Phenom100 MTM - Vol 3

TM
Embraer
Empresa Brasileira de Aeronáutica S.A.
PHONE: (55 12) 3927-7517
FAX: (55 12) 39277546
http://www.embraer.com.br
e-mail: training@embraer.com.br
MAINTENANCE
TRAINING
MANUAL
VOL. 3 OF 4
Copyright 2007 by Embraer - Empresa Brasileira de Aeronáutica S.A. All rights reserved.
This document shall not be copied or reproduced in whole or in part, in any form or by any means without the express written
Authorization of Embraer. The information, technical data, designs and drawings disclosed in this document are proprietary
information of Embraer or third parties and shall not be used or disclosed to any third party without permission of Embraer.
MAINTENANCE TRAINING MANUAL VOL. 3 TM
Embraer proprietary — Copyright © — 2008

Developed for Training Purposes Only

THIS PAGE INTENTIONALLY LEFT BLANK
FRONT MATTER - page 6

VOLUME 3
21 ...........................................................AIR CONDITIONING 55 .....................................................................STABILIZERS

22 ....................................................................AUTO FLIGHT 56 ..........................................................................WINDOWS

23 ...........................................................COMMUNICATIONS 57 ................................................................................WINGS
24 .......................................................ELECTRICAL POWER 71 ...................................................................POWERPLANT
25 .............................................EQUIPMENT/FURNISHINGS 72 ...............................................................................ENGINE
26 ............................................................FIRE PROTECTION 73 .........................................ENGINE FUEL AND CONTROL
27 ..........................................................FLIGHT CONTROLS 74 .............................................................................IGNITION
28 ...................................................................................FUEL 75 ......................................................................................AIR
29 .........................................................HYDRAULIC POWER 76 ..........................................................ENGINE CONTROLS
30 ...........................................ICE AND RAIN PROTECTION 77 .........................................................ENGINE INDICATING
31 ..............................INDICATING/RECORDING SYSTEMS 78 ...........................................................................EXHAUST
32 .................................................................LANDING GEAR 79 .......................................................................................OIL
33 ...............................................................................LIGHTS 80 ..........................................................................STARTING
34 ......................................................................NAVIGATION
35 .............................................................................OXYGEN
36 .......................................................................PNEUMATIC
38 .................................................................WATER/WASTE
44 ...............................................................CABIN SYSTEMS
45 ................................CENTRAL MAINTENANCE SYSTEM
50 ...............CARGO AND ACCESSORY COMPARTMENTS
51 ...............................Unknown ATA2200 Chapter number
52 ...............................................................................DOORS
53 .........................................................................FUSELAGE
54 .........................................................NACELLES/PYLONS
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80



CHAPTER 26 - FIRE PROTECTION
SECTION TITLE PAGE

26-00 FIRE PROTECTION 10

26-10 FIRE/SMOKE DETECTION SYSTEM 14
26-11 ENGINE FIRE/OVERHEAT DETECTION SYSTEM 18
26-20 FIRE EXTINGUISHING 28
26-21 ENGINE FIRE EXTINGUISHING SYSTEM 30
26-24 PORTABLE FIRE EXTINGUISHING SYSTEM 40
22-Aug-2008 CHAPTER 26 - page 9
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EFFECTIVITY: ALL
FIRE PROTECTION 26-00
Introduction • One portable fire extinguisher.
The function of the fire protection system is to monitor the aircraft for fire and The control modules that interface with this system are:
overheat conditions, and to permit the discharge of fire extinguishing agent
to eliminate these conditions. • GEA (Garmin Engine/Airframe unit)
General Description • GIA (Garmin Integrated Avionics unit)
The FIRE PROTECTION includes these subsystems: The figure FIRE PROTECTION - BLOCK DIAGRAM provides further data on
the preceding text.

• FIRE/SMOKE DETECTION SYS- (AMM SDS 26-10-00/1)
TEM
• FIRE EXTINGUISHING (AMM SDS 26-20-00/1)
Components
FIRE/SMOKE DETECTION SYSTEM (26-10)
The function of the fire detection system is to give conditions for detection of
fire and overheat in the engines, and alert the crew about these conditions.
FIRE EXTINGUISHING (26-20)
The function of the fire extinguishing system is to discharge fire extinguishing

agent in areas where fire/overheat can occur. This system has fixed and
portable components charged with Halon 1301 agent, which permits the crew
to extinguish the fire.
The fire protection system has these components:
• One fire detector per engine, in a single loop configuration.
• Two engine shutoff pushbuttons, one engine fire extinguishing switch and
one test button.
• One ENG FIRE EXTINGUISHER and one TEST control panels.
• One engine fire extinguishing bottle.

EFFECTIVITY: ALL
HOT BUS 1
INTEGRATED

AVIONICS FIRE/ ENG/ TRIM
TEST PANEL
UNIT 1 PANEL
(GIA 1)
RS485
DETECTOR 1
ENGINE/AIRFRAME
UNIT 1
(GEA 1)
ENGINE FIRE EXTINGUISHING BOTTLE
EM500ENSDS260007A.DGN
FIRE PROTECTION - BLOCK DIAGRAM
Sheet 1
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EFFECTIVITY: ALL
NOTES: NOTES:

EFFECTIVITY: ALL
HOT BUS 1

INTEGRATED
AVIONICS FIRE/ ENG/ TRIM
TEST PANEL
UNIT 2 PANEL
(GIA 2)
RS485
DETECTOR 2
ENGINE/AIRFRAME
UNIT 2
(GEA 2)
ENGINE FIRE EXTINGUISHING BOTTLE
EM500ENSDS260008B.DGN
FIRE PROTECTION - BLOCK DIAGRAM
Sheet 2
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EFFECTIVITY: ALL
FIRE/SMOKE DETECTION SYSTEM 26-10
Introduction
The function of the fire detection system is to give conditions for detection of
fire and overheat in the engines, and alert the crew about these conditions.
General Description
The FIRE/SMOKE DETECTION SYSTEM includes this subsystem:


• ENGINE FIRE/OVERHEAT DE- (AMM SDS 26-11-00/1)
TECTION SYSTEM
The engine fire/overheat detection subsystem has one fire detector in each
engine compartment, installed on the mid cowl compartment with its housing
mounted on the engine structure.
The CAS (Crew Alerting System) window on the PFD (Primary Flight Display)
and the CMC (Central Maintenance Computer) screen show the detector
failure messages. When the engine fire detector senses a fire/overheat
condition, the system alerts the crew by means of the FIRE message in the
respective engine ITT (Interstage Turbine Temperature) field on the EICAS
(Engine Indication Crew Alert System), a voice message, and a red light on
the engine shutoff pushbutton (red light stays on while the fire condition
persists).
Components
ENGINE FIRE/OVERHEAT DETECTION SYSTEM (26-11)
The function of the engine fire detection system is to provide conditions for
fire and overheat detection in both engine compartments, and to alert the
crew about these conditions.
The figure FIRE/SMOKE DETECTION SYSTEM - BLOCK DIAGRAM

provides further data on the preceding text.

EFFECTIVITY: ALL
HOT BUS 1

CAS
MESSAGES
FIRE/ENG/TRIM TEST
GIA 1
PANEL PANEL
CMC
MESSAGES RS485
GEA 1 DETECTOR 1
FIRE/SMOKE DETECTION SYSTEM - BLOCK DIAGRAM
Sheet 1
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EFFECTIVITY: ALL
NOTES: NOTES:

EFFECTIVITY: ALL
HOT BUS 1

CAS
MESSAGES
FIRE/ENG/TRIM TEST
GIA 2
PANEL PANEL
CMC
MESSAGES RS485
GEA 2 DETECTOR 2
FIRE/SMOKE DETECTION SYSTEM - BLOCK DIAGRAM
Sheet 2
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EFFECTIVITY: ALL
ENGINE FIRE/OVERHEAT DETECTION SYSTEM 26-11
Introduction A FIRE button on the TEST control panel is used to make sure of the integrity
of the detection system; when it is pressed, a fire condition on the engines is
The function of the engine fire detection system is to provide conditions for simulated, and the fire alarms are activated (red light in the shutoff pushbutton
fire and overheat detection in both engine compartments, and to alert the lamps, FIRE message in the ITT field on the EICAS and voice message
crew about these conditions. FIRE).
General Description The figure ENGINE FIRE/OVERHEAT DETECTION SYSTEM -
COMPONENTS LOCATION provides further data on the preceding text.
The engine fire detection system has one single loop-type fire detector for

each engine. The fire detector is installed on the mid cowl compartment and
its housing is mounted on the engine structure using a P-clamp and brackets.
The detector sensor tube is installed along the mid cowl compartment, close
to the main flammable fluid components, covering both left and right sides of
the engine. The detector sensor tube is mounted on the engine structure and
system lines using hinged clamps.
The system is able to detect either overheat (average temperature) or fire

(discrete air temperature). When the engine fire detector senses a fire/
overheat condition for an engine, a signal is sent to the GEA (Garmin Engine/
Airframe unit) and to the engine shutoff pushbutton in the ENG FIRE
EXTINGUISHER control panel.
Each engine fire detector is electrically connected to supply:
– Warning indication by means of a red light on the engine shutoff

pushbutton.
– Warning indication by means of a red FIRE message in the ITT (Interstage

Turbine Temperature) field on the EICAS (Engine Indication Crew Alert
System).
– Voice message FIRE.
The engine fire detection system is connected to the aircraft electrical buses
as follows:
– Engine 1(2) - Hot Bus 1.

EFFECTIVITY: ALL
A B
87.8 TO 87.8
ATR

C
OFF TEST
ANNUNCIATOR
77.5 N1%
FIRE FIRE
ITT FIELD
IGN IGN
__
544 ITT C
__
99.9 N2% STALL PROT
OIL PRES PSI
0 TEMP C 0
ENG FIRE EXTINGUISHER FUEL
FF KGH
SHUTOFF 1 BOTTLE SHUTOFF 2 TEST CONTROL PANEL
FQ KG
DISCH
C
OFF LWD
EICAS
EM500ENSDS260002D.DGN
ENG START/STOP
RUN RUN
STOP START STOP START B
ENG FIRE EXTINGUISHER CONTROL PANEL
ENGINE FIRE/OVERHEAT DETECTION SYSTEM - COMPONENTS LOCATION
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
NOTES: NOTES:

EFFECTIVITY: ALL
ENGINE FIRE
DETECTOR
B

DETECTOR
SENSING
ELEMENT
DETECTOR
A A
C
CLAMP
DETECTOR
C SENSING
ELEMENT
CLAMP
DETECTOR
SENSING
ELEMENT
B
TYPICAL
C
TYPICAL
ENGINE FIRE/OVERHEAT DETECTION SYSTEM - COMPONENTS LOCATION
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
Components ENGINE FIRE/OVERHEAT DETECTION SYSTEM - FIRE DETECTOR SET

POINT TABLE (Continued)
FIRE DETECTOR
DISCRETE TEMPER- AVERAGE TEMPER-
FIRE DETECTORS
The engine fire protection system LRU (Line Replaceable Unit) model 801- ATURE ATURE
DRH from Meggitt Safety Systems, is of the pneumatic fire/overheat type.
Engine Fire Detector 593 °C (1099 °F) 251 °C (484 °F)
The detector is an electromechanical device factory calibrated, hermetically
sealed, thermally sensitive, and is pneumatically actuated. Its primary NOTE: DISCRETE TEMPERATURE - High-intensity fire impingement on a

components are: small localized section.
AVERAGE TEMPERATURE - General overheat of the entire sensor
– Pneumatic sensor element. tube area.
– Responder assembly. The figure ENGINE FIRE/OVERHEAT DETECTION SYSTEM - WORK
PRINCIPLE provides further data on the preceding text.
The pneumatic sensing element consists of a rugged 0.229 cm (0.090 in.)
outside diameter stainless steel tube charged with helium (inert gas), for
compartment general overheat detection. It also contains hydrogen (active
gas) as a charged core material, for extremely localized heat detection.
The responder assembly has two pressure-sensitive switches (alarm and

integrity), an electrical connector assembly, and a housing assembly.
Each pressure switch has a preformed metal diaphragm and a stationary

contact pin on one side of the diaphragm that is also electrically isolated with
ceramic. The diaphragm is the moving contact of a normally-open single-pole
throw switch. It moves over the center, tightly contacting the stationary
contact pin. With the diaphragm against the center contact pin, there is a path
for the current flow. When the pressure against the diaphragm decreases to
below the activation force, the diaphragm moves back over the center away
from the stationary contact pin and opens the electrical path.
The housing assembly has a protective shell with all responder assembly
components inside. It has an electrical connector assembly on one end and
the sensor element on the other end.
The table FIRE DETECTOR SET POINT presents the set point of fire
detectors.

EFFECTIVITY: ALL
RESPONDER HOUSING
A
E F
B
D C

RESPONDER ALARM
SWITCH (N.O.)
DETECTOR HOUSING
ISOLATOR
A
+28VDC SENSOR
B
C
NO D
TEST
ALARM E
F
SPARE INTEGRITY SWITCH
(HELD CLOSED BY NORMAL
SENSOR PRESSURE)
INTERFACE WIRING
SHUNT PLATE
SCHEMATIC SENSOR/RESPONDER−TYP
ENGINE FIRE/OVERHEAT DETECTION SYSTEM - WORK PRINCIPLE
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EFFECTIVITY: ALL
Operation The fire detection sensor for RH en-

ENG 2 FIRE DETECTOR FAIL gine is unable to detect fire/overheat
If fire is detected: condition.
• The fire detector sends a signal to the GEA that communicates with GIA
(Garmin Integrated Avionics unit) by means of an RS-485 digital interface. The figure ENGINE FIRE/OVERHEAT DETECTION SYSTEM -
Refer to AMM SDS 31-41-00/1 for more details. SCHEMATIC DIAGRAM provides further data on the preceding text.
• The fire detector also sends a signal to the control panel to cause the

respective engine shutoff pushbutton to come on.
• The GIA provides a FIRE inscription in the ITT field on the EICAS and a
voice message FIRE.
The engine shutoff pushbutton stays on as long as the fire condition persists.
A single loop fire detector is installed in each engine and its integrity is
continuously monitored; in case of failure of power supply, bottle pressure,
cartridges and associated harnesses, fail messages come into view on the
CAS (Crew Alerting System) window and on the CMC (Central Maintenance
Computer) screen on the PFD (Primary Flight Display).
ENGINE FIRE/OVERHEAT DETECTION SYSTEM - CAS MESSAGES

(Continued)
INDICA- LEVEL
DESCRIPTION
TION (COLOR)
The fire detection sensor for LH (Left-Hand)
E1 FIRE Caution
engine is unable to detect fire/overheat con-
DET FAIL (Amber)
dition.
The fire detection sensor for RH (Right-
E2 FIRE Caution
Hand) engine is unable to detect fire/over-
DET FAIL (Amber)
heat condition.
The fire detection sensor for LH en-

ENG 1 FIRE DETECTOR FAIL gine is unable to detect fire/overheat
condition.

EFFECTIVITY: ALL
FIRE/ENG/TRIM PANEL TEST PANEL
WHITE (PRESS)
RED (FIRE)
NOT SHUTOFF
HOT BUS 1
SHUTOFF

ENG 1 SHUTOFF
E1 FIREX
INTEGRATED AVIONICS UNIT 1

(GIA 1)
ENG 1 FIRE EXT PWR

INPUT
ENGINE/AIRFRAME UNIT 1
(GEA 1)
FIRE
SW
ENG 1 FIRE DETECTOR

ENG 1 FIRE
INPUT
INTEGRITY
SW
ENG 1 DET FAIL
INPUT
ENGINE FIRE/OVERHEAT DETECTION SYSTEM - SCHEMATIC DIAGRAM

Sheet 1
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EFFECTIVITY: ALL
NOTES: NOTES:

EFFECTIVITY: ALL
FIRE/ENG/TRIM PANEL TEST PANEL
WHITE (PRESS)
RED (FIRE)
NOT SHUTOFF
HOT BUS 1
FIRE
SHUTOFF
ENG 2 SHUTOFF

E2 FIREX
INTEGRATED AVIONICS UNIT 2

(GIA 2)
ENG 2 FIRE EXT PWR

INPUT
(GEA 2)
FIRE
SW
ENG 2 FIRE DETECTOR

ENG 2 FIRE
INPUT
INTEGRITY
SW
ENG 2 DET FAIL
INPUT
ENGINE FIRE/OVERHEAT DETECTION SYSTEM - SCHEMATIC DIAGRAM

Sheet 2
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EFFECTIVITY: ALL
FIRE EXTINGUISHING 26-20
Introduction ENGINE FIRE EXTINGUISHING SYSTEM (26-21)
The function of the fire extinguishing system is to discharge fire extinguishing The engine fire extinguishing system has the function of discharging fire
agent in areas where fire/overheat can occur. This system has fixed and extinguishing agent in both engine compartments upon actuation of the
portable components charged with Halon 1301 agent, which permits the crew BOTTLE switch installed on the ENG FIRE EXTINGUISHER control panel in
to extinguish the fire. the cockpit.
General Description PORTABLE FIRE EXTINGUISHING SYSTEM (26-24)


The FIRE EXTINGUISHING includes these subsystems: The portable fire extinguishing system provides the flight crew with means to
control localized fire.
• ENGINE FIRE EXTINGUISHING (AMM SDS 26-21-00/1)
SYSTEM
• PORTABLE FIRE EXTINGUISH- (AMM SDS 26-24-00/1) The figure FIRE EXTINGUISHING - BLOCK DIAGRAM provides further data
ING SYSTEM on the preceding text.
The engine fire extinguishing system is basically composed of a single fixed

bottle that may be discharged in either LH (Left-Hand) or RH (Right-Hand)
engine by the related tubing to extinguish fire.
The bottle assembly is installed in the rear fuselage and is composed of a

container, two discharge outlets, three mounting lugs, two rupture disc
assemblies, two electroexplosive cartridges, one fill fitting assembly and one
TCPS (Temperature Compensated Pressure Switch).
Each discharge outlet has an explosive cartridge activated by the crew from
the cockpit by means of the ENG FIRE EXTINGUISHER control panel. The
fill fitting assembly works as a primary safety relief and the rupture disk
assembly as a secondary safety relief for overpressure. The TCPS is
responsible for monitoring the extinguishing agent for correct pressure.
There is also one portable fire extinguisher located in the cockpit that may be
discharged by the crew in areas where fire/overheat events occur.
The fire extinguishing agent used in the bottle is Halon 1301, and in the
portable extinguisher is a Halon 1301/1211 blend.
Components

EFFECTIVITY: ALL
FIRE EXTINGUISHING 26-20
EICAS

MESSAGES
ENGINE FIRE
EXTINGUISHING BOTTLE
CARTRIDGE 1
FIRE
PANEL
CARTRIDGE 2 GEA GIA
PRSOV
TCPS
FUEL
SOV
CMC
MESSAGES
FIRE EXTINGUISHING - BLOCK DIAGRAM
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
ENGINE FIRE EXTINGUISHING SYSTEM 26-21
Introduction rupture disc assemblies, three mounting lugs, a fill fitting assembly and a
TCPS (Temperature Compensated Pressure Switch).
The engine fire extinguishing system has the function of discharging fire
extinguishing agent in both engine compartments upon actuation of the The 573 cm³ (35 in³) container holds from 0.45 to 0.50 kg (1.0 to 1.1 lb) of
BOTTLE switch installed on the ENG FIRE EXTINGUISHER control panel in Halon 1301 agent, at pressure ranging from 4.136 to 4.309 MPa
the cockpit. (Megapascal) (600 to 625 psig) at 21 °C (70 °F) for an operating temperature
range of -62 to 85 °C (-80 to 185 °F).
General Description
The discharge outlet is machined from aluminum alloy and anodized for

The engine fire extinguishing system is capable of discharging extinguishing corrosion protection, except for mating surfaces that are alodized for electrical
agent (Halon 1301 - CBrF3) in both engines through the fire extinguishing conductivity. Debris from the disc and cartridge are prevented from entering
bottle installed in the aircraft rear fuselage. the discharge lines by a screen located in the discharge outlet. The discharge
head incorporates a drain valve and each one is responsible for discharging
Commands for the engine fire extinguishing discharges are provided through the agent in one engine.
the BOTTLE switch located on the ENG FIRE EXTINGUISHER control panel.
A control unit continuously monitors the readiness of the engine fire The pyrotechnic cartridge is hermetically sealed, and operates at a
extinguishing system. If the system fails, a caution indication in yellow is temperature range from 29 to 93 °C (85 to 200 °F). In normal operation, the
provided on the CAS (Crew Alerting System) window, and a maintenance explosive cartridge is fired and precipitates a high pressure shock wave,
message is recorded by the CMC (Central Maintenance Computer). which, in combination with high velocity cartridge fragments, cause the
prestressed disc to rupture and the agent to be released. The cartridge
Components connectors are keyed so as to prevent a misconnection between them.
ENGINE FIRE SHUTOFF BUTTONS AND EXTINGUISHING SWITCH The TCPS is responsible for monitoring the extinguishing agent for correct
pressure. The switch contact of the TCPS is normally open when the fire
The ENG FIRE EXTINGUISHER control panel comprises one shutoff extinguisher is properly charged and closed when sufficient pressure loss has
pushbutton for each engine (ENG 1 SHUTOFF and ENG 2 SHUTOFF) and occurred. The low pressure generates a CMC message (ENG FIREX
a fire extinguishing switch (BOTTLE). Pressing either engine shutoff BOTTLE LOW PRESS) in the PFD (Primary Flight Display). A press-to-test
pushbutton on the ENG FIRE EXTINGUISHER control panel enables the button is provided to check for low pressure condition. This press-to-test
BOTTLE switch. If the engine fire condition does not disappear, extinguishing button also tests the switch, TCPS internal wiring, connector, aircraft wiring
agent can be discharged on the respective engine selected through the and the CAS messages E1 FIREX FAIL and E2 FIREX FAIL. Failure of the
engine shutoff pushbutton upon actuation of the BOTTLE switch. The shutoff TCPS switch does not affect the fire extinguishing system operation.
pushbuttons are protected by a guard and the switch is protected by a lever
lock. ENGINE FIRE EXTINGUISHING DISCHARGE PIPING
ENGINE FIRE EXTINGUISHING BOTTLE The discharge piping has the function of allowing the discharge of the fire
extinguishing agent from the fire extinguishing bottle to the discharge outlet.
The engine fire extinguishing bottle consists of the following components: a The piping is designed to avoid ice blockage, is protected against corrosion
container, two discharge outlets and related electroexplosive cartridges and and is fire proof in the area of the engine compartment. Since the drain valve

EFFECTIVITY: ALL
NOTES: NOTES:

21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
is installed in the discharge outlet, the fire extinguishing bottle is located in
such a way as to avoid the installation of another valve in the tubing.
There is no piping inside the engine compartment; one discharge outlet for
each engine is available in the firewall to discharge Halon 1301 in the mid
cowl compartment.
The figure ENGINE FIRE EXTINGUISHING SYSTEM - COMPONENTS

LOCATION provides further data on the preceding text.

