pressure ratio in flight

49

7 Exhaust

The exhaust System is installed in a fire proof compartment separated from the APU compartment.The exhaust System functions äs a silencer. It conducts the not APU exhaust gases to the outside of theaircraft at a location providing high Installation pressure recovery in flight.



8 Control and monitoring 49

The central component of the control and monitoring System is the APU Control Unit (ACU). This unit isinstalled in the rear fuselage and is easily accessible.

It uses microprocessor/digital techniques to realise the following functions :

control the Start sequenceregulate the APU speed

- regulate the bleed air flowmonitor all important Operation parametersmonitor the sensor units on the APU to detect the development of a serious faultdrive indications and annunciation to the cockpit and provide adequate Signals to theECAM, FWS, QMSinitiale APU shut down in case of a fault

- control the APU surge System

An RS 232 Interface is provided on the ACU. This allows check out of the ACU circuits in the shop, äs well äsadjustment of the performance Outputs.

The controls to switch on or off the APU and to activate the APU starting sequence are located in the cockpit.The APU Operation is fully automatically controlled by the ACU.

APU monitoring for the flight crew is provided by indications on ECAM.

SEPTEMBER1990 Page:49.16

AUXILIARY POWER UNIT AIRCRAFT

Control and monitoring cont'd 49

For maintenance purposes APU Parameters are provided to the Onboard Maintenance System (OMS).A test button on the cockpit fire panel allows testing of the fire extinguisher circuit äs well äs functional integrityof the fire detection System

Besides the controls in the flight deck emergency shut down of the APU is possible :

- by pushing the APU fire handle- from the nose landing gear APU emergency shut down push-button- automatically on ground after an APU fire is detected

SEPTEMBER 1990 Pa9e: 49'16a

ENGINE CONTROLS MPCAIRCRAFT

73

Contents

General (FADEC)

Engine Control Unit

Engine Interface Vibration Monitoring Unit

September 1990

ENGINE CONTROLS AIRCRAFT

FÜLL AUTHORITY DIGITAL ENGINE CONTROL (FADEC)

G e n e r a l D e s c r i p t i o n

The MPC 75 engines are electronically controlled by the Füll Authority Digital Engine Control (FADEC). Each

engine is equipped with a fully redundant (dual) FADEC System which provides better engine protection and

permanent engine health monitoring.

The application of a FADEC provides advantages to the MPC 75 aircraft. It saves weight, reduces pilot workload

and maintenance cost, saves fuel by dimension free control of the gas generator and allows the Optimum

adaptation of thrust schedules to the aircraft needs.

FADEC is an electronic System which consists of a fully redundant (dual) Engine Control Unit (ECU) and an

(simplex) Engine Interface Vibration Monitoring Unit (EIVMU)) with built in failure .detection. In case of a FADEC,

malfunction, the System Switches to the other channel or an alternate control mode, the best working channel always

being in Charge of engine controls. This ensures engine performance and safety.



FADEC cont. 73

Each engine is equipped with a FADEC which provides:

- gas generator control

- engine limit protection

- engine automatic starting

- fly-by-wire System compatibility

- power management

- failure Identification and indication

- condition monitoring

September1990 ; Page-.73.02

ENGINE CONTROLS

Engine Control Unit (ECU) 73One ECU is located on the engine with dual redundant channels (active and standby) each having separate

aircraft power sources to ensure engine starting on the ground and in flight. In addition dedicated ECU

alternators assure independent power supply.

Dual redundancy for electrical input devices ( Throttle Lever Angle (TLAs), engine sensor Signals, Air Data System

(ADS 1 +2) ).

Dual redundancy for electrical part of control actuator.

The Simplex hydromechanical unit includes:

- the metering valve and servo valves to control the variable Stator position

- the variable bleed valve control System

- the HP turbine active clearance control System actuators

- the fuel System

- the ignition System

- the thrust reverser System

Fault tolerance and fail operational capability.

High level of protection against electromagnetic disturbance.September1990 _ Page: 73.03


Engine Interface Vibration Monitoring Unit (EIVMU)

The Interface between the FADEC System and the other aircraft Systems is mainly performed by the EIVMU

through digital data buses.

One EIVMU per engine is located in the avionics bay.

Gare is taken to preserve Systems segregation for safety and integrity.

The EIVMU includes the necessary data processing Provision to make engine parameters available for a condition

monitoring System.

The compatibility with the aircraft fly-by-wire System and incorporation of power management function allows

significant aircraft weight saving and fuel burn reduction.

73

Septembof1990 Page: 73.04

ENGINE CONTROLS MPCAIRCRAFT

Systemsisplay

Engine / Warningsdisplay

Throttle lever angle

ADS 1 +2

ADS1+2 [

Throttle lever angle

Automatic/Manual starting

FMS

FCGC

ENGINEINTERFACEVIBRATION

MONITORINGUNIT

(EIVMU)

EngineControl

Unit(ECU)

CHANNEL A

EngineControl

Unlt(ECU)

CHANNEL* B

Engine sensor Signals

Ground / Flight

Bleed Status

iHydromechanicalfuel unit

Start valve

Thrust ReverserSystem

Ignition System

Engine sensor Signals

73

September 1990Page: 73.05

Documents

MPC 75 System Definition MPC AIRCRAFT A AC ACARS ACMS ACP ACU ADF ADM ADS AFT Ah AHRS ALT AMM AOA APU ATC B BAT BFE BITE BMC BRT BSCU CAA GAB CAPT CAUT CDU CFDS CIDS CMC CMS COND CPM