EFFECTIVITY: ALL
B 87.5 GA 87.5
ATR
ITT FIELD
N1%

IGN IGN
__
544 ITT C 350 __
N2%
OIL PRES PSI
OIL TEMP C
FUEL
ENG FIRE EXTINGUISHER

EICAS
SHUTOFF 1 BOTTLE SHUTOFF 2
DISCH A
OFF LWD
ENG START/STOP
EM500ENSDS260001F.DGN
RUN RUN
STOP START STOP START
ENG FIRE EXTINGUISHER

CONTROL PANEL
ENGINE FIRE EXTINGUISHING SYSTEM - COMPONENTS LOCATION
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
NOTES: NOTES:

EFFECTIVITY: ALL
ZONES DISCHARGE PIPING CONNECTION

317 (BAGGAGE COMPARTMENT)
318
321
413
423
B
A

C
ACCESS PANEL
BAGGAGE
COMPARTMENT
DOOR
DISCHARGE PIPING OUTLET ENGINE FIRE
(ENGINE COMPARTMENT) EXTINGUISHING
BOTTLE
C B
ENGINE FIRE EXTINGUISHING SYSTEM - COMPONENTS LOCATION
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
Operation ENGINE FIRE EXTINGUISHING SYSTEM - CAS MESSAGES (Continued)

If fire/overheat condition is detected in an engine compartment, the message
FIRE comes into view in the ITT (Interstage Turbine Temperature) field on LEVEL
INDICATION DESCRIPTION
the EICAS (Engine Indication Crew Alert System), the related engine fire (COLOR)
shutoff pushbutton (ENG 1 SHUTOFF or ENG 2 SHUTOFF) red light comes The pressure of the fire extinguishing
on and the aural warning FIRE is heard. bottle for LH (Left-Hand) engine is below
E1 FIREX Caution
minimum, the cartridge is already shot,
Upon pressing the ENG 1 SHUTOFF or ENG 2 SHUTOFF pushbutton: FAIL (Amber)
or there is no power available for shot-

• The related PRSOV (Pressure Regulating and Shutoff Valve) (AMM SDS ting.
36-11-00/1) and fuel shutoff valve (AMM SDS 28-21-00/1) close, avoiding The pressure of the fire extinguishing
air bleeding and fuel flow in the fire zone. bottle for RH (Right-Hand) engine is be-
E2 FIREX Caution
• A white stripe comes on to indicate that the fire ENG 1 SHUTOFF or ENG low minimum, the cartridge is already
FAIL (Amber)
2 SHUTOFF pushbutton was pressed. shot, or there is no power available for
shotting.
If fire/overheat condition persists in the engine compartment:
ENG FIREX Advisory
The bottle was discharged.
• The ENG 1 SHUTOFF or ENG 2 SHUTOFF pushbutton red light remains DISCH (White)
on, the message FIRE in the ITT field on the EICAS continues and the
aural warning FIRE is still heard. ENGINE FIRE EXTINGUISHING SYSTEM - CMC MESSAGES (Contin-
ued)
Upon setting the BOTTLE switch to the DISCH position:
INDICATION DESCRIPTION
• Extinguishing agent is released to the respective engine selected by the
ENG 1(2) FIREX BOTTLE CAR- Indicates malfunction of the E1(2)
fire ENG 1 SHUTOFF or ENG 2 SHUTOFF pushbutton.
TRIDGE FAIL cartridge.
• The message ENG FIREX DISCH comes into view on the CAS window. Indicates low pressure of extin-
ENG FIREX BOTTLE LOW
guishing agent (cartridge is opera-
When the overheat/fire condition is extinguished, the FIRE message goes out PRESS
tive).
of view from the ITT field of the EICAS, the related engine fire shutoff
pushbutton red light goes off and the aural warning FIRE is cancelled.
The signals can also be carried on two segregated harnesses in the area,
preventing system failure when rotor burst occurs.
The figure ENGINE FIRE EXTINGUISHING SYSTEM - SCHEMATIC

DIAGRAM provides further data on the preceding text.

EFFECTIVITY: ALL
FIRE/ENG/TRIM PANEL (GEA 1)
ENG 1 EXT BTL

ENG 1 SHUTOFF
INPUT

OFF ON
E1 CARTRIDGE
EXTING
OFF
ON
FIRE EXTINGUISHING BOTTLE

ENG 2 SHUTOFF
E2 CARTRIDGE
01
(GEA 2)
EM500ENSDS260003C.DGN
PRESS
TO TEST
ENG 2 EXT BTL
INPUT
PRESS SW
NORMAL
EXT BTL LOW PRESS
INPUT
LOW
PRESS
01 ROTOR BURST ZONE
ENGINE FIRE EXTINGUISHING SYSTEM - SCHEMATIC DIAGRAM
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
Training Information Points
During the procedures of removal and installation of the extinguishing bottles,

the cartridges must be kept protected until the installation of electrical mating
connector to prevent accidental detonation. Use only protective caps for the
protection of connector pins or sockets in electrical connectors, because
other materials can cause damage to the connector pins or sockets, or let
unwanted materials stay in the connector.

The figure ENGINE FIRE EXTINGUISHING SYSTEM - CARTRIDGES
PROTECTION provides further data on the preceding text.

EFFECTIVITY: ALL
A
ZONE

321
C A
C ENGINE FIRE B
B
EXTINGUISHING
BOTTLE C
CAUTION
DO NOT REMOVE PROTEC
B
TIVE
COVER UNTIL INTALLATION
OF MATING ELECTRICAL
CONNECTOR
CARTRIDGES
C
ENGINE FIRE EXTINGUISHING SYSTEM - CARTRIDGES PROTECTION
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
PORTABLE FIRE EXTINGUISHING SYSTEM 26-24
Introduction
The portable fire extinguishing system provides the flight crew with means to
control localized fire.
General Description
The portable fire extinguishing system is composed of portable fire

extinguisher attached to the aircraft by means of quick release bracket.

The lightweight fire extinguisher installed in the cockpit area is charged with
Halon 1211/1301 blend which is highly effective against fires Class B and C,
and has low toxicity characteristics.
Components
The fire extinguisher consist of a cylinder made of aluminum with a handle,

activating lever, nozzle and a safety pin. The unit is compact, lightweight, safe
and easy to operate. The Cabin Fire Protection system include one portable
fire extinguisher, charged with 1.2 kg (2.5 lb) of Halon 1211/1301 blend,
installed inside the cockpit, behind the copilot seat.
Operation
The operation of the portable fire extinguisher is as follows:
• Hold the bottle upright
• Remove the safety pin
• Direct the nozzle toward the base of the fire
• Press the activating lever
• Sweep side to side
The figure PORTABLE FIRE EXTINGUISHING SYSTEM - COMPONENT


EFFECTIVITY: ALL
PORTABLE FIRE EXTINGUISHING SYSTEM 26-24

BRACKET
A
ZONE B
224
CLAMP
A
PORTABLE FIRE
EXTINGUISHER
B
PORTABLE FIRE EXTINGUISHING SYSTEM - COMPONENT LOCATION
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80



CHAPTER 28 - FUEL
SECTION TITLE PAGE

28-00 FUEL 44

28-10 STORAGE 50
28-11 WING TANK 54
28-12 TANK VENT 74
28-20 DISTRIBUTION 84
28-21 ENGINE FEED SYSTEM 86
28-40 INDICATING 104
28-41 ELECTRICAL FUEL QUANTITY INDICATING 110
28-43 FUEL TEMPERATURE INDICATION SYSTEM 124
28-45 FUEL LOW PRESSURE WARNING SYSTEM 130
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
FUEL 28-00
Introduction
The main function of the EMB-500 fuel system is to contain and supply fuel
to the engines, providing indication for the fuel carried on board the aircraft.
The figure FUEL - LOCATION provides further data on the preceding text.

EFFECTIVITY: ALL
FUEL 28-00

RIGHT WING TANK
(ZONES 631,632 AND 641)
RIB 15
RIB 14
RIB 7
RIB 3
RIB 1
RIB 3
RIB 7 LEFT WING TANK
(ZONES 531,532 AND 541)
RIB 14
RIB 15
FUEL - LOCATION
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
FUEL 28-00
General Description Inter wing balancing of fuel load is achieved by gravity, via an interconnecting
transfer valve.
The FUEL includes these subsystems:
Refueling is accomplished by gravity, through a filler neck on each wing upper
• STORAGE (AMM SDS 28-10-00/1) surface.
• DISTRIBUTION (AMM SDS 28-20-00/1)
• INDICATING (AMM SDS 28-40-00/1) The vent system has been sized to avoid exceeding the 5 psig fuel tanks
structural limit during normal aircraft operation.
The fuel system is powered by a 28 V DC power source. Control and

monitoring functions for the fuel system are provided by the EFCU (Electronic The figure FUEL - BLOCK DIAGRAM provides further data on the preceding
Fuel Control Unit) and other avionics units. text.
Operation
The fuel system continuously supplies fuel to the engines at a minimum

pressure of TVP (True Vapor Pressure) + 6.25 psia (43.0 kPa) or 2 psia (13.8
kPa) above ambient pressure, whichever is greater, in normal operating
conditions throughout the aircraft normal flight envelope.
Fuel is contained in two integral wing tanks, one in each wing. Each wing
supplies its respective engine through a feed system independent of the other
engine.
Normal engine feed is done through ejector pumps. The ejector pumps in
each wing are driven by high-pressure motive flow returned from the engines.
Electrical power is not required for normal engine fuel feed operation.
Scavenge ejectors in each wing are also used to minimize unusable fuel. Two
electrical pumps, one in each wing, are provided for engine start operation,
and to work under ejector pump failure condition.
There is no power wiring inside the fuel tanks.
The fuel gauging subsystem provides an accurate measure of the fuel mass
in the fuel tanks. The fuel gauging subsystem also provides fuel low level and
temperature indication. In addition, fuel conditions are displayed on the MFD
(Multi-Function Display) fuel synoptic page, in the cockpit.

EFFECTIVITY: ALL
FUEL 28-00
RIB 1
RIB 2
RIB 2
RIB 3
RIB 3
RIB 4
RIB 4
RIB 9
RIB 9
RIB 12
RIB 12
RIB 13
RIB 13
RIB 14
RIB 14
DCM
RIB 15
RIB 15
SPAR I
SPAR I

DCM DCM
SPAR II SPAR II
D D
SPAR III SPAR III

DCM
DCM
LEGEND:
SCAVENGE EJECTOR PUMP NACA INLET
ENGINE FEED EJECTOR PUMP PS PS FLOAT VENT VALVE

DCM
DC AUXILIARY BOOST PUMP FLAP VALVE

DCM
SHUTOFF VALVE ( DC MOTOR OPERATED) BAFFLE CHECK VALVE
ENGINE
ENGINE
CHECK VALVE DRAIN VALVE
PS ENGINE PRESSURE SWITCH MOTIVE FLOW LINE
VENT LINE DRAIN ORIFICE FUEL FEED LINE
COLLECTOR TANK VENT ORIFICE SCAVENGE/ TRANSFER LINE
GRAVITY REFUELING ADAPTER VENT LINE
D DUMP VALVE
FUEL - SCHEMATIC DIAGRAM
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
FUEL 28-00
NOTES: NOTES:

EFFECTIVITY: ALL
FUEL 28-00
EMERGENCY BUS DC BUS 2 EMERGENCY BUS
VDC VDC
AUXILIARY AUXILIARY
BOOST BOOST
PUMP 1 PUMP 2
ENG 1 FIRE
DETECTOR VDC AUXILIARY VDC AUXILIARY ENG 2 FIRE
BOOST PUMP 1 BOOST PUMP 2 DETECTOR
RELAY RELAY

ENGINE/
AIRFRAME
OPEN ENG 1 FUEL INTEGRATED UNIT (GEA 1)
SHUTOFF AVIONICS OPEN
SHUTOFF
VALVE (SOV) UNIT (GIA 1) ON ON SHUTOFF
ENG 1 SHUTOFF AUTO AUTO
PUSHBUTTON ENG 2 SHUTOFF
OFF OFF
PUSHBUTTON
ENGINE/ ENGINE/
FUEL TO CLOSE AIRFRAME AIRFRAME ENG 2 FUEL
TRANSFER UNIT (GEA 2) VDC VDC UNIT (GEA 3) SHUTOFF
VALVE (SOV) FUEL AUXILIARY AUXILIARY VALVE (SOV)
TO OPEN TRANSFER BOOST BOOST
VALVE FUEL TEMP PUMP 1 PUMP 2
PANEL SW SENSOR PANEL SW PANEL SW
DATA ARINC 429 EFCU SERIAL DATA LINK EFCU INTEGRATED

CONCENTRATOR ARINC 429 AVIONICS
CHANNEL 1 CHANNEL 2
UNIT UNITS (GIA 2)
LH TANK RH TANK
HI UNIT ARRAY HI
UNIT ARRAY
LOW LOW
ENG 1 PRESS SW ENG 2 PRESS SW
FUEL - BLOCK DIAGRAM
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
STORAGE 28-10
Introduction
The storage subsystem is responsible for keeping the fuel in correct

conditions. It contains the fuel tanks and the vent system.
The figure STORAGE - LOCATION provides further data on the preceding

text.

EFFECTIVITY: ALL
STORAGE 28-10

RIB 15
(YW/=5504.92)
RIB 14
(YW/=4994.92)
RIGHT MAIN
TANK
SURGE TANK
RIB 7
(Y=1899.09)
RIB 3
(Y=635.00)
RIB 1
(Y=0.00) RIB 3 RIB 7
(Y=−635.00) (Y=−1899.09)
LEFT MAIN
TANK
RIB 14
(YW/=−4994.92)
RIB 15
COLLECTOR (YW/=−5504.92)
TANK
SURGE TANK
STORAGE - LOCATION
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
STORAGE 28-10
General Description
The STORAGE includes these subsystems:
• WING TANK (AMM SDS 28-11-00/1)

• TANK VENT (AMM SDS 28-12-00/1)
The storage subsystem contains the fuel tanks, the pressure relief
components, and a vent system. The total capacity of the fuel storage

subsystem is approximately 1610 (425 gal.).
The figure STORAGE - SCHEMATIC DIAGRAM provides further data on the

preceding text.

EFFECTIVITY: ALL
STORAGE 28-10
RIB 1
RIB 2
RIB 2
RIB 3
RIB 3
RIB 4
RIB 4
MAIN VENT MAIN VENT
RIB 9
RIB 9
RIB 12
RIB 12
LINE LINE
RIB 13
RIB 13
RIB 14
RIB 14
RIB 15
RIB 15

SPAR I
SPAR I
SPAR II SPAR II
D D
SPAR III SPAR III
SURGE TANK COLLECTOR TANK SURGE TANK

MAIN TANK MAIN TANK
LEGEND:
VENT LINE DRAIN ORIFICE FLOAT VENT VALVE
COLLECTOR TANK VENT ORIFICE FLAP VALVE
GRAVITY REFUELING ADAPTER BAFFLE CHECK VALVE
D DUMP VALVE DRAIN VALVE
NACA INLET VENT LINE
STORAGE - SCHEMATIC DIAGRAM
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
WING TANK 28-11
Introduction
The aircraft is provided with two integral (wet) wing tanks. The wing tanks are
the main structure for the storage and distribution of fuel.
The figure WING TANK - LOCATION provides further data on the preceding
text.

EFFECTIVITY: ALL
WING TANK 28-11

RIB 15
(YW/=5504.92)
RIB 14
(YW/=4994.92)
RIGHT MAIN
TANK
SURGE TANK
RIB 7
(Y=1899.09)
RIB 3
(Y=635.00)
RIB 1
(Y=0.00) RIB 3 RIB 7
(Y=−635.00) (Y=−1899.09)
LEFT MAIN
TANK
RIB 14
(YW/=−4994.92)
RIB 15
COLLECTOR (YW/=−5504.92)
TANK
SURGE TANK
WING TANK - LOCATION
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
WING TANK 28-11
General Description The compartments between rib 14 and rib 15 in the wing tips serve as surge
tanks and do not normally carry fuel. The surge tanks collect fuel that enters
The two wing tanks are physically isolated and are independently gauged and the fuel tank vent sub-subsystem (AMM SDS 28-12-00/1) during wing-down
refueled. The wing tanks are bounded by the wing spar 1 and spar 3, the and uncoordinated maneuvers. At the end of the maneuver, the fuel returns
upper and lower wing skin surfaces, and the rib 1 (center wing rib) and rib 14, to the main tank through a flap valve located at the lowest point of rib 14.
at stations Y=0 and Y=4994.92 respectively. The wing tanks are also
bounded by the main landing gear wheelwell (by spar 2, between rib 3 and The top of each rib has openings to prevent air pockets from forming in the
rib 7, at stations Y=635.00 and Y=1899.09 respectively). The arrangement of wing tanks. The ribs also have provisions to prevent accumulation of water
the tank structure is designed to permit the fuel to flow from the wing tip to and trapped fuel. Holes at the bottom of each open rib provide passage for

the wing root. The bottom portions of all the ribs in the wing are free from drainage in the direction of the drain valves. Baffle valves are installed at
obstructions to allow the fuel flow. This prevents the collection of fuel on the closed ribs 9 and 12 to restrict outboard fuel movement during maneuvers.
ribs and stiffeners.
The fuel dumping is accomplished by means of a dump valve located in the
The interior of the tanks is chemically treated against corrosion and coated wing bottom skin in each collector tank to which a hose can be connected.
with a biocide compound. The fuel system components and the plumbing One manually operated water drain valve is integrally installed in the dump
inside the tanks are electrically bonded to prevent arcing due to lightning valve. The primary sealing is achieved through the water drain valve seals.
strikes. Sealing compound is applied to all joints and riveting areas to ensure A secondary sealing (metal/metal) is provided in the dump valve in case of
tank seal integrity. failure of drain valve seals.
Each wing tank is divided into three compartments: Lightning protected access panels are provided in the wing lower surface in
order to allow inspection and repair of internal tank structure as well as
• Collector Tank removal and replacement of any component located inside the tanks.
• Surge Tank Refueling is accomplished by gravity through a filler neck on the upper
surface of each wing, the location of which prevents the refueling operator
• Main Tank
from exceeding the fuel capacity. If desired, both wings can be filled from one
The inboard part of each wing tank is used as a partially sealed collector tank. side up to 60% of total capacity by opening the gravity transfer shutoff valve
The collector tanks are located between rib 1 (center wing rib), at Y=0, and (see AMM SDS 28-21-00/1).
rib 3, at Y=635, and between spar 2 and spar 3. These tanks supply
The wing tanks have the components that follow:
continuous fuel feed to the engines and minimize the amount of unusable
fuel. The collector tanks are supplied with fuel by gravity through the three • Baffle Check/Flap Valves
flap valves installed at spar 2. Scavenge ejector pumps (AMM SDS
28-21-00/1) installed in the main tanks are required to maintain the collector • Fuel Tank Access Panels
tanks fuel supply during all attitudes in the operational envelope. The capacity
of each collector tank, which forms part of the usable capacity, is • Drain/Dump Valves
approximately 135 (35.7 gal.).
• Gravity Refueling Adapters/Gravity Refueling Caps

EFFECTIVITY: ALL
WING TANK 28-11
NOTES: NOTES:

21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
WING TANK 28-11
The figure WING TANK - SCHEMATIC DIAGRAM provides further data on

the preceding text.

EFFECTIVITY: ALL
WING TANK 28-11
RIB 1
RIB 2
RIB 2
RIB 3
RIB 3
RIB 9
RIB 9
RIB 12
RIB 12
RIB 13
RIB 13
RIB 14
RIB 14
RIB 15
RIB 15
SPAR I
SPAR I

SPAR II SPAR II
D D
SPAR III SPAR III
LEGEND:
GRAVITY REFUELING ADAPTER
D DUMP VALVE
FLAP VALVE
BAFFLE CHECK VALVE
DRAIN VALVE
WING TANK - SCHEMATIC DIAGRAM
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
WING TANK 28-11
Components GRAVITY FILLER NECKS/CAPS
BAFFLE CHECK/FLAP VALVES Two gravity filler necks are installed on the aircraft for gravity refueling. There
is one filler neck on the upper surface of each wing between rib 12 and rib
The 2.2 in. baffle check valves are one-way flapper valves that control the 13.
flow of fuel inboard. There are three baffle check valves in each wing tank.
Two baffle check valves near the bottom of rib 12 and one near the bottom The gravity filler necks are provided with caps. The caps are flush mounted
of rib 9. to minimize aerodynamic drag and ensure that no fuel can be trapped. The
caps have an integral seal to prevent fuel leaks and are lightning-strike proof.

The flap valves are one-way flapper valves that control the flow of fuel Lanyards retain the caps when they are removed from the gravity filler necks.
inboard. There are four 1 in. flap valves in each wing tank. Three flap valves The caps are key locked for security against unauthorized entry.
are installed at the bottom of the collector tank spar 2 boundary. One flap
valve is installed at the bottom of rib 14, at Y = 4994.92, in each wing. Gravity refueling protection nets are installed in both gravity filler necks to
provide a protection for the bottom wing skin against damage from the
FUEL TANK ACCESS PANELS refueling nozzle.
Each wing tank has 16 access panels installed on the lower wing skin. Each The figure WING TANK - COMPONENT LOCATION provides further data on
surge tank has one access panel installed on the lower wing skin. One access the preceding text.
panel also gives access to each collector tank. The access panels allow the
inspection and repair of the internal structure of the tank. The access panels
also allow the inspection, repair, and replacement of components located
inside the wing tanks.
DUMP/DRAIN VALVES
The water drain valves are operated manually and allow the removal of water
and contaminants from the wing tanks. They are also used to remove
remaining fuel from the wing tanks after they have been defueled. The
primary sealing is achieved through the water drain valve seals. The drain
valves are spring-loaded poppet valves. There is one drain valve in each wing
tank located in the bottom skin of each wing, at the collector tank, and fitted
inside the dump valve assembly.
The dump valves are operated manually and allow the defueling operation.
A secondary sealing (metal/metal) is provided in the dump valve in case of
failure of drain valve seals. There are two dump valves: one located in the
bottom skin of each wing, and the other at the collector tank, to which a hose
can be connected.

EFFECTIVITY: ALL
WING TANK 28-11
RIB 2
(Y=340) RIB 3
B (Y=635)
RIB 9
(Y=2731.00)
C
RIB 12
RIB 1 (YW/=4102.00)
(Y=0)

RIB 14
FLAP VALVES C (YW/=4994.92)
B
A
ZONES
530/541
A
630/641
BAFFLE CHECK
VALVES
FLAP VALVE
BAFFLE
CHECK VALVE
FLAP VALVE
C B
TYPICAL TYPICAL
WING TANK - COMPONENT LOCATION

Sheet 1
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
WING TANK 28-11
NOTES: NOTES:

EFFECTIVITY: ALL
WING TANK 28-11
ZONE
641 B RIB 12
A
RIB 13
ZONE
RIB 13 541
RIB 12
A

RIB 12
RIB 13
GRAVITY
GRAVITY REFUELING
REFUELING PROTECTION−NET
ADAPTER
A
TYPICAL
GRAVITY FILLER GRAVITY
CAP KEY LOCK FILL CAP
B

Sheet 2
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
WING TANK 28-11
NOTES: NOTES:

EFFECTIVITY: ALL
WING TANK 28-11
LOWER WING
SKIN (REF.)

A
ZONES
530/541 SEAL
630/641
B
WING TANK SCREW
ACCESS PANEL (16x)
B
TYPICAL
A

Sheet 3
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
WING TANK 28-11
NOTES: NOTES:

EFFECTIVITY: ALL
WING TANK 28-11

A
ZONES
532
632
LOWER WING
B
SKIN (REF.)
DUMP VALVE
DRAIN VALVE

Sheet 4
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
WING TANK 28-11
Operation
DRAIN VALVE OPERATION
The drain valves have three positions:
POSITION FUNCTION
Closed Normal position of drain valve.

Draining water, contaminants, and
Open
remaining fuel from wing tanks.
Service Replacing drain valve seal packing.
The service position allows the replacement of the packing without removal
of the drain valve or defueling of the wing tanks.
DUMP VALVE OPERATION
The dump valve is opened by a dump tool TOOL, DEFUELING (GSE 023).
The figure WING TANK - DUMP VALVE OPERATION provides further data
on the preceding text.

EFFECTIVITY: ALL
WING TANK 28-11

A 523 BL B
ZONES
532
632 A
TO LOCK OPEN FROM TO CLOSE FROM FOR MAINTENANCE

CLOSED POSITION: OPEN POSITION: FROM CLOSED POSITION:
1. PUSH UP AND TURN 1. TURN CLOCKWISE 1. TURN CLOCKWISE
COUNTER−CLOCKWISE TO CLOSE POSITION TO POPPET−DOWN
TO OPEN POSITION POSITION
B B B
WING TANK - DRAIN VALVE OPERATION

Sheet 1
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
WING TANK 28-11
NOTES: NOTES:

EFFECTIVITY: ALL
WING TANK 28-11

TO OPEN AND DRAIN
WATER FROM CLOSED TO CLOSE FROM DRAIN
POSITION: POSITION:
1. PUSH THE FUEL SAMPLE 1. PULL THE FUEL SAMPLE

TEST TOOL UP TEST TOOL DOWN
B B
WING TANK - DRAIN VALVE OPERATION

Sheet 2
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
WING TANK 28-11
NOTES: NOTES:

EFFECTIVITY: ALL
WING TANK 28-11
523 BL B

A A B
ZONES
532
632
− TO DEFUEL THE FUEL TANK
1 2 3
INSTALL THE DEFUELING TOOL
IN THE DUMP VALVE.
REMOVE THE DRAIN VALVE FROM INSTALL THE SCREWS IN THE TURN COUNTER−CLOCKWISE UNTIL THE
THE DUMP VALVE. DUMP VALVE. SCREWS HOLD THE TOOL ASSEMBLY
WING TANK - DUMP VALVE OPERATION
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
TANK VENT 28-12
Introduction
The fuel tank vent sub-subsystem, keeps the fuel pressure differential
between the fuel tanks and the atmosphere within the +5 psig structural limit
during all operating conditions.
The vent sub-subsystem also prevents fuel spillage during flight maneuvers
and hard braking.

The figure TANK VENT - LOCATION provides further data on the preceding
text.

EFFECTIVITY: ALL
TANK VENT 28-12
ZONES
541
641
A

NACA INLET
RIB 1 NACA CONNECTING
VENT LINE
RIB 3
FLOAT VENT
VALVE
MAIN TANK
VENT LINE
A
RIB 12
RIB 14
TANK VENT - LOCATION
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
TANK VENT 28-12
General Description
Each wing tank is vented through two independent 3/4 in. main vent lines
connected to the surge tanks (AMM SDS 28-11-00/1). The surge tank,
between rib 14 and rib 15, is vented through a NACA (National Advisory
Committee for Aeronautics) air inlet installed on the lower wing skin 1.9 m
inboard of the wing tip, to provide means of protecting the vent inlet against
the effects of lightning strike and corona discharge. The NACA inlet, installed
in the wet fuel tank zone, is connected to the surge tank via a 5/8 in. pipe in

each wing.
The inboard part of both wing tanks is vented through the 3/4 in. vent lines.
The vent line in each wing runs from the inboard part of the tank, near the
center rib, to the surge tank.
The outboard part of the wing tank is vented directly to the surge tank through
a float valve attached to rib 14, at station Y=4994.92.
The vent lines are so arranged that at least one line is always open during all
flight conditions. The vent lines provide adequate protection for the wing tanks
during all flight and ground operations.
The top portions of all the ribs in the wing are free from obstructions to allow
air to flow between the wing compartments.
Fuel or water trapped in the vent pipes is drained into the fuel tanks through
orifices at the lowest points of each vent line.
The fuel tank vent sub-subsystem has the components that follow:
• Float Vent Valves
The figure TANK VENT - SCHEMATIC DIAGRAM provides further data on

the preceding text.

EFFECTIVITY: ALL
TANK VENT 28-12
RIB 1
RIB 2
RIB 2
RIB 3
RIB 3
RIB 4
RIB 4
MAIN TANK MAIN TANK
RIB 9
RIB 9
VENT LINE VENT LINE
RIB 12
RIB 12
RIB 13
RIB 13
RIB 14
RIB 14
RIB 15
RIB 15

SPAR I
SPAR I
SPAR II SPAR II
SPAR III SPAR III
LEGEND:
NACA INLET
FLOAT VENT VALVE
VENT LINE DRAIN ORIFICE
COLLECTOR TANK VENT ORIFICE
VENT LINE
TANK VENT - SCHEMATIC DIAGRAM
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
TANK VENT 28-12
Components
FLOAT VENT VALVES
There are two float vent valves in the fuel tank vent sub-subsystem. A float
valve is installed in each wing tank on rib 14. The float valve consists of a
check valve attached to a float arm. The float valves vent pressure from the
outboard area of the wing tanks and prevent fuel flow into the surge tanks.

The figure TANK VENT - COMPONENT LOCATION provides further data on
the preceding text.

EFFECTIVITY: ALL
TANK VENT 28-12
ZONES RIB 14
541
641 RIB 12
A
C
RIB 5
RIB 1

B RIB 5
RIB 12
RIB 14
A
RIB 5
C
RIB 4
RIGHT MAIN RIB 3 LEFT MAIN

VENT LINE VENT LINE
RIB 2
RIB 1
RIB 2
RIB 3
RIB 4
B RIB 5
TANK VENT - COMPONENT LOCATION

Sheet 1
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
TANK VENT 28-12
NOTES: NOTES:

EFFECTIVITY: ALL
TANK VENT 28-12
D

RIB 12
RIB 13
E
C
RIB 14
FLOAT VENT
VALVE
NACA
D E
TANK VENT - COMPONENT LOCATION

Sheet 2
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
TANK VENT 28-12
Operation
FLOAT VENT VALVE OPERATION
When the fuel level in the wing tank decreases, the float valve opens to vent
pressure from the outboard area of the wing tank. When the fuel level rises
due to aircraft maneuvers or refueling, the float valve closes to prevent fuel
spillage into the surge tank.

The figure TANK VENT - FLOAT VENT VALVE OPERATION provides further
data on the preceding text.

EFFECTIVITY: ALL
TANK VENT 28-12
FLOAT VENT
FLOAT VENT VALVE
VALVE

FLOAT
FUEL
LEVEL
FLOAT
FUEL
LEVEL
NOTE: OPEN POSITION NOTE: CLOSED POSITION
ACTION: THE FUEL LEVEL IS BELOW THE ACTION: THE FUEL LEVEL IS AT OR ABOVE THE
FLOAT VENT VALVE. FLOAT VENT VALVE.
RESULT: THE FLOAT LOWERS (THE VENT LINE IS OPEN). RESULT: THE FLOAT RAISES (THE VENT LINE IS CLOSED).
RESULT: THE FUEL TANK IS VENTED.
TANK VENT - FLOAT VENT VALVE OPERATION
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
DISTRIBUTION 28-20
Introduction
The distribution subsystem controls the movement of fuel within the wing
tanks and engines. The distribution subsystem supplies the flow of fuel for
the aircraft fuel feed.
General Description
The DISTRIBUTION includes this subsystem:


• ENGINE FEED SYSTEM (AMM SDS 28-21-00/1)
The figure DISTRIBUTION - SCHEMATIC DIAGRAM provides further data


EFFECTIVITY: ALL
DISTRIBUTION 28-20
RIB 1
RIB 2
RIB 2
RIB 3
RIB 3
RIB 9
RIB 9
RIB 12
RIB 12
RIB 13
RIB 13
RIB 14
RIB 14
DCM
RIB 15
RIB 15
SPAR I
SPAR I

DCM DCM
SPAR II SPAR II
SPAR III SPAR III

DCM
DCM
LEGEND:
PS PS
SCAVENGE EJECTOR PUMP
ENGINE FEED EJECTOR PUMP

DCM
DC AUXILIARY BOOST PUMP

ENGINE
ENGINE
DCM
SHUTOFF VALVE ( DC MOTOR OPERATED)
CHECK VALVE
PS ENGINE PRESSURE SWITCH
MOTIVE FLOW LINE
FUEL FEED LINE
SCAVENGE/ TRANSFER LINE
DISTRIBUTION - SCHEMATIC DIAGRAM
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
ENGINE FEED SYSTEM 28-21
Introduction
The primary function of the engine fuel feed sub-subsystem is to supply fuel
to the engines during aircraft operation. There is a separate system for each
engine in the fuel feed sub-subsystem. The engine fuel feed sub-subsystem
also transfers fuel to the collector tank, isolates the fuel if there is an engine
fire, and equalizes the fuel quantity between the two wings (gravity transfer).
The figure ENGINE FEED SYSTEM - LOCATION provides further data on


the preceding text.

EFFECTIVITY: ALL
ENGINE
SHUTOFF
VALVE ENGINE
FEED LINE
MOTIVE FLOW
CHECK VALVE
ENGINE FEED
MOTIVE FLOW
EJECTOR PUMP
LINE

ENGINE FEED MOTIVE FLOW
CHECK VALVE LINE
A ENGINE
FEED LINE
ZONES MOTIVE FLOW
522/622 CHECK VALVE
530/630
ENGINE
SHUTOFF
VALVE
VDC AUXILIARY
BOOST PUMP
SCAVENGE
EJECTOR
PUMP
FUEL TRANSFER
VALVE VDC AUXILIARY
BOOST PUMP
A
ENGINE FEED SYSTEM - LOCATION
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
General Description The FUEL and the FIRE extinguisher control panels control the operation of
the engine fuel feed sub-subsystem.
The engine fuel feed sub-subsystem supplies correct fuel flow to the engines The position and function of the applicable control panel and fire switches are
during all operational conditions at a pressure that obeys limits given by the given in the table below.
engine manufacturer.
The figure ENGINE FEED SYSTEM - SCHEMATIC DIAGRAM provides
The engine fuel feed sub-subsystem comprises these components: further data on the preceding text.
• Engine Feed Ejector Pumps

• V DC Auxiliary Boost Pump Cartridges
• V DC Auxiliary Boost Pump Canisters
• V DC Pump Pressure Switches
• Engine SOV (Shutoff Valve)s
• Engine SOV Actuator Assemblies
• Fuel Transfer Valve
• Fuel Transfer Valve Actuator
• Scavenge Ejector Pumps
• Engine Feed Check Valves
• Motive Flow Check Valves
• FUEL Control Panel
The engines are normally fed by the engine feed ejector pumps. A V DC
auxiliary pump in each collector tank is provided for the engines during start
and in case of ejector pump failure. The V DC auxiliary pumps operation is
controlled by the EFCU (Electronic Fuel Control Unit) and powered by the
EMERGENCY BUS bar.

EFFECTIVITY: ALL
RIB 1
RIB 2
RIB 2
RIB 3
RIB 3
RIB 9
RIB 9
RIB 12
RIB 12
RIB 13
RIB 13
RIB 14
RIB 14
DCM
RIB 15
RIB 15
SPAR I
SPAR I

DCM DCM
SPAR II SPAR II
SPAR III SPAR III

DCM
DCM
LEGEND:
PS PS
SCAVENGE EJECTOR PUMP
ENGINE FEED EJECTOR PUMP

DCM
DC AUXILIARY BOOST PUMP

ENGINE
ENGINE
DCM
CHECK VALVE
MOTIVE FLOW LINE
FUEL FEED LINE
SCAVENGE/ TRANSFER LINE
ENGINE FEED SYSTEM - SCHEMATIC DIAGRAM
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
(Continued)
REF CONTROL POSITION FUNCTION
1 XFR Pushbutton Pushed Opens the fuel transfer valve.
Not Pushed (normal posi-
Closes the fuel transfer valve.
tion)
2 PUMP 1 Switch OFF Turns the LH (Left-Hand) V DC auxiliary pump off.

Allows automatic operation of the LHV DC auxiliary pump during engine
AUTO
start, or when the engine feed ejector pump fails.
ON Turns the LHV DC auxiliary pump on.
3 PUMP 2 Switch OFF Turns the RH (Right-Hand) V DC auxiliary pump off.
Allows automatic operation of the RHV DC auxiliary pump during engine
AUTO
start, or when the engine feed ejector pump fails.
ON Turns the RHV DC auxiliary pump on.
4 ENG 1 SHUTOFF Keeps the engine 1 SOV open.
tion)
Pushed Closes the engine 1 SOV.
5 ENG 2 SHUTOFF Keeps the engine 2 SOV open.
tion)
Pushed Closes the engine 2 SOV.
6 EXTING DISCH Activates the fire extinguishing system for the applicable engine.

EFFECTIVITY: ALL
ENG FIRE EXTINGUISHER TRIM

BOTTLE YAW
SHUTOFF 1 SHUTOFF 2
LEFT RIGHT
DISCH
ROLL
OFF LWD RWD
ENG START/STOP
A STOP
RUN
START STOP
RUN
START
PITCH BKP

DN
UP
1 2
ENG IGNITION MODE
+
ON BKP
AUTO
OFF OFF
1 2
FIRE CONTROL PANEL
B A
FUEL STALL WRN
PUMP 1 XFR PUMP 2 INHIB
ON ON
AUTO AUTO
OFF OFF
PAX SIGNS ELT HYD PUMP

AUTO
OFF ON
PED−BELTS ON
BELTS ARMED
OFF TEST/RESET
FUEL CONTROL PANEL
ENGINE FEED SYSTEM - CONTROL
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
The CAS (Crew Alerting System) messages related to the engine fuel feed
sub-subsystem are listed in the table below:
(Continued)
INDICATION LEVEL (COLOR) DESCRIPTION
FUEL 1 SOV FAIL Caution (Amber) The left engine SOV has failed.
FUEL 2 SOV FAIL Caution (Amber) The right engine SOV has failed.

There is a discrepancy between the transfer valve command and
FUEL XFR FAIL Caution (Amber)
its feedback.
The transfer valve status can lead to loss of fuel through the vent
FUEL OVERFILL Caution (Amber)
system.
FUEL PUMP 1 FAIL Advisory (White) The LHV DC auxiliary pump has failed.
FUEL PUMP 2 FAIL Advisory (White) The RHV DC auxiliary pump has failed.
FUEL EQUAL Advisory (White) The fuel transfer valve is open and there is not a fuel imbalance.
The LHV DC auxiliary pump is on due to a low pressure detected
FUEL 1 FEED FAULT Advisory (White)
by the pressure switch.
The RHV DC auxiliary pump is on due to a low pressure detected
FUEL 2 FEED FAULT Advisory (White)
by the pressure switch.

EFFECTIVITY: ALL
DC AUXILIARY FUEL TRANSFER DC AUXILIARY
BOOST PUMP 1 VALVE (SOV) BOOST PUMP 2
A
RH ENGINE FEED
LH ENGINE FEED EJECTOR PUMP
EJECTOR PUMP
XFR

XXX LB XXX LB
RH ENGINE
LH ENGINE SHUTOFF VALVE
SHUTOFF VALVE TOTAL (SOV)
(SOV) XXXX LB
USED
LH FUEL XXX LB
PRESSURE RH FUEL
LEGEND: SWITCH PRESSURE
SWITCH
SOV IS OPEN (COLOUR FOLLOWS LINE DOWNSTREAM)
SOV IS CLOSED (WHITE)

MFD
(SYNOPTIC)
SOV IS FAILED (WHITE UNDER X IN RED)
TRANSITIONAL STATE − OPEN TO CLOSED A

OR CLOSED TO OPEN (WHITE)
GREEN = EJECTOR PUMP IS IN OPERATION GREEN = HIGH PRESSURE
WHITE = EJECTOR PUMP IS NOT IN OPERATION WHITE = LOW PRESSURE
WHITE UNDER X IN RED = EJECTOR PUMP IS FAILED WHITE UNDER X IN RED = PRESSURE SWITCH IS FAILED
GREEN = DC PUMP IS IN OPERATION GREEN = LINE IS UNDER SYSTEM OPERATION

WHITE = DC PUMP IS NOT IN OPERATION WHITE = LINE IS NOT UNDER SYSTEM OPERATION
WHITE UNDER X IN RED = DC PUMP IS FAILED WHITE UNDER X IN RED = SYSTEM COMMUNICATION IS FAILED
ENGINE FEED SYSTEM - FUEL SYNOPTIC PAGE
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
Components ingestion of foreign objects. A check valve is installed to the outlet of the V
DC auxiliary pump canister to prevent fuel flow in the wrong direction.
ENGINE FEED EJECTOR PUMPS
ENGINE SHUTOFF VALVES (SOV)
There are two engine feed ejector pumps in the engine fuel feed sub-
subsystem. There is one ejector pump installed in each collector tank. The There are two engine SOVs in the engine fuel feed sub-subsystem. A SOV
ejector pumps are the primary source of fuel supply to the engines. The is installed in each engine feed line to stop the flow of fuel in case of engine
ejector pumps are venturi-type pumps with no moving parts that draw fuel fire. The SOVs are installed on the wing-to-fuselage fairing, outside the fuel
from the collector tanks when fed with motive flow. The ejector pumps receive tank. They are ball valves, controlled through the engine SOV actuator

their motive flow from the engine-driven fuel pumps. assemblies. Thermal relief is incorporated to each side of the ball to vent out
excessive pressure caused by thermal expansion after engine shutdown.
A strainer is incorporated in the inlet of each ejector pump to prevent ingestion
of foreign objects. ENGINE SHUTOFF VALVE (SOV) ACTUATOR ASSEMBLIES
DC AUXILIARY BOOST PUMP CARTRIDGE There are two engine SOV actuator assemblies in the engine fuel feed sub-
subsystem. The actuator assemblies are installed in the engine SOVs,
There are two V DC auxiliary boost pump cartridges in the engine fuel feed outside the fuel tank. The actuator assemblies are electrically operated and
sub-subsystem. There is one V DC auxiliary boost pump cartridge installed control the open/closed positions of the SOVs. The ENG SHUTOFF switches,
in each wing tank collector box. The V DC auxiliary pump cartridge is installed on the FIRE extinguisher panel in the cockpit, operate the actuator
in the V DC auxiliary boost pump canister and can be removed and installed assemblies. Indication switches in the actuators provide feedback regarding
without defueling the fuel tank. They supply fuel to the engines for engine valve position. The EFCU monitors the status of the left engine SOV switches
start, or in the event of engine feed ejector pump failure. and transmits the data for the CAS display. The EFCU also monitors the
signals from the right engine SOV switches and transmits the data for the
The V DC auxiliary pump cartridges are centrifugal, wet-motor pumps that CAS display.
use pressurized fuel for cooling. They are brushless 28 V DC electronically
controlled motor supplied by the EMERGENCY BUS bar. The electronic FUEL TRANSFER VALVE
control is integral to the motor.
The fuel transfer valve is installed on the inside face of the front spar, in the
DC AUXILIARY BOOST PUMP CANISTER left tank, with its spindle passing through the spar to a separately removable
actuator located on the outside face. The fuel transfer valve is an electrically
There are two V DC auxiliary boost pump canisters in the engine fuel feed actuated ball valve that opens in the event of fuel load imbalance occurring
sub-subsystem. There is one V DC auxiliary boost pump canister installed in between wings (e.g. following engine failure). Lateral balance is maintained
each wing tank collector box, on the lower wing surface. They house the V by opening the fuel transfer valve by means of a switch on the fuel control
DC auxiliary boost pump cartridge. panel and allowing fuel to be transferred by gravity. The fuel transfer valve is
a ball valve and is controlled by the fuel transfer SOV actuator.
The V DC auxiliary pump cartridge and canister are designed not to exceed
200 °C (392 °F) external case temperature. The V DC auxiliary pump canister FUEL TRANSFER VALVE (SOV) ACTUATOR
has thermal fuses to prevent hazardous temperatures. A strainer is
incorporated to the inlet of the V DC auxiliary pump canister to prevent

EFFECTIVITY: ALL
NOTES: NOTES:

21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
The fuel transfer SOV actuator is installed on the fuel transfer valve, on the The check valve is an in-line swing check valve. The fittings at the ends are
outside face of the front spar in the left fuel tank. The actuator is electrically different to prevent reverse installation.
operated and controls the open/closed positions of the fuel transfer valve. A
pushbutton on the FUEL control panel in the cockpit operates the actuator. FUEL CONTROL PANEL
Indication switches in the actuator provide feedback regarding valve position.
The FUEL control panel is located on the main instrument panel in the cockpit.
The EFCU monitors the status of the switches and transmits the data for the
The FUEL control panel provides control of engine fuel feed and fuel transfer.
CAS display.
The two DC (Direct Current) pump switches and the XFR pushbutton on the
SCAVENGE EJECTOR PUMPS control panel are used to set the mode of operation for the V DC pump and

the fuel transfer valve.
There are two scavenge ejector pumps in the engine fuel feed sub-
subsystem. The ejector pumps are venturi-type pumps, with no moving parts, The default positions of the FUEL control panel are shown in the table below.
that draw fuel from the main tanks to the collector tanks when fed with motive
flow. The ejector pumps receive their motive flow from engine-driven fuel XFR Pushbutton OFF (Not Pushed)
pumps.
DC PUMP Switches AUTO
One scavenge pump is installed in each main fuel tank, forward of the
collector box, between ribs 1 and 2 and spars 1 and 2. These ejector pumps The figure ENGINE FEED SYSTEM - COMPONENT LOCATION provides
collect fuel from the wing tanks and transfer it to the collector tank. further data on the preceding text.
ENGINE FEED CHECK VALVES
There are four engine feed check valves in the engine fuel feed sub-
subsystem. Two of these check valves are installed in both engine feed lines,
downstream of the engine feed ejector pump. The other two check valves are
installed in both engine feed lines, downstream of the V DC auxiliary boost
pump. The check valves control the flow of fuel from the engine feed ejector
pumps to the engines. The check valves also prevent fuel flow from the V DC
auxiliary boost pumps in the wrong direction.
MOTIVE FLOW CHECK VALVES
There are two motive flow check valves in the engine fuel feed sub-
subsystem. A check valve is installed in each motive flow line, upstream of
the engine feed ejector pump. The check valves prevent excessive fuel loss
if the motive flow line is open due to failure or maintenance activity.

EFFECTIVITY: ALL
C
B
H

D
A
ZONES
522/622 B
530/630
G
F
E
A
ENGINE FEED SYSTEM - COMPONENT LOCATION

Sheet 1
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
NOTES: NOTES:

EFFECTIVITY: ALL
ENGINE
SHUTOFF ENGINE FEED
MOTIVE FLOW VALVE CHECK VALVE
CHECK VALVE

B D
C TYPICAL
TYPICAL
TYPICAL
SCAVENGE
EJECTOR
PUMP FUEL TRANSFER
VALVE
ENGINE FEED
VDC AUXILIARY
EJECTOR PUMP
BOOST PUMP
H G
TYPICAL TYPICAL F E
ENGINE FEED SYSTEM - COMPONENT LOCATION

Sheet 2
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
Operation The figure ENGINE FEED SYSTEM - BLOCK DIAGRAM provides further
ENGINE FUEL FEED OPERATION
With both engines and engine-driven motive flow pumps operating normally,
motive flow is supplied to the engine feed and scavenge ejector pumps. The
scavenge ejector pumps transfer fuel to the collector tanks to maintain them
with a correct fuel level even during uncoordinated maneuvers. The engine
feed ejector pumps supply fuel to the engines.

Pressure switches are installed in the engine feed lines (see AMM SDS
28-45-00/1). If a pressure switch senses that the fuel pressure is less than
approximately 41.4 kPa (6 psig), the FUEL 1(2) LO PRES caution message
shows on the PFD (Primary Flight Display), in the CAS display. If the V DC
PUMP switch is set to AUTO, the EFCU sends a discrete signal for the V DC
auxiliary pump to start on, and the FUEL 1(2) FEED FAULT advisory
message shows on the PFD, in the CAS display.
FUEL TRANSFER OPERATION
A fuel transfer function is provided to allow the fuel imbalance between the
left and right wing tanks to be lower than 140 kg.
If an imbalance of more than approximately 140 kg (308 lb) between the left
and right wing tanks occurs for a period longer than approximately 10
seconds, the FUEL IMBALANCE caution message shows on the PFD, in the
CAS display. The operator must then set the XFR pushbutton to OPEN to
initiate a fuel transfer. When the operator does that, the fuel transfer valve
opens and the lateral balance is achieved through gravity. Once the fuel
imbalance becomes approximately 60 kg (132 lb), the FUEL IMBALANCE
caution message goes out of view. When the fuel balance is achieved (fuel
imbalance is less than approximately 40 kg (88 lb)), the FUEL EQUAL
advisory message comes into view, warning the operator to stop the fuel
transfer. Then the operator must set the XFR switch to CLOSE.
If there is an engine failure, the fuel transfer function can be used to prevent
fuel imbalance.

EFFECTIVITY: ALL
RIB 1
RIB 2
RIB 2
FUEL
RIB 3
RIB 3
TRANSFER
RIGHT MAIN
RIB 9
LEFT MAIN
RIB 9
VALVE
RIB 12
RIB 12
TANK
RIB 13
TANK
RIB 13
RIB 14
RIB 14
DCM
RIB 15
RIB 15
SPAR I
SPAR I

DCM DCM
SPAR II SPAR II
COLLECTOR
TANK
SPAR III SPAR III

DCM
DCM
ENGINE 1 ENGINE 2
SURGE TANK SURGE TANK
SHUTOFF SHUTOFF
VALVE VALVE
LEGEND: PS PS
MP MP
SCAVENGE EJECTOR PUMP MOTIVE FLOW LINE
ENGINE FEED EJECTOR PUMP FUEL FEED LINE

ENGINE
ENGINE
DCM
DC AUXILIARY BOOST PUMP SCAVENGE/ TRANSFER LINE
DCM
CHECK VALVE
MP MOTIVE PUMP
COLLECTOR TANK VENT ORIFICE
ENGINE FEED SYSTEM - NORMAL ENGINE FEED OPERATION
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
NOTES: NOTES:

EFFECTIVITY: ALL
VDC VDC
AUXILIARY AUXILIARY
BOOST BOOST
PUMP 1 PUMP 2
ENG 1 FIRE
RELAY RELAY

ENGINE/
AIRFRAME
SHUTOFF
OFF OFF
PUSHBUTTON
ENGINE/ ENGINE/

CHANNEL 1 CHANNEL 2
UNIT UNITS (GIA 2)
LH TANK RH TANK
HI UNIT ARRAY HI
UNIT ARRAY
LOW LOW
ENGINE FEED SYSTEM - BLOCK DIAGRAM
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
INDICATING 28-40
Introduction
The indicating subsystem gives electrical fuel quantity, fuel low-level, and fuel
temperature indications and warnings to the crew.
The figure INDICATING - SCHEMATIC DIAGRAM provides further data on

the preceding text.

EFFECTIVITY: ALL
INDICATING 28-40
RIB 1
RIB 2
RIB 2
RIB 3
RIB 3
RIB 7
RIB 7
RIB 8
RIB 8
RIB 11
RIB 11
RIB 12
RIB 12

SPAR I
SPAR I
SPAR II SPAR II
SPAR III SPAR III
EFCU EFCU
TO AVIONICS ARINC 429 ARINC 429 TO AVIONICS
CH 1 CH 2
LEGEND:
TANK UNIT
FUEL TEMPERATURE SENSOR

(INSTALLED ON LEFT WING TANK ONLY)
INDICATING - SCHEMATIC DIAGRAM
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
INDICATING 28-40
General Description
The INDICATING includes these subsystems:
• ELECTRICAL FUEL QUANTITY (AMM SDS 28-41-00/1)

INDICATING
• FUEL TEMPERATURE INDICA- (AMM SDS 28-43-00/1)
TION SYSTEM
• FUEL LOW PRESSURE WARN- (AMM SDS 28-45-00/1)

ING SYSTEM
Some fuel indications and warnings are shown in the EICAS (Engine
Indication Crew Alert System) fuel indication field, on the CAS (Crew Alerting
System) display, and on the MFD (Multi-Function Display) fuel synoptic page.
Some fuel indicating failures are reported to and stored in the CMC (Central
Maintenance Computer).
The figure INDICATING - BLOCK DIAGRAM provides further data on the

preceding text.

EFFECTIVITY: ALL
INDICATING 28-40
87.8 TO 87.8
A B ATR
2.5 N1% 2.5

FUEL INDICATING
AREA
ITT C
IGN ____ ____ IGN

OFF OFF
RIGHT WING FUEL
55.1 N2% 55.1
QUANTITY (REF.)
OIL PRES PSI
OIL TEMP C
FUEL
FF PPH
FQ LB
XFR
TEMP XX C
ELEC CABIN
BATT1 0V
ALT
BATT2 0V
RATE
C LEFT WING
FUEL XXX LB XXX LB
SPDBRK DELTA-P
LFE
QUANTITY
OXY
(REF.)
LG FLAPS
CAS
LG LEVER DISAG
E1 FIREX FAIL TOTAL TOTAL
BLEED 2 FAIL FUEL XXXX LB DN
BLEED 1 FAIL QUANTITY TAKEOFF DATA SET
FUEL XFR FAIL (REF.) USED OAT -237 C
FUEL 2 SOV FAIL XXXX LB
FUEL 1 SOV FAIL ATR ON
E2 FADEC FAULT
BRK FAIL
OXY LO PRES FUEL
D−I WINGSTB FAIL QUANTITY EICAS
STALL ICE SPEED USED
ADS−AOA NOT AUTO (REF.) A
MFD
CAS WINDOW
(FUEL SYNOPTIC PAGE)
C B
INDICATING - DISPLAYS
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
INDICATING 28-40
NOTES: NOTES:

EFFECTIVITY: ALL
INDICATING 28-40
VDC VDC
AUXILIARY AUXILIARY
BOOST BOOST
PUMP 1 PUMP 2
ENG 1 FIRE
RELAY RELAY

ENGINE/
AIRFRAME
SHUTOFF
OFF OFF
PUSHBUTTON
ENGINE/ ENGINE/

CHANNEL 1 CHANNEL 2
UNIT UNITS (GIA 2)
LH TANK RH TANK
HI UNIT ARRAY HI
UNIT ARRAY
LOW LOW
INDICATING - BLOCK DIAGRAM
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
ELECTRICAL FUEL QUANTITY INDICATING 28-41
Introduction
The electrical fuel quantity indicating sub-subsystem gives indication of the

fuel quantity to the crew.
The figure ELECTRICAL FUEL QUANTITY INDICATING - LOCATION


EFFECTIVITY: ALL
RIB 1
RIB 2
RIB 2
RIB 3
RIB 3
RIB 7
RIB 7
RIB 8
RIB 8
RIB 11
RIB 11
RIB 12
RIB 12

2LH 2RH
SPAR I
SPAR I
SPAR II SPAR II
3LH 3RH
4LH 4RH
1LH 1RH
SPAR III SPAR III
EFCU EFCU
CH 1 CH 2
LEGEND:
TANK UNIT
ELECTRICAL FUEL QUANTITY INDICATING - LOCATION
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
General Description
The left wing fuel tank
The electrical fuel quantity indicating sub-subsystem has a dual-channel quantity inside tank is
EFCU (Electronic Fuel Control Unit), two sets of tank units (unit arrays), and small. There is less
two harness assemblies. Each unit array has 4 fuel quantity probes. The tank than 30 minutes of fuel
FUEL 1 LO LEVEL Caution (Amber)
units sense the fuel level and send this information to the related channel of remaining at cruising
the EFCU. The EFCU channels are fully segregated. The EFCU sends this speed. The low level
information to the MFD (Multi-Function Display) through an ARINC set point is equal to 90
(Aeronautical Radio Incorporated) 429 digital data bus. kg (200 lb).

The right wing fuel
Some fuel indicating failures are reported to and stored in the CMC (Central
tank quantity inside
Maintenance Computer).
tank is small. There is
Fuel quantities in each tank are shown on the EICAS (Engine Indication Crew less than 30 minutes
FUEL 2 LO LEVEL Caution (Amber)
Alert System) fuel indication field and on the MFD fuel synoptic page, where of fuel remaining at
the total and used fuel quantities are also shown. cruising speed. The
low level set point is
On the EICAS fuel indication field, the digits are provided green on black equal to 90 kg (200 lb).
background and turn black on amber background (equal to or less than 90
kg / 200 lb) or white on red background (equal to 0 kg / 0 lb) to indicate the There is a lateral fuel
different fuel quantity levels in each tank when it reaches the predetermined imbalance between
quantities. The EICAS fuel indication field also shows the total fuel quantity; FUEL IMBALANCE Caution (Amber) the left and right wing
in this case the digits are provided green on black background and turn black tanks equal to 140 kg
on amber background (equal to or less than 180 kg / 400 lb) or white on red (308 lb).
background (equal to 0 kg / 0 lb) to indicate the different fuel quantity levels
when it reaches the predetermined quantities. On the MFD digital fuel quantity and analogue fuel quantity, indicated by
means of a vertical bar with a level bug calibrated in percent, are provided for
The CAS (Crew Alerting System) messages are shown on the PFD (Primary each tank. Digital indication is also provided for TOTAL and USED tank
Flight Display) and on the MFD in reversionary mode. contents:
The CAS messages related to the electrical fuel quantity indicating sub- • Between 90 kg (200 lb) and full tank quantity, digital fuel quantity for each
subsystem are listed in the table below: tank is displayed in green. The analogue fuel quantity bar for each tank is
displayed in white with the indicating arrow in green. Between 180 kg (400
lb) and full tank quantity, total fuel quantity is displayed in green on black
background.
• Between 0 and 90 kg (0 and 200 lb) digital fuel quantity for each tank is
displayed in black on amber background and analogue fuel quantity for

EFFECTIVITY: ALL
NOTES: NOTES:

21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
each tank is displayed in amber. Between 0 and 180 kg (0 and 400 lb),
total fuel quantity is displayed in black on amber background.
• For fuel quantity equal to 0 kg (0 lb), digital fuel quantity for each tank is
displayed in white on red background and analogue fuel quantity for each
tank is displayed in red. For fuel quantity equal to 0 kg (0 lb), total fuel
quantity is displayed in white on red background.
The figure ELECTRICAL FUEL QUANTITY INDICATING - DISPLAYS



EFFECTIVITY: ALL
RIGHT TANK ANALOGUE

FUEL QUANTITY BAR RIGHT TANK
A B LEFT TANK ANALOGUE
FUEL QUANTITY
FUEL QUANTITY BAR
87.8 TO 87.8
ATR
LEFT XFR
2.5 N1% 2.5 TANK
FUEL

QUANTITY
XXX LB XXX LB
ITT C
IGN ____ ____ IGN
OFF OFF
TOTAL FUEL
55.1 N2% 55.1
QUANTITY
OIL PRES PSI
OIL TEMP C TOTAL
FUEL FUEL USED XXXX LB
FF PPH
RIGH
FQ LB
FUEL USED
LEFT FUEL XXXX LB
FLOW FLOW
TEMP XX C
ELEC CABIN
BATT1 0V
LEFT FUEL BATT2 0V
ALT
QUANTITY RATE RIGHT
SPDBRK DELTA-P FUEL
LFE
QUANTITY
OXY
MFD
TOTAL FUEL LG FLAPS
(FUEL SYNOPTIC PAGE)
QUANTITY
DN B
TAKEOFF DATA SET
OAT -237 C
ATR ON
EICAS
ELECTRICAL FUEL QUANTITY INDICATING - DISPLAYS
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
Components Each channel of the EFCU supplies these outputs:
ELECTRONIC FUEL CONDITIONING UNIT • Discrete outputs for low level warning.
The EFCU is a dual-channel microprocessor controlled unit that processes • Digital data to the aircraft through the ARINC 429 data bus.
fuel quantity data. The channels are the left and right fuel quantity processors.
Each processor channel receives fuel quantity data from the other via a serial The EFCU is installed in the Center LH (Left-Hand) Compartment.
data link internal to the EFCU.
FUEL QUANTITY PROBES

An independent 28 V DC electrical supply is provided for each of the two
There are 8 fuel quantity probes in the electrical fuel quantity indicating sub-
segregated channels within the EFCU. The left channel is powered by the
subsystem. The probes provide capacitance values to the EFCU for fuel
EMERGENCY BUS. The right channel is powered by the DC2 BUS.
quantity indication. The probes are composed of two concentric composite
The EFCU performs these functions: cylinders, a terminal block, and mounting brackets. The cylinders form the
capacitor elements. The inner cylinder is the high-impedance element and
• Performs continuous self-test (BIT (Built-in Test)). the outer cylinder is the low-impedance element. Changes in fuel level around
the probes cause changes in the capacitance of the probes.
• Monitors fuel quantity failures.
TANK-UNIT HARNESS ASSEMBLIES
• Provides excitation signals to the tank units.
There are two identical tank unit harness assemblies, one for each tank array.
• Receives return signals from the tank units. The harness assemblies use separate connectors at the fuel tank wall
interface for the Hi-Z (high impedance) and Lo-Z (low impedance) signal wires
• Conditions and supplies the signals to the displays (PFD and MFD).
to prevent stray capacitance. The terminal lugs at the tank units have 3 wires,
Fuel mass is continuously computed in both pounds and kilograms. The pilot each one carrying an integral captive screw on a ring tag, in which the
can view the fuel weight displayed on the EICAS fuel indication field or by harness/probe connection is crimped and soldered.
selecting the fuel synoptic page on the MFD in the units designated for the
The figure ELECTRICAL FUEL QUANTITY INDICATING - COMPONENT
aircraft.
Each processor within the EFCU also monitors the status of low level
information derived from fuel quantity data. Low level status is output in two
discrete signals.
Each channel of the EFCU receives these inputs:
• Capacitance inputs (return signals) from 4 tank units.
• Digital data from the aircraft through ARINC 429 data bus.

EFFECTIVITY: ALL
ZONES
531/541 B
631/641
B
A
B
RIB 1
B

RIB 2
RIB 7
RIB 8
A RIB 11
HI Z
CONNECTION
PROBE
C
LI
CONNECTION
C B
ELECTRICAL FUEL QUANTITY INDICATING - COMPONENT LOCATION

Sheet 1
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
NOTES: NOTES:

EFFECTIVITY: ALL

RIB 1
RIB 2
RIB 3
A
ZONES
530/541
630/641 RIB 7
RIB 8
RIB 10
RIB 11
TANK−UNIT
HARNESS ASSENBLY
A

Sheet 2
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
NOTES: NOTES:

EFFECTIVITY: ALL
ZONES
241
242
A

B
EFCU

Sheet 3
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
Operation
LEFT TANK FUEL QUANTITY INDICATING OPERATION
When the aircraft is energized, EFCU channel 1 receives 28 V DC through

the V DC EMERGENCY BUS. The EFCU excites the left tank unit array and
provides the signals of the amount of fuel remaining in the left tank. The EFCU
sends these signals to the PFD, in the CAS display, and fuel synoptic page
on the MFD.

The EFCU monitors the tank unit array and sends fault messages to the CMC.
The EFCU receives low level signal from the left tank fuel quantity probes
and sends the discrete signals for low level warning.
RIGHT TANK FUEL QUANTITY INDICATING OPERATION
When the aircraft is energized, the EFCU receives 28 V DC through the V

DC DC2 BUS. The EFCU excites the right tank unit array and provides the
signals of the amount of fuel remaining in the right tank. The EFCU sends
these signals to be displayed on the PFD, in the CAS display, and on the fuel
synoptic page on the MFD.
The EFCU monitors the tank unit array and sends fault messages to the CMC.
The EFCU receives low level signal from the right tank fuel quantity probes
and sends the discrete signals for low level warning.
The figure ELECTRICAL FUEL QUANTITY INDICATING - BLOCK


EFFECTIVITY: ALL
VDC VDC
AUXILIARY AUXILIARY
BOOST BOOST
PUMP 1 PUMP 2
ENG 1 FIRE
RELAY RELAY

ENGINE/
AIRFRAME
SHUTOFF
OFF OFF
PUSHBUTTON
ENGINE/ ENGINE/

CHANNEL 1 CHANNEL 2
UNIT UNITS (GIA 2)
LH TANK RH TANK
HI UNIT ARRAY HI
UNIT ARRAY
LOW LOW
ELECTRICAL FUEL QUANTITY INDICATING - BLOCK DIAGRAM
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
FUEL TEMPERATURE INDICATION SYSTEM 28-43
Introduction
The fuel temperature indicating sub-subsystem gives fuel temperature

indication to the crew.
General Description
The fuel temperature indicating sub-subsystem has a temperature sensor in

the left collector tank. The EFCU (Electronic Fuel Control Unit) monitors the

resistance value of the temperature sensor and provides the fuel temperature
to be displayed on the MFD (Multi-Function Display).
A cockpit display of fuel temperature is provided on the EICAS (Engine

Indication Crew Alert System) fuel indicating field. The fuel temperature signal
is sent to Channel 1 of the EFCU. In the event of sensor failure, a red "X" is
shown in place of the temperature indication.
The temperature value is shown in green if the fuel temperature is more than
−37 °C (−34.6 °F) and less than 80 °C (176 °F). The temperature value is
also shown in black (amber background) if the fuel temperature is less than
−37 °C (−34.6 °F) or more than 80 °C (176 °F).
The figure FUEL TEMPERATURE INDICATION SYSTEM - LOCATION


EFFECTIVITY: ALL
RIB 1
RIB 2
RIB 2
RIB 3
RIB 3
RIB 7
RIB 7
RIB 8
RIB 8
RIB 11
RIB 11
RIB 12
RIB 12

SPAR I
SPAR I
SPAR II SPAR II
SPAR III SPAR III
EFCU EFCU
CH 1 CH 2
LEGEND:

(INSTALLED ON LEFT WING TANK ONLY)
FUEL TEMPERATURE INDICATION SYSTEM - LOCATION
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
Components
The fuel temperature sensor is located at spar III, outside the left collector
tank and can be accessed through the wing-to-fuselage fairing access
panels.
A PT-500 three-wire configuration is used for the temperature sensor. It


consists of a platinum RTD (Resistance Temperature Detector). The
temperature sensor can be removed and installed without opening the fuel
tank access panels; however, it is necessary to defuel the tanks to remove
the temperature sensor.
The figure FUEL TEMPERATURE INDICATION SYSTEM - COMPONENT


EFFECTIVITY: ALL
B

A
ZONES
522
FUEL
TEMPERATURE
SENSOR
B
FUEL TEMPERATURE INDICATION SYSTEM - COMPONENT LOCATION
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
Operation
FUEL TEMPERATURE INDICATING OPERATION
When the fuel temperature is less than −37 °C (−34.6 °F) or more than 80 °C
(176 °F), the temperature value is shown in black in an amber background
on the EICAS.
If this condition occurs, the crew must:


• Lower the aircraft altitude.
• Increase the airspeed.
• Monitor the fuel temperature.
If this condition occurs prior to takeoff, the aircraft cannot be dispatched

unless it has been fueled with fuel (Jet A-1) at temperatures between -37 °C
(-34.6 °F) and 80 °C (176 °F).
The figure FUEL TEMPERATURE INDICATION SYSTEM - DISPLAYS


EFFECTIVITY: ALL
A 87.8 TO 87.8
ATR
2.5 N1% 2.5


ITT C
IGN ____ ____ IGN
OFF OFF
55.1 N2% 55.1

OIL PRES PSI
TEMP C
FUEL
FF PPH
FQ LB
FUEL TEMPERATURE TEMP XX C

ELEC CABIN
INDICATION
BATT1 0V
ALT
BATT2 0V
RATE
SPDBRK DELTA-P
LFE
OXY
LG FLAPS
DN
TAKEOFF DATA SET
OAT -237 C
ATR ON
EICAS
FUEL TEMPERATURE INDICATION SYSTEM - DISPLAYS
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
FUEL LOW PRESSURE WARNING SYSTEM 28-45
Introduction
The low-pressure warning sub-subsystem monitors the fuel pressure in the

engine feed lines. This sub-subsystem gives indication of low fuel pressure
to the crew.
The figure FUEL LOW PRESSURE WARNING SYSTEM - LOCATION


EFFECTIVITY: ALL
RIB 1
RIB 2
RIB 2
RIB 3
RIB 3
RIB 9
RIB 9
RIB 12
RIB 12
RIB 13
RIB 13
RIB 14
RIB 14
DCM
RIB 15
RIB 15
SPAR I
SPAR I

DCM DCM
SPAR II SPAR II
SPAR III SPAR III

DCM
DCM
PS PS
ENGINE
ENGINE
LEGEND:
MOTIVE FLOW LINE
FUEL FEED LINE
FUEL LOW PRESSURE WARNING SYSTEM - LOCATION
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
General Description Right fuel pressure

switch indicates the
The low pressure warning sub-subsystem has two low pressure switches to pressure is not low
monitor the engine feed lines. when all fuel pumps
FUEL 2 PSW FAIL Advisory (White) are off; or right pres-
One low pressure switch is installed in the left engine feed line, downstream
sure switch did not de-
of the left engine SOV (Shutoff Valve). The other low pressure switch is
tect high pressure
installed in the right engine feed line, downstream of the right engine SOV.
when DC pump 2 was
commanded to ON.

Each engine low pressure switch monitors the related feed line. If the fuel
pressure decreases below 41.4 kPa (6 psi), each pressure switch sends a
signal to both EFCU (Electronic Fuel Control Unit) channels, which send a The figure FUEL LOW PRESSURE WARNING SYSTEM - COMPONENT
signal to cause the automatic operation of the applicable V DC auxiliary boost LOCATION provides further data on the preceding text.
fuel pump. The DC PUMP switches set at AUTO enables the automatic
operation of the V DC auxiliary pumps (see AMM SDS 28-21-00/1).
After receiving the left engine 1 and left engine 2 fuel low pressure signals,
the EFCU sends them to the MFD (Multi-Function Display).
The CAS (Crew Alerting System) message related to the fuel low pressure
warning sub-subsystem is given in the table below:
The fuel pressure of

FUEL 1 LO PRES Caution (Amber)
the left engine is low.
The fuel pressure of
FUEL 2 LO PRES Caution (Amber)
the right engine is low.
Left fuel pressure
switch indicates the
pressure is not low
when all fuel pumps
FUEL 1 PSW FAIL Advisory (White) are off; or left pressure
switch did not detect
high pressure when
DC pump 1 was com-
manded to ON.

EFFECTIVITY: ALL

A B
ZONES
522
622
C A
ENGINE 2 FUEL
FEED PRESS SW
ENGINE 1 FUEL ENGINE 2 FUEL
FEED PRESS SW ENGINE 1 FEED LINE
MOTIVE FLOW
LINE
ENGINE 1 FUEL C
FEED LINE
B ENGINE 2
MOTIVE FLOW
LINE
FUEL LOW PRESSURE WARNING SYSTEM - COMPONENT LOCATION
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
Components
LOW PRESSURE SWITCH
A pressure switch is installed in the engine feed line, downstream of each

engine SOV.
This pressure switch monitors the pressure in the feed line. When the
pressure is lower than 41.4 kPa (6 psig), the pressure switch energizes the

applicable remaining pump and sends a signal to the PFD (Primary Flight
Display). When this occurs, the CAS display will show the FUEL 1(2) LO
PRES caution indication and the V DC auxiliary boost pump starts its
operation.
The automatic operation of the V DC auxiliary pumps occurs when the DC

PUMP switches are set to AUTO (see AMM SDS 28-21-00/1).
The figure FUEL LOW PRESSURE WARNING SYSTEM - DISPLAYS


EFFECTIVITY: ALL
A

CAS
FUEL 1 LO PRESS
FUEL 2 LO PRESS
FUEL 1 PSW FAIL
FUEL 2 PSW FAIL
CAS WINDOW
A
FUEL LOW PRESSURE WARNING SYSTEM - DISPLAYS
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
Operation
LOW PRESSURE WARNING
If the fuel pressure is too low in an engine feed line:
• The caution message FUEL 1(2) LO PRES shows on the CAS display.
• The V DC applicable auxiliary boost pump is energized.


• The fuel pressure increases in comparison with the operating pressure
(41.4 kPa or 6 psig).
The automatic operation of the V DC auxiliary pumps occur when the DC

PUMP switches are set to AUTO (see AMM SDS 28-21-00/1).
The figure FUEL LOW PRESSURE WARNING SYSTEM - BLOCK


EFFECTIVITY: ALL
VDC VDC
AUXILIARY AUXILIARY
BOOST BOOST
PUMP 1 PUMP 2
ENG 1 FIRE
RELAY RELAY

ENGINE/
AIRFRAME
SHUTOFF
OFF OFF
PUSHBUTTON
ENGINE/ ENGINE/

CHANNEL 1 CHANNEL 2
UNIT UNITS (GIA 2)
LH TANK RH TANK
HI UNIT ARRAY HI
UNIT ARRAY
LOW LOW
FUEL LOW PRESSURE WARNING SYSTEM - BLOCK DIAGRAM
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80



CHAPTER 71 - POWERPLANT
SECTION TITLE PAGE

71-00 POWERPLANT 140

71-10 COWLING 148
71-20 MOUNTS 152
71-30 FIRESEAL 154
71-50 ELECTRICAL HARNESS 156
71-60 AIR INLET 162
71-70 ENGINE DRAINS 164
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
POWERPLANT 71-00
Introduction is the need for safety. This has been achieved by providing redundancy and
independence into the control system.
The powerplant system is basically composed of two pylon-mounted Pratt &
Whitney PW617F turbofan engines on the rear fuselage. A twin-channel Full Authority Digital Electronic Control (FADEC) controls the
engine and regulates its operation in response to inputs from the pilot,
The powerplant provides thrust for the aircraft, as well as pneumatic and airframe, and engine mounted sensors.
electrical power.
The powerplant indications are displayed on the EICAS (Engine Indication
The engines are controlled from the cockpit control stand and powerplant Crew Alert System) on the left stripe of the center MFD (Multi-Function

control panel through the FADEC (Full Authority Digital Engine Control). Each Display) unit of the cockpit panel. The powerplant indications can also be
engine is controlled and monitored by two FADEC channels. When one shown on the PFD (Primary Flight Display) in reversionary mode. The CAS
channel is in control, the other is in standby mode. (Crew Alerting System) messages are shown on the CAS window on the PFD
and on the MFD in reversionary mode.
General Description
The color scheme adopted for the propulsion system warning, caution, and
The POWERPLANT includes these subsystems: advisory indications are shown below:
• COWLING (AMM SDS 71-10-00/1) • Red, for warning lights - lights indicating a hazard which may require
• MOUNTS (AMM SDS 71-20-00/1) immediate corrective action.
• FIRESEAL (AMM SDS 71-30-00/1)
• ELECTRICAL HARNESS (AMM SDS 71-50-00/1) • Amber, for caution lights - lights indicating the possible need for future
• AIR INLET (AMM SDS 71-60-00/1) corrective action.
• ENGINE DRAINS (AMM SDS 71-70-00/1)
• Cyan, for advisory lights.
The PW617F is a two-spool turbofan engine with a full length annular bypass • White, for status lights.
duct. The engine is designed, developed, and manufactured by the Pratt & • Green, for safe operation lights.
Whitney Company.
Rotary and pushbutton switches for ignition, and engine START/STOP are
The PW617F control system is a computer-based electronic engine control located on the cockpit panel.
system. It is composed of a twin-channel FADEC, a FMU (Fuel Metering
Unit), PMA (Permanent Magnet Alternator), engine sensors, a BVA (Bleed The FADEC is able to transfer control from one channel to another in the
Valve Actuator), an ignition system for each engine, TCQ (Thrust Control event of a failure on that channel that results in loss of functionality. Control
Quadrant) and engine cockpit switches (ignition and start/stop switches). of the engine is maintained in the presence of multiple faults through a
hierarchical scheme that maintains the most fit channel in control as long as
The system controls the engine in response to thrust command inputs from possible.
the aircraft and provides information to the aircraft for cockpit indication,
maintenance reporting and engine condition monitoring. Due to the criticality The channels are designated Channel A and Channel B. Identical software
of the functions, the main aspect of the design of the PW617F FADEC system is loaded into each channel. Only one channel can be in control of the output
EFFECTIVITY: ALL
POWERPLANT 71-00
NOTES: NOTES:

21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
POWERPLANT 71-00
devices at a time. The software in each channel reads the other channel’s electrical current return under wiring faulty conditions, the engine
health status and determines which is the healthier channel to remain in compartment is electrically bonded to the airframe structure.
control. During each start, the channel in control is switched to confirm that
the standby channel is capable of controlling and is free from faults only AIR INLET (71-60)
detectable by having control of the engine. This process reduces the
The main function of the air inlet is to supply proper engine air flow, assure
probability of dormant failures.
minimum air temperature increase and reduce the total pressure loss as well
The PW617F control system is composed of the following main components: as decrease drag in different operating conditions.
ENGINE DRAINS (71-70)


• FADEC
• EDCU (Engine Data Collector Unit) The drain lines collect fuel, and oil, from some points of the powerplant and
discharge the fluids overboard.
• PMA
• BVA Operation
• Engine Sensors All the interfaces between the cockpit and the engine nacelle are electrically
done. The control stand has two thrust levers, one for each engine thrust
Components control. The powerplant panel has dedicated switches to select the IGNITION
system (OFF/AUTO/ON), and engine START/STOP.
COWLING (71-10)
The engine indications are displayed on the EICAS.
The main function of the engine cowlings is to permit a smooth, undisturbed
air flow around the engine, and provide a protective covering for the engine Training Information Point
and its components.
The powerplant, owing to its importance for the aircraft, is the system that
MOUNTS (71-20) needs most attention and care.
The main function of the engine mount system is to attach the engine to the When the aircraft is not in service, the protection devices must be installed
nacelle pylons, and absorb noise and vibration. on the engine air inlet.
FIRESEAL (71-30) Over the aircraft operational life, the powerplant must be constantly checked,
independently of the scheduled inspections.
The engine compartment inside the nacelle is a single fire zone.
Whenever possible, the aircraft should be visually inspected for:
ELECTRICAL HARNESS (71-50)
• Fluid leak evidence.
The powerplant electrical harness links the engine accessories and the
aircraft systems. To prevent electrostatic discharges, lightning current, or • Cowling conditions.

EFFECTIVITY: ALL
POWERPLANT 71-00
NOTES: NOTES:

21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
POWERPLANT 71-00
• Engine air inlet conditions.
• Engine exhaust duct general conditions.
• Clogged drains.
• Clogged engine cooling air inlets and outlets.
The figure POWERPLANT - ENGINE CONTROLS AND OPERATING

INTERFACES provides further data on the preceding text.

EFFECTIVITY: ALL
POWERPLANT 71-00
CAS
A
B 87.8 TO
ATR
87.8
2.5 N1% 2.5


ITT C
IGN ____ ____ IGN
C CAS WINDOW 55.1 N2% 55.1

OIL PRES PSI
A OIL TEMP C
FUEL
FF KGH
FQ KG
TEMP XX C
ENG FIRE EXTINGUISHER TRIM ELEC CABIN
D SHUTOFF 1
BOTTLE
SHUTOFF 2
YAW BATT1 0V
ALT
TO/GA DISCH
LEFT RIGHT BATT2 0V
RATE
SWITCH ROLL SPDBRK DELTA-P
OFF LWD RWD
LFE
ENG START/STOP OXY

RUN RUN LG FLAPS
PITCH BKP
DN
TO/GA
DN
SWITCH
UP
1 2 TAKEOFF DATA SET
ENG IGNITION MODE
THROTTLE + OAT -237 C
ON BKP
LEVERS AUTO
ATR ON
OFF OFF
1 2
EICAS DISPLAY
FIRE/ENG/TRIM
THRUST CONTROL QUADRANT CONTROL PANEL B
D C
POWERPLANT - ENGINE CONTROLS AND OPERATING INTERFACES

Sheet 1
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
POWERPLANT 71-00
NOTES: NOTES:

EFFECTIVITY: ALL
POWERPLANT 71-00
EICAS
87.5 TO 87.5
ATR AIRFRAME AIR DATA REMOTE ENGINE
− WOW
(AVIONICS) FADECs
− FCV CLOSE − ENGINE POSITION
ELECTRICAL AMB. DATA
77.5 N1% 27.4 COMMAND − AIRCRAFT ID
POWER SUPPLY (MACH, P o )
− IGN A ON − MAINT. FAULT RESET
(FADEC, IGNITION,
− IGN B ON − TEST MODE ENABLE
TT0 HEATERS)
−TT0 HEATER INTERNACELLE X TALK (CAN BUS)
IGN IGN
544 ITT C 350
__ __ 28VDC
ENG START/STOP
N2%
RUN RUN

OIL PRES PSI STOP START STOP START
OIL TEMP C
− N1 RED LINE
FADEC CHANNEL A START/ STOP/ IGN 1 2
ENG IGNITION
− N2 RED LINE ON
− ITT RED LINE AUTO
ARINC 429 − N1 MAX OFF

1 2
− THRUST RATING
X TALK
INTEGRATED ARINC 429
AVIONICS UNIT − N1 TARGET
ARINC 429 − N1 REQUEST
STOP
GSD FOR CHANNELS A
GIA FOR CHANNELS B ARINC 429 − N1
− N2 TLA
− ITT RVDTs
− CMC
TT0 HEATER 28VDC
IGNITION 28 VDC
− FUEL FILTER IMP. BYPASS

FADEC CHANNEL B
ENGINE/AIRFRAME ENGINE/AIRFRAME
− OIL PRESSURE UNIT
UNIT
− OIL TEMPERATURE (GEA)
GEA 1 − ENGINE LH
GEA 2 − ENGINE RH
− OIL FILTER IMP. BYPASS
− CHIP DETECTOR
FUEL FLOW COMMAND
BLEED VALVE COMMAND (BVA)

STATIC PRESSURE
ELEC. POWER SUPPLY (PMA) EDCU
SENSOR
SHUTDOWN SOLENOID
ENG SENSORS (N1, N2, T6)
TT0
POWERPLANT - ENGINE CONTROLS AND OPERATING INTERFACES

Sheet 2
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
COWLING 71-10
Introduction • BVA drain line
The main function of the engine cowlings is to permit a smooth, undisturbed • GGC
air flow around the engine, and provide a protective covering for the engine
and its components. • Exciter Box
The engine cowling is composed of an upper mid cowl, a lower mid cowl, and • N1 (Fan Rotor Speed) sensor
a titanium apron.
• TT0 (Inlet Total Temperature) sensor

General Description
• Rear Mount
The cowling streamlined surface minimizes drag and gives a protective
• Front Mounts
enclosure for the engine and its accessories. A drain mast and a GGC (Gas
Generator Case) drain hole in the engine lower mid cowl receive all engine ENGINE LOWER MID COWL
drain lines. There are also two additional drain holes in the lowest area of the
engine lower mid cowl. For each engine lower mid cowl, a ground service/ The engine lower mid cowl rear edge is attached to the engine aft body
inspection quick-access door is available for oil tank level check/ interface with latches, whereas its forward and inboard/outboard edges are
replenishment and a dedicated access door is available for inspection of the attached to the engine air intake, apron and engine lower mid cowl with
oil filter mechanical pop up impending bypass indicator. Full access to the screws. The engine lower mid cowl permits access to the components below:
engine and aircraft systems installed in the engine compartment is available
through removal of the upper and lower mid cowls. The cowling also gives • FMU (Fuel Metering Unit)
lightning and fire protection for the engine compartment, ventilation for the
• AGB (Accessory Gearbox)
engine compartment, and outlet for the waste engine fluid.
• FDV (Flow Divider / Shutoff Valve)
Components
• Starter/Generator
ENGINE UPPER MID COWL
• Chip Detector
The engine upper mid cowl rear edge is attached to the engine aft body
interface with latches, whereas its forward and inboard edges are attached • Front Mounts
to the engine air intake and apron with screws. The engine upper mid cowl
permits access to the components below: • Rear Mount
• EDCU (Engine Data Collector Unit) • Oil Filter
• ACOC (Air-Cooled Oil Cooler) • Fuel Pump
• BVA (Bleed Valve Actuator) • Fuel Filter

EFFECTIVITY: ALL
COWLING 71-10
NOTES: NOTES:

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EFFECTIVITY: ALL
COWLING 71-10
• Engine Harness
• Ignitors
• Oil Tank Drain
• Oil Pump
• Engine Line Drains


The nacelle drain system consists of drain holes and drain mast collector at
the lowest part of the engine lower mid cowl compartment to prevent any
flammable fluid accumulation inside the fire zone and eliminate any fluid
puddle in this zone.
A quick-access door located on the engine lower mid cowl left side gives
access to the oil level sightglass.
During removal/installation of the engine cowling, be careful not to cause

damage to the engine components.
The figure COWLING - COMPONENT LOCATION provides further data on

the preceding text.

EFFECTIVITY: ALL
COWLING 71-10
NACELE
APRON
OIL INPENDING
BYPASS POP UP
DOOR

STARTER
GENERATOR ENGINE
VENT OUTLET COMPARTMENT
AIR INLET
DRAIN HOLE ENGINE UPPER
MID COWL
DRAIN
MAST
DRAIN HOLE
VENT OUTLET
GRILLE
GAS GENERATOR
CASE DRAIN HOLE
ENGINE LOWER
MID COWL
ENGINE OIL
SERVICE
ACCESS
COWLING - COMPONENT LOCATION
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EFFECTIVITY: ALL
MOUNTS 71-20
Introduction engine cowlings are opened, even if partially open, a visual inspection be
made in the compartment, checking for general conditions and cleanliness.
The main function of the engine mount system is to attach the engine to the It is also very important to check for possible corrosion points, cracks, and
nacelle pylons, and absorb noise and vibration. ruptures.
General Description The figure MOUNTS - COMPONENT LOCATION provides further data on
the preceding text.
The engine uses a two-plane three-point mount system:
• The front mounts are attached to the engine and to the yoke of the pylon.

• The rear mount is attached to the pylon and to the aft mount bracket
located in the engine.
The engine front plane in the front frame casing has four mounting pads.
The mounting pads are symmetrically spaced in relation to the vertical
centerline (two pads per side) to permit left or right fuselage pylon
installation.
The aft mount bracket is installed on the left or right side of the engine,
according to the buildup configuration of the engine (left or right).
Components
FRONT MOUNTS
The forward mount is an elastomeric soft mount designed to react to the

thrust, torque, lateral and vertical forces and it consists of bearings, bolts and
a yoke assembly necessary to transfer loads between the engine front mount
pads and the pylon.
REAR MOUNT
The rear mount is an elastomeric soft mount, which attaches to the engine
outer bypass duct flange and is designed to react vertical and lateral forces.
The rear mount consists of two links, end fitting, bearings, and bolts.
Owing to the importance of the engine mount and vibration-dampener

assembly in the aircraft operation, it is recommended that, whenever the

EFFECTIVITY: ALL
MOUNTS 71-20
ENGINE FORWARD
MOUNT

A PYLON YOKE
B (REF.)
PYLON
(REF.)
A
TYPICAL
AFT MOUNT
BRACKET ENGINE AFT
MOUNT
B
MOUNTS - COMPONENT LOCATION
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EFFECTIVITY: ALL
FIRESEAL 71-30
Introduction
The engine compartment inside the nacelle is a single fire zone.
General Description
The air inlet module rear wall and the exhaust module front wall have fireseals
that isolate the engine fire zone. It does not include the pylon firewall, which
is described in AMM SDS 54-52-00/1.

The front and rear fireseals, the pylon firewall, and the engine upper and lower
mid cowls are the limits of the engine and the accessory fire zone.
Drainage and ventilation are provided for the nacelle engine compartment.
Electrically controlled shutoff means are available for fuel, hydraulic, and
bleed air lines crossing the nacelle to the pylon firewall. Full segregation
between flammable fluid carrying lines and bleed ducts/electrical harnesses
is ensured inside the pylon region.
The figure FIRESEAL - COMPONENT LOCATION provides further data on

the preceding text.

EFFECTIVITY: ALL
FIRESEAL 71-30
EXHAUST
(REF.)

A
FIRE SEAL
B
AIR INLET
(REF.)
FIRE SEAL
FIRESEAL - COMPONENT LOCATION
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EFFECTIVITY: ALL
ELECTRICAL HARNESS 71-50
Introduction The figure ELECTRICAL HARNESS - COMPONENT LOCATION provides

further data on the preceding text.
The powerplant electrical harness links the engine accessories and the
aircraft systems. To prevent electrostatic discharges, lightning current, or
electrical current return under wiring faulty conditions, the engine
compartment is electrically bonded to the airframe structure.
General Description

The electrical system distributes the power required by the aircraft and engine
components. The electrical system also transmits control and indicating
signals between the engine and aircraft. The engine harnesses are routed
through the pylon.
Components
HARNESS
Most of the powerplant electrical harnesses are designed for disconnection

at the pylon firewall by means of connectors. These connectors have visual
locking advisory (colored neck) and do not require safety-wiring. The
generator electrical power harnesses have no connectors (at the pylon
firewall), but they can be disconnected at the generator terminals (two per
engine).
BONDING STRAPS
The engine compartment is electrically bonded to the airframe structure. The

bonding straps give an adequate path for lightning current and also for
electrical current return under wiring faulty conditions. It also prevents the
buildup of electrostatic charges between the engine and airframe structure.
The engine compartment has four bonding straps located as follows:
• 2 at the forward pylon.
• 2 at the rearward pylon.
All these bonding straps link the pylon firewall to the engine.

EFFECTIVITY: ALL
TTO HARNESS
NAI PRESSURE

IGNITER FADEC
CABLE CHANNEL B TRANSDUCER
HARNESS SENSOR
HARNESS
EMISSIVE
HARNESS
FADEC
CHANNEL A
HARNESS
SUSCEPTIBLE
HARNESS
MOPT
HARNESS
FLOWMETER
HARNESS
FUEL IMPENDING
BYPASS HARNESS
ELECTRICAL HARNESS - COMPONENT LOCATION

Sheet 1
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EFFECTIVITY: ALL
NOTES: NOTES:

EFFECTIVITY: ALL
BVA
HARNESS
IGNITER
CABLE

TO
AIRCRAFT
STARTER
GENERATOR
POWER CABLE
STARTER GENERATOR
EXCITATION HARNESS FMU HARNESS

Sheet 2
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EFFECTIVITY: ALL
NOTES: NOTES:

EFFECTIVITY: ALL
B

A
BONDING
STRAP A
BONDING
STRAP
B
BONDING
C STRAPS

Sheet 3
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EFFECTIVITY: ALL
AIR INLET 71-60
Introduction
The main function of the air inlet is to supply proper engine air flow, assure
minimum air temperature increase and reduce the total pressure loss as well
as decrease drag in different operating conditions.
General Description
The air inlet is installed on the forward face of the engine fan case. As it is a

part of the engine nacelle, the air inlet must also assist in minimizing the
nacelle drag and noise. Therefore, its function is to give a smooth
aerodynamic surface for airflow into and around the engine.
To remove or prevent ice formation around the engine inlet, the air inlet has
an engine ice protection system. The system is supplied with bleed air from
the related engine. A valve adjusts the pressure of the bleed air for the piccolo
tube. The air inlet also contains one air intake which supplies cooling air to
the starter/generator.
Components
The air inlet assembly contains: lip skin, forward bulkhead, starter/generator
air inlet, piccolo tube and a barrel with flange. The lip skin is "D" shaped in its
cross-section and manufactured from a one-piece heat-resistant aluminum
alloy. The forward bulkhead section is made of corrosion and fire resistant
material.
The air inlet must be protected with proper covers when the airplane is not
operating and parked at seaside areas or dust concentration regions.
The use of protective covers prevents any foreign objects or dust deposits
from being formed in the air inlet, which could cause damage to the engine.
The figure AIR INLET - LOCATION provides further data on the preceding
text.

EFFECTIVITY: ALL
AIR INLET 71-60
BARREL
AIR INLET

FORWARD
BULKHEAD
PICCOLO
TUBE
A
STARTER/GENERATOR
B AIR INLET
FLANGE
LIP SKIN
B
A
AIR INLET - LOCATION
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EFFECTIVITY: ALL
ENGINE DRAINS 71-70
Introduction Therefore, periodic visual inspections must be performed to make sure that
all drain points are unplugged
The drain lines collect fuel, and oil, from some points of the powerplant and
discharge the fluids overboard. The figure ENGINE DRAINS - COMPONENT LOCATION provides further
General Description
The drain points of the BVA (Bleed Valve Actuator), FMU (Fuel Metering Unit)
and starter/generator pad are connected through of tubing lines to a drain

outlet mast located at the lowest region on the lower mid cowl. The drain lines
are routed isolated from each other.
The drain point of GGC (Gas Generator Case) is collected through of tubing
line. and discharged overboard. The fluid drained is discharged overboard
through an outlet on the engine lower mid cowl. In addition, the engine cowling
door, engine exhaust duct, and engine air inlet have drain holes in their lowest
areas to avoid any fluid.
In addition, the engine lower mid cowl have drain holes in their lowest areas
to avoid any fluid accumulation.
Components
The drain system consists of lines and openings that convey overboard any
waste fluids from the engine, accessories, and nacelle.
The components that are connected to the drain system are as follows:
• BVA
• FMU
• GGC
• Starter/generator pad
All drains, besides their important function during the powerplant operation,
help in the identification of possible leakage in systems and components.

EFFECTIVITY: ALL
ENGINE DRAINS 71-70

STARTER/GENERATOR
PAD DRAIN
BLEED VALVE
ACTUATOR DRAIN
FUEL METERING
UNIT DRAIN
GAS GENERATOR
CASE DRAIN
A
ENGINE DRAINS - COMPONENT LOCATION
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CHAPTER 72 - ENGINE
SECTION TITLE PAGE

72-00 ENGINE 168

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EFFECTIVITY: ALL
ENGINE 72-00
Introduction • Engine Control System
The PW617F-E series engine is a high-bypass, two-spool, axial-flow, Components

turbofan engine.
For additional information about the engine components, refer to the last
General Description revision of PW617F Engine Maintenance Manual PN 3072162 (EMM TASK
72-00-00-800-801/91).
The PW617F-E engine is a two-spool turbofan engine with a full length
annular bypass duct. A concentric shaft system supports the LP (Low Training Information Points

Pressure) and HP (High Pressure) rotors. The inner LP shaft supports the LP
compressor (fan) which is driven by a single stage LP turbine. The outer HP Make sure that there are no persons near the engine air intake and exhaust
shaft system is mechanically independent of the LP shaft and supports a areas when you operate the engine. This will prevent injury to persons.
single mixed flow stage and one centrifugal stage HP compressor driven by
Do not touch the exhaust duct and engine components until they are cool.
a single-stage HP turbine. Thrust and roller anti-friction bearings provide
The temperature can stay hot for a long time after the engine stops.
support on each shaft.
Make sure that the engine harnesses are correctly attached and there is no
A twin channel FADEC (Full Authority Digital Engine Control) controls the
chafing against the engine components, cowlings or pylon structure. Such a
engine and regulates its operation in response to inputs from the pilot and
chafing can cause serious problems to the correct engine operation.
engine mounted sensors. For more details regarding the engine control
system, see AMM SDS 73-00-00/1. The engine ignition system has high energy. This makes the system a
dangerous source of electrical shock.
The PW617F-E engine is divided into 10 modules as follows:
Do not operate the engine near flammable material or fuel vent. Explosion
• LP Compressor (Fan)
can occur.
• HP Compressor
Before you start the engine, make sure that there are no objects near the
• Combustor and Diffuser Case engine air intake. These objects can go into the engine and cause damage
to it.
• HP Turbine
The figure ENGINE - COMPONENT LOCATION provides further data on the
• LP Turbine preceding text.
• Monocase
• Accessory Gearbox, Bearings, LP and HP Shafts
• Bypass Ducting and Externals
• External Accessories

EFFECTIVITY: ALL
ENGINE 72-00
AIR COOLER
OIL COOLER
(ACOC)
BLEED VALVE
ACTUATOR
(BVA)
IGNITION
EXCITER

ENGINE DATA
COLLECTOR UNIT
(EDCU)
T1 IGNITION
SENSOR CABLE
IGNITER
FAN SPINNER
FRONT
MOUNTS
PADS
FMU ASSEMBLY
STARTER/
GENERATOR OIL SIGHT GLASS
(LH ENGINE)
OIL FILLER
NECK
ENGINE - COMPONENT LOCATION
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80



CHAPTER 73 - ENGINE FUEL AND CONTROL
SECTION TITLE PAGE

73-00 ENGINE FUEL AND CONTROL 172

73-10 ENGINE FUEL DISTRIBUTION 174
73-11 FMU ASSEMBLY 178
73-20 ENGINE FUEL CONTROLLING 184
73-21 FUEL CONTROLLING SYSTEM 186
73-30 ENGINE FUEL INDICATING 196
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EFFECTIVITY: ALL
ENGINE FUEL AND CONTROL 73-00
Introduction
The purpose of the engine fuel and control system is to deliver scheduled fuel
to the engine to provide the combustion required to generate propulsive
power. The system pressurizes, heats and filters the fuel and then delivers it
to the combustion chamber for burning.
General Description

The ENGINE FUEL AND CONTROL includes these subsystems:
• ENGINE FUEL DISTRIBUTION (AMM SDS 73-10-00/1)

• ENGINE FUEL CONTROLLING (AMM SDS 73-20-00/1)
• ENGINE FUEL INDICATING (AMM SDS 73-30-00/1)
The PW617F engine fuel system consists of a FMU (Fuel Metering Unit) that
contains seven major elements: the fuel pump, the PMA (Permanent Magnet
Alternator), the fuel metering system, the flow divider valve, the motive flow
system, the ecology system and shaft shear protection. The centrifugal boost
pump raises the pressure of the fuel supply to a level sufficient to charge the
inlets of the engine gear pump. The centrifugal boost pump supply is routed
through an engine oil/fuel heat exchanger before charging the inlets of the
engine gear pump. The first purpose is to cool the engine oil, which prolongs
the life of the engine bearings. The second purpose is to heat up the fuel so
that, during operation with ice in the fuel, the engine oil heat helps keeping
the fuel filter temperature above freezing. Yet, the fuel flows through a fuel
filter included in this assembly in order to protect sensitive components from
possible contaminants in the fuel. Should the fuel filter blockage become too
great, a bypass valve on the unit opens to ensure the engine is never starved
of fuel.
Afterwards, the fuel flows through the fuel metering system and then is
directed to the flow divider and to the manifolds.
The figure ENGINE FUEL AND CONTROL - SCHEMATIC DIAGRAM


EFFECTIVITY: ALL
ENGINE FUEL AND CONTROL 73-00
FUEL FILTER ASSEMBLY
FUEL

FILTER
BYPASS
INDICATOR
FUEL FILTER
INPENDING
BYPASS
SWITCH
OIL SYSTEM
LINE
FUEL
FILTER
BYPASS
LP VALVE
INLET FLOW CENTRIFUGAL
FROM FUEL HP
AIRCRAFT PUMP FUEL GEAR
FOHE
FILTER FUEL
PUMP
ENGINE FUEL AND CONTROL - SCHEMATIC DIAGRAM
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
ENGINE FUEL DISTRIBUTION 73-10
The fuel distribution system has the following functions: fuel pressurization,
filtering, injection, heat exchange, and operation of FMU (Fuel Metering Unit)
hydraulic systems.
General Description
The ENGINE FUEL DISTRIBUTION includes this subsystem:


• FMU ASSEMBLY (AMM SDS 73-11-00/1)
The components of the engine fuel distribution subsystem are mounted on

the AGB (Accessory Gearbox), except for the fuel manifold and the fuel
injectors. The propulsion system fuel distribution subsystem consists of the
following items: fuel primary ejector pump (AMM SDS 28-21-00/1), fuel
metering unit (FMU) (AMM SDS 73-11-00/1), fuel filter and fuel/oil heat
exchanger, fuel manifold, and fuel injectors.
The figure ENGINE FUEL DISTRIBUTION - BLOCK DIAGRAM provides


EFFECTIVITY: ALL
ENGINE
OIL
COOL
HEAT

HEAT FUEL FILTER
FOHE
COOL ASSEMBLY
FUEL
FUEL SUPPLY LP
MANIFOLD/
FUEL FMU
INJECTOR
PUMP
NOZZLES
BVA
ENGINE FUEL DISTRIBUTION - BLOCK DIAGRAM
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
Components All nozzles have identical pressure atomizer nozzle tips, with a single fuel
passage and nozzle jet, and two rows of holes in the air swirlier, to further
FMU ASSEMBLY (73-11) assist in atomizing the fuel spray into small fuel droplets.
The FMU (Fuel Metering Unit) assembly performs the following major During start-up, the FMU delivers fuel to the primary manifold to provide
functions: pressurization of fuel supply, regulation of fuel flow to be burnt, ignition and initial spool up of the engine. A cross-flow orifice in the FMU flow
division of primary and secondary flow, engine shutdown in normal, divider shutoff valve delivers some flow to the secondary manifold to help pre-
uncontrolled thrust and shaft shear circumstances, supply of motive fuel flow fill this manifold.
for airframe usage, and prevention of fuel discharge after engine shutdown.

The figure ENGINE FUEL DISTRIBUTION - FUEL DISTRIBUTION
FUEL MANIFOLD COMPONENT LOCATION provides further data on the preceding text.
The fuel manifold is an integral single looped ring, with dual fuel channels
machined into the ring and sealed by a cover plate. The manifold is mounted
around the gas generator case, via support pins. The fuel manifold assembly
contains connections for 7 fuel nozzles per each fuel channel, totalizing 14
fuel-nozzle connections.
FUEL-NOZZLE INJECTORS
The 14 air assisted fuel nozzle injectors are mounted equally spaced around
the gas generator case. Each fuel injector also has a check valve that closes
at engine shutdown to prevent the manifolds from draining into the
combustor. The injectors deliver atomized fuel into the combustion chamber,
where it mixes with compressor discharge air and is burned.
Operation
The FMU has two ports for the manifold, primary and secondary. The fuel is
delivered through two tubes to the bottom fairing, in the bypass duct and
interfaces with the inlet stem of the integral fuel manifold and nozzle
assembly. Independent fuel passages in the manifold inlet stem, lead to the
manifold ring. The integrated manifold and fuel nozzle assembly is protected
with last chance inlet screens at the primary and secondary inlets in the inlet
stem and in each fuel nozzle. The distribution of the nozzles is such to
maximize starting performance and to provide uniform radial temperature
distribution for the engine combustor.

EFFECTIVITY: ALL
MONOCASE

FUEL MANIFOLD AND
NOZZLE ASSEMBLY
ENGINE FUEL DISTRIBUTION - FUEL DISTRIBUTION COMPONENT LOCATION
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
FMU ASSEMBLY 73-11
Introduction operating modes except start-up. At start-up, the PRV (Pressure Regulating
Valve) automatically cuts off the BVA and motive flow to minimize the
The FMU (Fuel Metering Unit) assembly performs the following major displacement of the pump. The FMU contains the AGB (Accessory Gearbox)
functions: pressurization of fuel supply, regulation of fuel flow to be burnt, drive gear, located on the pump drive shaft. The gear is supported by two oil
division of primary and secondary flow, engine shutdown in normal, mist lubricated ball bearings.
uncontrolled thrust and shaft shear circumstances, supply of motive fuel flow
for airframe usage, and prevention of fuel discharge after engine shutdown. HIGH PRESSURE RELIEF VALVE
General Description The high pressure relief valve is incorporated to the FMU to limit the fuel

pressure in the outlet of the high pressure gear stage. The high pressure relief
The FMU assembly contains seven major elements: the LP (Low Pressure) valve limits the HP pump discharge pressure to no greater than 1.4 times the
centrifugal fuel pump and HP (High Pressure) gear fuel pump, the permanent maximum operating pressure. The design of the FMU is such that the fuel
magnet alternator (AMM SDS 73-21-00/1), the fuel metering system, the flow schedules are not significantly affected after the high pressure relief valve
divider valve, the motive flow system, the ecology system, and the shaft shear operation has occurred.
protection valve (see AMM SDS 76-20-00/1).
PROPORTIONAL MODULE AND METERING VALVE
Components
The fuel metering function is provided by a proportional module. The
LP FUEL PUMP proportional module incorporates a torque motor and the metered flow is
proportional to the current supplied to the torque motor. The torque motor is
The LP centrifugal fuel pump is mounted on the FMU. It consists of a two made up of dual coils, one per FADEC (Full Authority Digital Engine Control)
stage pump, which includes an inducer and regenerative impeller. The channel, suspended in a permanent magnet circuit. The motor is operated
inducer stage comprises multiple axial blades that provide the suction wet with electrical wires routed through the motor header.
capability. The fuel is then passed through a regenerative low-pressure boost
pump. It is designed as a second stage pump that provides a positive To minimize pump sizing, the servo flow uses a part of the fuel flow. The
pressure rise over the first stage pressure envelope (see AMM SDS metering valve is connected to the servo via a feedback spring system, which
28-21-00/1) to provide a reference pressure for the operation of the FMU results in valve position and hence burn flow, being proportional to the servo
hydraulic system. The pump pressurizes the fuel sufficiently to support the valve driver current from the FADEC.
inter-stage pressure losses and the HP pump gear fill requirements.
PRESSURE REGULATING VALVE (PRV)
HP FUEL PUMP
The PRV is located downstream of the gear pump flow. It is required to
The HP gear fuel pump is into the FMU. After passing through the LP maintain a constant differential pressure across the metering valve by
centrifugal pump, the fuel is ported to a separated filter and heat exchanger returning the pump flow excess to the inlet of the gear stage, so that the burn
assembly and then, the filtered fuel is directed to the positive gear pump to flow is proportional to the position of the metering valve ensuring its adequate
provide adequate pressurization for the fuel nozzles. The HP gear pump operation.
provides the fuel flow to be burnt in the engine (or burn flow), the flow to the
BVA (Bleed Valve Actuator) and the motive flow to the airframe ejector at all

EFFECTIVITY: ALL
FMU ASSEMBLY 73-11
NOTES: NOTES:

21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
FMU ASSEMBLY 73-11
The excess flow is determined based on the burn flow requirements. Once connects fuel drain from the manifold to the motive flow line. The residual
the burn flow requirements are satisfied, the PRV also ports high pressure combustor pressure and the gravity between the FMU and the airframe fuel
supply to the BVA and the motive flow line. tank purges the engine fuel manifold during the time the engine is spooling
down from ground idle.
WASH SCREEN
The figure FMU ASSEMBLY - COMPONENT LOCATION provides further
A wash screen filter is located upstream of the PRV and proportional module data on the preceding text.
to protect the fuel control from contamination. This screen has a wash flow
through the middle of the screen to prevent build up of contamination. The

wash flow routes through the PRV and either goes to the second stage of the
proportional module, inlet to the HP pump or to motive flow to the aircraft. The
fine filtered fuel routes to the first stage of the proportional module.
FLOW DIVIDER / SHUTOFF VALVE
The flow divider splits the flow from the metering valve between the primary
and secondary manifolds. The flow divider also provides regulation of the fuel
during the engine regimes. During starting regime the flow divider regulates
more flow to the primary nozzles and also provides the equalization of the
pressures between the primary and secondary manifolds during the light-off
regime and after this.
The flow divider is driven by the pressure differential between the metered
burn flow from the proportional module and the interstage pressure, and a
internal spring.
ECOLOGY SYSTEM
The ecology system ensures that unburned fuel is removed from the fuel
manifolds after the engine has been shutdown. This system works
automatically and drains the excess of fuel back into the motive flow line
under the influence of residual engine combustor pressure (P3) and ejector
pump suction.
MANIFOLD DRAIN VALVE
The manifold drain valve is fully closed at idle and above and routes motives
flow direct to the airframe, thereby, bypassing the manifold drain. When the
engine is commanded to shutdown, the manifold drain valve opens and

EFFECTIVITY: ALL
FMU ASSEMBLY 73-11

FMU
ACCESSORY
GEAR BOX
(REF.)
A
FMU ASSEMBLY - COMPONENT LOCATION
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
FMU ASSEMBLY 73-11
Operation The figure FMU ASSEMBLY - SCHEMATIC DIAGRAM provides further data
Fuel is supplied to the FMU from the aircraft fuel system. It is then pressurized
in three stages: a fixed ejector pump, a regenerative low pressure centrifugal
pump and a gear positive displacement pump.
The first stage, a fixed orifice ejector pump, is powered from the second stage
element. Its purpose is to keep the pump inlet filled with wing tanks fuel. The

second stage is a two-stage boost pump, which comprises an inducer and a
regenerative centrifugal pump that provides a positive pressure rise over the
full operating envelope and also a reference pressure for the operation of the
FMU hydraulic system. After passing through this two stages the fuel is ported
to a separate filter and heat exchanger assembly. Filtered fuel is then passed
to the gear positive displacement pump to provide adequate pressurization
for the fuel nozzles.
The fuel is also regulated in the metering valve and then is divided for the
primary and secondary flows by the flow divider to regulate more flow to the
primary nozzles during starting. The flow divider provides regulation of the
primary and secondary flows during the light-off regime and equalization of
the primary and secondary manifold pressures after light-off, ensuring
smooth distribution of the fuel around the combustor. This is achieved through
control lands in the flow divider valve.
Motive flow is required above idle speed to power the main airframe ejector
pump in the collector tank. Motive flow is drawn from the high-pressure supply
line. The switching of motive flow is achieved through the position of the
pressure regulating valve (PRV) that opens a second port at speed above
idle to provide fuel to the motive flow port. To minimize the pump size, the
motive flow is not supplied during engine starting.
The motive flow is also used by the ecology system ejector to provide fuel
purge from the manifold during engine shutdown. During engine spool down,
excess fuel in the manifolds is drawn back into the motive flow line under the
influences of residual engine combustor pressure (P3) and the ecology
ejector pump suction. A check valve prevents backflow from motive flow to
the flow divider and engine manifold.

EFFECTIVITY: ALL
FMU ASSEMBLY 73-11
FUEL NOZZLES
(TO COMBUSTION
CHAMBER)
FUEL FILTER ASSEMBLY FMU
FUEL FLOW DIVIDER/

FUEL FUEL SHUTOFF VALVE
FILTER

FILTER FILTER FUEL
BYPASS INPENDING BYPASS HPRV
BYPASS FILTER
INDICATOR VALVE FLOW
SWITCH
METER
WASH
HP SCREEN
GEAR
PROPORTIONAL
ESOV
MODULE
FUEL
PUMP ESOV
CABLE FROM
EMERGENCY
FUEL
FOHE SHUTDOWN SHUTOFF
PRV
SOLENOID MECHANISM
LP MANIFOLD
INLET FLOW
CENTRIFUGAL DRAIN VALVE
FROM
FUEL
AIRCRAFT
PUMP
INTEGRATED
PMA/N2
BVA MOTIVE FLOW:

TO TANK EJECTOR PUMP
FMU ASSEMBLY - SCHEMATIC DIAGRAM
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
ENGINE FUEL CONTROLLING 73-20
Introduction
The controlling system is based on a fully redundant dual channel FADEC

(Full Authority Digital Engine Control) system.
General Description
The ENGINE FUEL CONTROLLING includes this subsystem:


• FUEL CONTROLLING SYSTEM (AMM SDS 73-21-00/1)
Components
FUEL CONTROLLING SYSTEM (73-21)
The fuel controlling system provides a full range of engine control in response
to thrust command inputs from the aircraft under all conditions. The system
also provides information for aircraft indication, maintenance reporting, and
engine condition monitoring.
The figure ENGINE FUEL CONTROLLING - BLOCK DIAGRAM provides


EFFECTIVITY: ALL
ENGINE FUEL CONTROLLING 73-20
BVA

FADEC
FMU
METERED FUEL
PMA FUEL PUMP
WIRING
PIPING (FUEL)
ENGINE FUEL CONTROLLING - BLOCK DIAGRAM
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
FUEL CONTROLLING SYSTEM 73-21
Introduction The FADEC has de-

ENG SHORT DIS-
Advisory (White) tected a short time dis-
The fuel controlling system provides a full range of engine control in response PATCH
patch fault condition.
to thrust command inputs from the aircraft under all conditions. The system
also provides information for aircraft indication, maintenance reporting, and The DCU (Data Con-
engine condition monitoring. centrator Unit) has lost
the reception of
General Description ARINC (Aeronautical
E1 FADEC FAULT Advisory (White)
Radio Incorporated)

The PW617F engine control system is a computer-based electronic engine 429 data from at least
control system. It comprises: one FADEC channel
of engine 1.
• Dual channel FADEC (Full Authority Digital Engine Control).
The DCU has lost the
• EDCU (Engine Data Collector Unit). reception of ARINC
• FMU (Fuel Metering Unit) (AMM SDS 73-11-00/1). E2 FADEC FAULT Advisory (White) 429 data from at least
one FADEC channel
• PMA (Permanent Magnet Alternator). of engine 2.
• BVA (Bleed Valve Actuator) (AMM SDS 75-30-00/1). The figure FUEL CONTROLLING SYSTEM - BLOCK DIAGRAM provides
• BOV (Bleed-Off Valve) (AMM SDS 75-30-00/1).
• Fuel filtering assembly.
• Engine sensors.
• Ignition system.
The CAS (Crew Alerting System) messages related to the fuel controlling
system are shown on PFD (Primary Flight Display), in the CAS messages
field. The messages are given in the table below:
The FADEC has de-

tected a no dispatch
ENG NO DISPATCH Caution (Amber) fault condition on at
least one of the en-
gines.

EFFECTIVITY: ALL
AIRFRAME DUAL CHANNEL FADEC PW ENGINE
N1. CJ RTD. Tfuel

AIRCRAFT DISCRETE OUT Tto, T6
ENGINE SENSORS
RELAYS

METER VALVE CMD
TLA
AIRCRAFT CONTROLS FAN SPEED GOVERNING PMA (INCL. N2 SPEED) INTEGRAL FUEL PUMP/
AND SWITCHES ALTERNATOR/FMU
DISCRETE IN
RATING CALCULATION SD SOV
N2 DOT ACCEL/DECEL
CONTROL
ARINC 429 BLEED VALVE DMD
ENGINE PROTECTION LOOPS BLEED VALVE
ARINC 429
BLEED VALVE CONTROL
BUILT IN TEST
AIRCRAFT TEST UART RS422
DATA SYSTEM
DAS ETHERNET IGNITOR
IGNITION
SYSTEM
CROSS ENGINE CAN
PAMB
SENSOR RS422 UART
(EDCU)
28 VDC DCU POWER
ESSENTIAL BUS
FUEL CONTROLLING SYSTEM - BLOCK DIAGRAM
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
Components assigned to each block. When the FADEC requests a data block, this unique
identifier is referenced in the transmit block command.
FADEC (FULL AUTHORITY DIGITAL ENGINE CONTROL)
FMU (FUEL METERING UNIT)
The dual channel FADEC (also found as EEC (Electronic Engine Control) in
Pratt & Whitney publications) is located in the middle electronic compartment. The FMU is attached to the accessory gearbox at approximately the 7:00
Each FADEC is a single LRU (Line Replaceable Unit) containing two o'clock position, looking aft. The FMU performs the following main functions:
channels on two separate printed circuit boards. Each channel has two pressurization of fuel supply, regulation of fuel flow to be burnt, division of
connectors, one for the interface with the engine and the other for the primary and secondary flows, engine shutdown in normal, uncontrolled thrust,

airframe. The channels are designated channel A and channel B. An identical and shaft shear circumstances, supply of motive flow for aircraft usage, and
software is loaded into each channel and, while one channel remains in prevention of discharge of fuel after engine shutdown.
control, the other channel is health monitored. During each start, the channel The FMU contains seven major elements: the fuel pump, the permanent
in control is switched. magnet alternator, the fuel metering system, the flow divider valve, motive
The FADEC software provides thrust management and fuel supply flow system, the ecology system, and shaft shear protection valve.
commands based on inputs such as analog thrust lever angle information, FUEL FILTER ASSEMBLY
electrical power, and various hard wired discretes such as air data from the
engine and aircraft sensors. These inputs are used to control the engine low The fuel filter assembly is an integral part of the FOHE (Fuel-Oil Heat
rotor speed (N1 (Fan Rotor Speed)) and thereby the engine thrust. To achieve Exchanger). It consists of a filter head assembly, a filter element, and a filter
this, the FADEC modulates the fuel flow by means of a torque motor in the bowl. The filter element is the paper medium type supported by a spring
FMU and modulates the bleed valve position by means of a torque motor in support core made from aluminum alloy, aluminum end caps being epoxy-
the BVA. bonded to the ends of the filter medium. The fuel filter rating is 10 microns
nominal and 25 microns (0.001 in.) absolute. The assembly removes
Power for the FADEC is primarily supplied by the PMA during engine contaminants from the engine fuel in order to protect sensitive components
operation and by a 28 V DC airframe input during engine starting. This 28 V from possible contaminants in the fuel. Should the fuel filter blockage become
DC airframe input also serves as a backup source in case the PMA fails. too great, a bypass valve on the unit opens to ensure the engine is never
EDCU (ENGINE DATA COLLECTION UNIT) starved of fuel. The fuel filter assembly also includes a fuel filter impending
bypass switch that signals to the crew the filter replacement necessity to avoid
The EDCU is an electronic storage device attached to the engine. Its purpose operating in bypass mode and the fuel filter bypass indicator that
is to provide storage for engine specific information such as: power engine mechanically signals the necessity for flushing the fuel system due to opening
trims, plant identification, rating selection, cumulative engine running data, of the bypass valve.
trend monitoring data, event and fault logging, trace recording, etc. These
features may be expanded without modification to the EDCU in order to PMA (PERMANENT MAGNET ALTERNATOR)
facilitate storage of other engine specific information and to maintain a history The PMA is the primary source of power for the FADEC. The engine supplied
of the engine to which it is attached. The FADEC downloads information into PMA is a dual-wound three-phase unit, which is integrated into the FMU and
the EDCU by transmitting formatted data packets with a block number is driven by the fuel pump drive shaft. The PMA provides AC (Alternating
Current) power for both channels of the FADEC only when the engine is
EFFECTIVITY: ALL
NOTES: NOTES:

21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
running. The FADEC also derives the N2 (Core Rotor Speed) speed signal
from the frequency of the AC power provided by the PMA. The PMA is
designed as an additional source of DC (Direct Current) power in the event
of loss or interruption of the aircraft 28 V DC bus. The selection of the power
source between the aircraft 28 V DC bus and the PMA to supply the FADEC
is made considering the higher source of power available.
ENGINE SENSORS

The FADEC system includes sensors for the parameters that it needs: to
control the engine in accordance with pilot demands and airframe
requirements, to protect the engine by maintaining it within safe operating
limits, to provide the necessary data for the cockpit indicating system. Each
engine FADEC has sensors that are independent of the other, except for the
cross-communication of the engine speeds for synchronization and fault
detection purposes.
Each engine FADEC channel has a set of sensors electrically independent

of the other channel. Most sensors are duplicated, some are a part of the
engine and others are airframe items. The PW617F sensors are: N1 speed
sensor, T6 and TT0 (Inlet Total Temperature) temperature sensors, and
ambient pressure sensor, a pressure transducer which is located on the
circuit board of each channel inside the FADEC. Signals such as TLA (Thrust
Lever Angle), WOW (Weight-on-Wheels), engine anti-icing ON, wing anti-
icing ON, and others are derived from airframe sources.
BVA (BLEED VALVE ACTUATOR)
There is a BOV in the PW617F engine that requires a handling bleed for
satisfactory operation across the full operational range. The BVA is a single
stage, dual-wound electrohydraulic servomotor that modulates the bleed
valve position. The BVA is controlled by the FADEC throughout the engine
operating envelope. The BVA linear motion is converted into rotary action as
required by the bleed valve through a lever mechanism.
The figure FUEL CONTROLLING SYSTEM - COMPONENT LOCATION


EFFECTIVITY: ALL

PMA
(N2 SIGNAL)
ACCESSORY
GEAR BOX
(REF.)
FMU
(REF.)
FUEL CONTROLLING SYSTEM - COMPONENT LOCATION

Sheet 1
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
NOTES: NOTES:

EFFECTIVITY: ALL
B BVA
EDCU

B
C
FUEL FILTER
ASSEMBLY
FUEL CONTROLLING SYSTEM - COMPONENT LOCATION

Sheet 2
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
Operation
FADEC
One FADEC channel operates as the "in-control" FADEC channel providing

electronic control outputs. The other FADEC channel operates as the standby
FADEC channel, processing all inputs and software, but with electronic
control outputs disabled during normal engine operation. The standby
FADEC channel also shares selected sensor inputs, airframe commands and

FADEC status information through the Canbus cross communication links.
During operation with two capable FADEC channels, in-control software logic
causes the FADEC channels to alternate control on each successive engine
start. The FADEC power supply is primarily provided by the PMA during the
engine operation and by a 28 V DC airframe input during starting operation.
This 28 V DC airframe input also serves as a backup source in case the PMA
fails.
FUEL FILTER ASSEMBLY
Fuel from the first-stage fuel pump passes through the FOHE and then enters
the fuel filter assembly. It filtrates from the outside through the element to the
core and then exits the assembly to feed the second-stage fuel pump.
BVA
The BVA provides differential pressure across the piston as commanded by

the FADEC to position the compressor BOV to optimize overalll engine
performance and operability. The air discharged by the bleed valve is ducted
into the bypass duct and, hence, no additional nacelle ventilation is required.
At high core speed in steady condition the valve is held closed for maximum
efficiency. At lower speeds the valve is modulated open by the BVA to prevent
compressor surge. An electrical failure with the BVA results in failed
operation, in which case a careful movement of the thrust lever is required.
The figure FUEL CONTROLLING SYSTEM - COMPONENT LOCATIONS


EFFECTIVITY: ALL
ZONES CENTER
434 COMPARTMENT
(REF.)
ZONES 444
241
242 D
A

B
D
C E
PAMB
PRESSURE
SENSOR PORT
FADEC 2 ENGINE
DATA PLATE
B
ACCESSORY
PAMB GEAR BOX
FADEC 1 PRESSURE (REF.)
C
SENSOR PORT
E
FUEL CONTROLLING SYSTEM - COMPONENT LOCATIONS
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
ENGINE FUEL INDICATING 73-30
Introduction of a poppet stem, a compression spring, a valve body and a spring pin set
with cracking point of 15 psid in the fuel filter blockage condition. The bypass
The indicating system includes the components related to the indications of valve opens to ensure the engine is never starved of fuel.
the fuel system:
FUEL FILTER BYPASS INDICATOR
• Fuel flow value indication.
The fuel filter bypass indicator is installed through a threaded interface into
• Fuel filter impending bypass indication. the fuel filter housing. The indicator is a mechanical device with a colored
button popping up to visually indicate that the filter needs to be replaced. The
General Description

indicator consists of an upper cavity and a lower cavity. The upper cavity
houses a spring loaded magnetic indicator. The upstream pressure of the fuel
The engine fuel flow indication is provided by a dedicated flow meter installed filter acts on the upper side of the piston against the spring. The downstream
is each engine fuel feed line. The fuel flow display provides an indication of pressure of the fuel filter acts on the bottom side of the piston. As
the correct functioning of the fuel shutoff valve, which is a MOV (Motor- contamination increases, the pressure drops across the filter and is detected
Operated-Valve) operated by the flight crew. The cockpit engine start/stop by the bypass indicator. Once the impending delta pressure of 16 ± 2 psid is
switch signals the FADEC (Full Authority Digital Engine Control) to open or reached, this will result in a gap within the indicator, large enough to break
close the engine fuel valve. The FADEC sends a command signal to the FMU the magnetic attraction to the upper cavity indicator. The indicator spring will
(Fuel Metering Unit). The engine fuel supply is also shut-off by the shaft shear then pop up the indicator. After actuation, the indicator has to be manually
shutoff valve in the FMU. reset.
The fuel flow value is shown in green digits on the CAS (Crew Alerting FUEL FILTER IMPENDING BYPASS SWITCH
System) display, in PPH (Pounds Per Hour) or KPH (Kilograms Per Hour). In
the case of invalid data, the fuel flow display is re-configured to a red "X". The fuel filter impending bypass switch is installed in the FOHE module
through a two-bolt flange mounted in parallel with the fuel filter line and its
Components bypass valve. It senses excessive fuel filter pressure across the filter element
FUEL FLOW METER indicating the filter blockage condition. If the differential pressure across the
fuel filter exceeds 8 ± 2 psid, a mechanism actuates an electrical microswitch
The fuel flow meter is located in the HP (High Pressure) gear fuel pump line, that causes the following advisory messages on the CAS display:
downstream of the ESOV (Emergency Fuel Shutoff Valve) and upstream of
the flow divider, both internal the FMU. It consists of a electronic unit that • E 1 FUEL IMP BYP, or
measures the fuel flow and sends the fuel flow signal to the avionics to be • E 2 FUEL IMP BYP
shown in the cockpit CAS display.
When the differential pressure drops below 3 psid the mechanism reverses
FUEL FILTER BYPASS VALVE itself resulting in the microswitch changing back to its normally closed state.
The fuel filter bypass valve is located in the LP (Low Pressure) centrifugal
pump line, downstream of the FOHE (Fuel-Oil Heat Exchanger), connected
in parallel to the fuel filter line and its impending bypass indicator. It consists

EFFECTIVITY: ALL
NOTES: NOTES:

21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
The fuel filter impend-

ing bypass discrete for
E 1 FUEL IMP BYP Caution (Amber)
LH (Left-Hand) engine
is set.
The fuel filter impend-
ing bypass discrete for
E 2 FUEL IMP BYP Caution (Amber)
RH (Right-Hand) en-
gine is set.

The figure ENGINE FUEL INDICATING - COMPONENT LOCATION

EFFECTIVITY: ALL

FUEL FILTER
IMPENDING
BYPASS SWITCH
FUEL FILTER
A ASSEMBLY
(REF.)
FUEL FILTER
BYPASS
INDICATOR
FUEL FILTER
BYPASS
VALVE
A
ENGINE FUEL INDICATING - COMPONENT LOCATION

Sheet 1
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
NOTES: NOTES:

EFFECTIVITY: ALL

A
FLOW METER
ENGINE FUEL INDICATING - COMPONENT LOCATION

Sheet 2
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
Operation
A flow meter installed in each engine fuel feed line sends a signal to avionics
providing indication of fuel flow for each engine and total fuel used for both
engines with its transmitter, which is shown in green digits on the CAS display,
in PPH (Pounds Per Hour) or KPH (Kilograms Per Hour). In the case of invalid
data, the fuel flow display is re-configured to a red "X".
The figure ENGINE FUEL INDICATING - DISPLAYS provides further data



EFFECTIVITY: ALL
A
87.8 TO 87.8
ATR
2.5 N1% 2.5


ITT C
IGN ____ ____ IGN
OFF OFF
55.1 N2% 55.1

OIL PRES PSI
OIL TEMP C
LEFT FUEL RIGHT FUEL
FUEL
FLOW FLOW
FF PPH CAS
FQ LB ENG 1 FUEL IMP BYPASS
ENG 2 FUEL IMP BYPASS
TEMP XX C
ELEC CABIN
B BATT1 0V
ALT
BATT2 0V
RATE
SPDBRK DELTA-P
LFE
OXY
LG FLAPS
DN
CAS WINDOW
TAKEOFF DATA SET
OAT -237 C B
ATR ON
EICAS
ENGINE FUEL INDICATING - DISPLAYS
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80



CHAPTER 74 - IGNITION
SECTION TITLE PAGE

74-00 ENGINE IGNITION SYSTEM 206

74-10 IGNITION ELECTRICAL POWER SUPPLY 210
74-20 ENGINE IGNITION DISTRIBUTION 212
74-30 SWITCHING 214
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
ENGINE IGNITION SYSTEM 74-00
Introduction Components
The purpose of the ignition system is to provide the electrical spark to initiate IGNITION ELECTRICAL POWER SUPPLY (74-10)
the combustion of the fuel/air mixture in the engine during start, auto-relight
and when continuous ignition is required. The aircraft provides power to the engine mounted ignition exciter that
supplies the igniters.
General Description
ENGINE IGNITION DISTRIBUTION (74-20)
The ENGINE IGNITION SYSTEM includes these subsystems:
The high voltage necessary for the sparking on the spark igniters during the

• IGNITION ELECTRICAL POWER (AMM SDS 74-10-00/1) engine starting is supplied by the ignition exciter box and fed to the spark
SUPPLY igniters through ignition cables.
• ENGINE IGNITION DISTRIBU- (AMM SDS 74-20-00/1)
TION SWITCHING (74-30)
• SWITCHING (AMM SDS 74-30-00/1) The ignition switching system controls the operation of the ignition circuit.
For additional information about the ignition system components, refer to the
The ignition system is controlled by the FADEC (Full Authority Digital Engine
last revision of PW617F Engine Maintenance Manual PN 3072162 (EMM
Control) for automatic engine starting and auto-relight. Continuous ignition
TASK 72-00-00-800-801/91).
can be manually set through a cockpit switch (AMM SDS 74-30-00/1).
Operation
The engine is equipped with a dual ignition system that is under the control
of both channels of the FADEC. The system comprises two independent The FADEC controls the ignition system by discrete outputs from each
ignition exciters in a single housing together with leads and igniters. channel of the FADEC. The ignition exciter uses 28 V power supply from the
airframe.
An IGN A and B icon is displayed for each engine showing which of the ignition
systems are being commanded by the FADEC. Normally during ground starts The cockpit ignition interface is composed of two three-position switches: ON
only one ignition channel is used and the channel selected alternates on each – AUTO – OFF. The “OFF” position is used for a dry motoring run. When the
start. In flight starts use both ignition systems. Similarly, the auto-relight switch is in the position “ON”, the FADEC provides continuous power to the
function will command both ignition systems on if the engine is detected to ignition. The “AUTO” position is for normal operation. The “Auto” position puts
have flamed out. If the pilot moves the Ignition selector switch to override the control of the igniters under the direction of the FADEC such that ignition
position, both ignition channels will be commanded to operate. The "A" and/ can be synchronized with the start sequence and for automatic relight in flight.
or "B" indication will only illuminate if the FADEC has commanded an ignition For ground starts and temperatures below 0 °C, the FADEC automatically
channel to operate. The ignition indication presents the following: "A", "B", "A commands both exciters via auxiliary ignition command relay.
B”, “OFF” or blank. The "OFF" indication provides confirmation to the crew
that the controls are correctly set for the dry motoring procedure. Blank An IGN A and B icon is displayed on the EICAS (Engine Indication Crew Alert
indication will be provided when the FADEC is in the automatic mode to System) for each engine showing which of the ignition systems are being
command the ignition, but neither ignition is active. commanded by the FADEC.

EFFECTIVITY: ALL
NOTES: NOTES:

21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
The "A" and/or "B" indication will only illuminate if the FADEC has
commanded an ignition channel to operate. The ignition indication presents
the following: "A", "B", "A B”, “OFF” or blank. The OFF indication provides
confirmation to the crew that the controls are correctly set for the dry motoring
procedure. Blank indication will be provided when the FADEC is in the
automatic mode to command the ignition, but neither ignition is active.


WARNING: MAKE SURE THAT THE IGNITION SYSTEM WAS OFF FOR
A MINIMUM OF 6 MINUTES BEFORE YOU DISCONNECT IT.
AN ELECTRIC SHOCK FROM THE HIGH VOLTAGE OF THE
IGNITION EXCITER CAN KILL YOU.
Residual voltage in the ignition exciter box may be dangerously high. It is

important that the ignition be turned off and the system be inoperative for at
least 6 minutes before the removal of any component of the ignition system.
Obey the WARNING placards on the ignition exciter box.
The figure ENGINE IGNITION SYSTEM - CONTROLS AND INDICATIONS


EFFECTIVITY: ALL
B CAS
87.8 TO 87.8
A ENG NO ATR
DESPATCH
E1 / 2 SHORT
DSPTCH
A 2.5 N1% 2.5
IGNITION IGNITION
INDICATION INDICATION

ITT C
IGN ____ ____ IGN
AB AB
55.1 N2% 55.1
OIL PRES PSI
C CAS WINDOW OIL TEMP C
FUEL
A FF KGH
FQ KG
TEMP XX C
ELEC CABIN
BATT1 0V
ALT
ENG FIRE EXTINGUISHER TRIM BATT2 0V
YAW RATE
SHUTOFF 1 BOTTLE SHUTOFF 2
SPDBRK DELTA-P
LEFT RIGHT
DISCH
LFE
ROLL
OXY
OFF LWD RWD
LG FLAPS
ENG START/STOP
RUN RUN
PITCH BKP DN
DN
TAKEOFF DATA SET
OAT -237 C
UP
1
ENG IGNITION
2
MODE
ATR ON
+
ON BKP
AUTO
OFF OFF
EICAS DISPLAY
1 2
FIRE/ENG/TRIM
B
CONTROL PANEL
C
ENGINE IGNITION SYSTEM - CONTROLS AND INDICATIONS
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
IGNITION ELECTRICAL POWER SUPPLY 74-10
Introduction connected to the DC bus 2. Note that in electrical emergency conditions,

power supply to the channel B ignition exciters are automatically switched
The aircraft provides power to the engine mounted ignition exciter that from the DC busses to the emergency bus. The FADEC channel A has the
supplies the igniters. capability of activating channel B ignition exciter in an electrical emergency
condition by commanding a single aircraft relay.
This power is fed to the ignition exciter box. The ignition exciter box is the unit
responsible for supplying high tension pulses to the spark igniters. The position of the cockpit ignition switches selects the mode of operation
(AMM SDS 74-30-00/1).
General Description

The ignition exciter is located on the compressor case. The exciter is a sealed
unit containing electronic components encased in epoxy resin. The unit is The following recommendation regarding the ignition components is
energized during the engine starting sequence and when selected by the important:
pilot. Refer to (AMM SDS 80-00-00/1).
WARNING: MAKE SURE THAT THE IGNITION SYSTEM WAS OFF FOR
The two individual ignition cable assemblies carry the electrical energy output A MINIMUM OF 6 MINUTES BEFORE YOU DISCONNECT IT.
from the ignition exciters to the spark igniters. Each lead assembly consists AN ELECTRIC SHOCK FROM THE HIGH VOLTAGE OF THE
of an electrical lead contained in a flexible metal braiding. One coupling nut IGNITION EXCITER CAN KILL YOU.
at each end of the assembly connects to the ignition exciter and the spark
igniters. The figure IGNITION ELECTRICAL POWER SUPPLY - COMPONENTS
Components.
IGNITION EXCITER
The exciter is a sealed unit containing electronic components encased in

epoxy resin.
Operation
The system is energized from the aircraft nominal 28 V DC power supply and
operates in the 9 to 30 V range. The operating output voltage range is 14 to
17 kV.
The FADEC (Full Authority Digital Engine Control) controls the ignition by
discrete outputs from each channel of the FADEC. The ignition exciter uses
28 V power supply from the airframe and fires the sparks at a fixed interval
when commanded ON. The left channel B ignition exciter is normally
connected to DC Bus 1, whereas the right channel B ignition exciter is

EFFECTIVITY: ALL
IGNITION ELECTRICAL POWER SUPPLY 74-10
IGNITION
EXCITER
A

IGNITER
IGNITER
A−A
A B
IGNITER
B
IGNITION ELECTRICAL POWER SUPPLY - COMPONENTS LOCATION
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
ENGINE IGNITION DISTRIBUTION 74-20
Introduction
The high voltage necessary for the sparking on the spark igniters during the
engine starting is supplied by the ignition exciter box and fed to the spark
igniters through ignition cables.
General Description
The distribution system comprises two ignition cables and two igniters. The

system receives pulsed high voltage from the ignition exciter. The high
voltage is carried from the ignition exciter to the igniters through the ignition
cables.
Components
IGNITION CABLES
The two individual ignition cable assemblies carry the electrical energy output
from the ignition exciters to the spark igniters. Each lead assembly consists
of an electrical lead contained in a flexible metal braiding. One coupling nut
at each end of the assembly connects to the ignition exciter and the spark
igniters.
SPARK IGNITERS
Two spark igniters, positioned at 4 and 8 o’clock respectively, are installed

through igniter support tubes on the gas generator case on the combustion
chamber.
Spark igniters are fragile components that must be handled with care. Should
a spark igniter be dropped, internal damage not detectable by a visual
inspection can occur. In this case, replace the spark igniter.
The figure ENGINE IGNITION DISTRIBUTION - COMPONENT LOCATION


EFFECTIVITY: ALL
ENGINE IGNITION DISTRIBUTION 74-20
ENG IGNITION
ON
AUTO
B 1
OFF
2
FIRE/ENG/TRIM

CONTROL PANEL
FADEC
A B
CONTROL PEDESTAL
C
ZONES
223 TYPICAL
224 A
IGNITION
EXCITER
CENTER
COMPARTMENT
(REF.)
FADEC 2
C IGNITER (2x)
FADEC 1
ENGINE IGNITION DISTRIBUTION - COMPONENT LOCATION
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
SWITCHING 74-30
Introduction The figure SWITCHING - RH ENGINE SCHEMATIC DIAGRAM provides

The ignition switching system controls the operation of the ignition circuit.
General Discription
The system has one ignition switch for each engine, installed on the Fire/Eng/
Trim control panel, in the cockpit.
The switch sends discrete outputs to the related FADEC (Full Authority Digital

Engine Control), so the FADEC can control the operation of the Ignition
exciter.
Components
IGNITION SWITCH
The ignition switch is a rotary switch with three positions. It operates as

follows:
• OFF: Inhibits the ignition.
• AUTO: Gives the FADEC full authority on ignition control.
• ON: Commands the FADEC to continuously activate the two ignition

channels.
Operation
With the ignition switch in the AUTO position during a ground start, only the
FADEC in control commands ignition. At the end of the starting cycle, the
FADEC deactivates the ignition exciter.
With the ignition switch in the ON position, the two FADECs command ignition
during start. The ignition is not deactivated at the end of the starting cycle.
With the ignition switch in the OFF position, the ignition exciter is always
inhibited.

EFFECTIVITY: ALL
SWITCHING 74-30
B
A 87.8 TO
ATR
87.8
2.5 N1% 2.5


IGNITION IGNITION
INDICATION INDICATION
ITT C
IGN ____ ____ IGN
AB AB
55.1 N2% 55.1
OIL PRES PSI
OIL TEMP C
FUEL
FF KGH
FQ KG

YAW
TEMP XX C
SHUTOFF 1 BOTTLE SHUTOFF 2 ELEC CABIN
DISCH
LEFT RIGHT BATT1 0V
ALT
BATT2 0V
ROLL RATE
OFF LWD RWD SPDBRK DELTA-P
LFE
ENG START/STOP
RUN RUN OXY
PITCH BKP LG FLAPS
DN
UP DN
1 2
ENG IGNITION MODE
+ TAKEOFF DATA SET
ON BKP
OAT -237 C
AUTO
OFF OFF
ATR ON
1 2
FIRE/ENG/TRIM EICAS DISPLAY

CONTROL PANEL
B
A
SWITCHING - CONTROLS/INDICATIONS
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
SWITCHING 74-30
NOTES: NOTES:

EFFECTIVITY: ALL
SWITCHING 74-30
EPDU FWD COMPARTMENT
A3 EXCITER 1A NACELLE
A2
EMERG BUS
A1
+28VDC
CB0225
RETURN
1 2 SP0046
B3 A1
B2 A2
B1 A3
B1
B2
B3 EXCITER 1B NACELLE

LPDU LH CENTER COMPARTMENT +28VDC
DC BUS 1
RETURN
A3
CB0107 A2
A1
B3
B2
B1 ENG IGNITION
ON
AUTO
OFF
1 2
LH CENTER COMPARTMENT
CONTROL PEDESTAL
IGNITION COMMAND
IGNITION 1
OFF 4
IGNITION OFF INPUT 5
AUTO
IGNITION ON INPUT
6
ON
RETURN 2
PIN A−C
CHANNEL B
CHANNEL A
IGNITION COMMAND
IGNITION AUXILIAR
FADEC 1
PIN B−C
RETURN 2
1
IGNITION ON INPUT
3
SWITCHING - LH ENGINE SCHEMATIC DIAGRAM
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
SWITCHING 74-30
NOTES: NOTES:

EFFECTIVITY: ALL
SWITCHING 74-30
EPDU FWD COMPARTMENT
A3 EXCITER 2A NACELLE
A2
EMERG BUS
A1
+28VDC
CB0226
RETURN
1 2 SP0046
B3 A1
B2 A2
B1 A3
B1
B2
B3 EXCITER 2B NACELLE

RPDU RH CENTER COMPARTMENT +28VDC
DC BUS 2
RETURN
A3
CB0198 A2
A1
B3
B2
B1 ENG IGNITION
ON
AUTO
OFF
1 2
RH CENTER COMPARTMENT
CONTROL PEDESTAL
IGNITION COMMAND
IGNITION 2
OFF 4
AUTO
IGNITION ON INPUT
6
ON
RETURN 2
PIN A−C
CHANNEL B
CHANNEL A
IGNITION COMMAND
IGNITION AUXILIAR
FADEC 2
PIN B−C
RETURN 2
1
IGNITION ON INPUT
3
SWITCHING - RH ENGINE SCHEMATIC DIAGRAM
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80



CHAPTER 75 - AIR
SECTION TITLE PAGE

75-00 AIR 222

75-30 COMPRESSOR CONTROL 226
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
AIR 75-00
Introduction The compressor control is achieved through the BOV (Bleed-Off Valve),
controlled by the BVA (Bleed Valve Actuator). This system is necessary to
The engine air system provides air from and through the engine for airframe control the air flow through the engine and to maintain the compressor
services, wing deicing, engine compressor control and engine sealing and operability margins across the full range of operation of the engine. Refer to
cooling. AMM SDS 75-30-00/1 for details.
General Description Components
The AIR includes this subsystem: COMPRESSOR CONTROL (75-30)

• COMPRESSOR CONTROL (AMM SDS 75-30-00/1) The compressor control system is used to control the air flow through the
engine and to maintain the compressor operability margins across the full
The PW617F is equipped with an inboard and an outboard HPC (High range of operation of the engine.
Pressure Compressor) delivery air bleed ports.
For additional information about the engine air system components, refer to
The engine air flow is used outside the engine and within the engine for the the last revision of PW617F Engine Maintenance Manual PN 3072162 (EMM
following purposes: TASK 72-00-00-800-801/91).
• Supply external bleed air to the ECS (Environmental Control System) and The figure AIR - AIR SYSTEM DUCTS - COMPONENT LOCATION provides
to the EAI (Engine Anti-Icing) System. further data on the preceding text.
• Supply internal bleed air for compressor control.
• Supply internal bleed air for sealing the bearing compartments and cooling
engine.
The ECS provides controlled air for cockpit and cabin pressurization and
heating as well as pneumatic supply for wing and horizontal stabilizer deicing
boots, using discharge air from the inboard engine bleed port. This air bled
from the engine passes through the ECS / Deicing bleed duct before going
to the PRSOV (Pressure Regulating and Shutoff Valve) to be distributed.
Refer to AMM SDS 36-11-00/1 for details.
The EAI system provides thermal energy to prevent ice accretion on air inlet,
using discharge air from the outboard engine bleed port. This air bled from
the engine passes through the EAI tubing with a shutoff valve and a pressure
transducer mounted on it, before going to the air inlet. Refer to AMM SDS
30-21-00/1 for details.

EFFECTIVITY: ALL
AIR 75-00
A

PRSOV
(REF.)
ECS/DEICING
BLEED DUCT
A
AIR - AIR SYSTEM DUCTS - COMPONENT LOCATION

Sheet 1
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
AIR 75-00
NOTES: NOTES:

EFFECTIVITY: ALL
AIR 75-00

EAI PRESSURE
TRANSDUCER
EAI TUBING
EAI SHUTOFF VALVE
AIR - AIR SYSTEM DUCTS - COMPONENT LOCATION

Sheet 2
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
COMPRESSOR CONTROL 75-30
Introduction The BVA operates the BOV under FADEC (Full Authority Digital Engine
Control) direction and is hydraulically operated using the FMU (Fuel Metering
The compressor control system is used to control the air flow through the Unit) fuel. Refer to AMM SDS 73-11-00/1 for details about the FMU.
engine and to maintain the compressor operability margins across the full
range of operation of the engine. The position of the BVA is computed in the FADEC software to optimize the
compressor discharge bleed off-take from the engine as a function of the
General Description current steady-state and transient engine operating condition. During steady-
state at high core speed the valve is held closed for maximum efficiency.
The compressor control is achieved by the operation of a single BOV (Bleed-
During transient operation, combustor relights and engine starting at lower

Off Valve), optimizing the compressor bleed discharge off-take. The BOV is
speeds the valve is modulated open for compressor surge avoidance.
controlled by the BVA (Bleed Valve Actuator).
The air discharged by the BOV is directed into the bypass duct, and hence,
A single-stage dual wound electro-hydraulic servomotor is linked to a lever
no additional nacelle ventilation is required to accommodate this flow.
mechanism that converts the linear motion to rotation, as required by the
BOV. The rotary motion modulates the BOV position to control the engine The figure COMPRESSOR CONTROL - SCHEMATIC DIAGRAM provides
operating limits at high core speed in steady state and during transient further data on the preceding text.
conditions.
Components
The PW617F engine compressor bleed-control-system contains the following

components:
BOV
The compressor BOV is a piston valve that has the function of regulating the
air flow between the mixed flow rotor and centrifugal impeller of HPC (High
Pressure Compressor), directing exceeding air from the mixed stage to
optimize overall engine performance and to prevent engine compressor
surges. This valve stays completely open during idle engine operation and
starts to close as thrust lever angle increases.
BVA
The BVA is a bi-directional fully modulated linear actuator used to position

the engine’s compressor BOV via a lever mechanism.
Operation

EFFECTIVITY: ALL
COMPRESSOR CONTROL 75-30
FADEC 1 FADEC 2
CHANNEL A
CHANNEL B
CHANNEL A
CHANNEL B
BVA T/M
BVA T/M
BVA T/M
BVA T/M

LO
LO
LO
LO
HI
HI
HI
HI
BVA T/M A BVA T/M B BVA T/M A BVA T/M B
BLEED VALVE ACTUATOR LH ENGINE BLEED VALVE ACTUATOR RH ENGINE
COMPRESSOR CONTROL - SCHEMATIC DIAGRAM
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80



CHAPTER 76 - ENGINE CONTROLS
SECTION TITLE PAGE

76-00 ENGINE CONTROLS 230

76-10 POWER CONTROL 232
76-11 MECHANICAL CONTROL SYSTEM 234
76-12 ELECTRONIC CONTROL SYSTEM 242
76-20 EMERGENCY SHUTDOWN 250
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
ENGINE CONTROLS 76-00
Introduction
The engine control system provides means of controlling the PW617F engine
operation under all thrust requirements allowed, as well as during emergency
shutdown.
General Description
The ENGINE CONTROLS includes these subsystems:


• POWER CONTROL (AMM SDS 76-10-00/1)
• EMERGENCY SHUTDOWN (AMM SDS 76-20-00/1)
Components
POWER CONTROL (76-10)
The thrust control subsystem furnishes means of controlling the fuel control
system of the PW617F engine.
EMERGENCY SHUTDOWN (76-20)
The emergency shutdown subsystem furnishes means of controlling the flow

of fluids to and from the engine during emergency procedures.
The figure ENGINE CONTROLS - ENGINE CONTROL SCHEMATIC


EFFECTIVITY: ALL
ENGINE CONTROLS 76-00
ENGINE CONTROL

THRUST EMERGENCY
CONTROL SHUTDOWN
ENGINE
− THRUST − EMERGENCY
MANAGEMENT STOP
ENGINE CONTROLS - ENGINE CONTROL SCHEMATIC
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
POWER CONTROL 76-10
Introduction and the air data from engine and airframe sensors (transmitted through the
ARINC (Aeronautical Radio Incorporated)). The FADEC uses these inputs
The thrust control subsystem furnishes means of controlling the fuel control that represent pilot demands, through the TLA, and ambient conditions,
system of the PW617F engine. through the sensors, also taking into account the engine operating limits to
calculate appropriate reference corrected and physical N1 (Fan Rotor Speed)
General Description and N2 (Core Rotor Speed) speeds for any given throttle position.
The POWER CONTROL includes these subsystems: The N1 and N2 control is achieved through the actuation of the FADEC,
configuring the engine to a required fuel flow demand as a function of thrust

• MECHANICAL CONTROL SYS- (AMM SDS 76-11-00/1) requirement.
TEM
• ELECTRONIC CONTROL SYS- (AMM SDS 76-12-00/1) Some of the thrust setting references are also modified by various hardwired
TEM discrete inputs.
The aircraft provides thrust requirements and ADC (Air Data Computer) The figure POWER CONTROL - THRUST CONTROL SCHEMATIC provides
inputs to the related two-channel FADEC (Full Authority Digital Engine further data on the preceding text.
Control), which manages thrust through adjusting fuel flow.
Components
MECHANICAL CONTROL SYSTEM (76-11)
The thrust control quadrant sends TLA (Thrust Lever Angle) signal to the
FADEC (Full Authority Digital Engine Control) for thrust management
purposes, and for other aircraft systems, for control purposes.
ELECTRONIC CONTROL SYSTEM (76-12)
The electronic control system of the PW617F engine is a computer-based

electronic control system (FADEC (Full Authority Digital Engine Control)
based), which provides full range of engine control under all conditions and
provides information for cockpit indication, maintenance reporting and engine
condition monitoring.
Operation
The main thrust setting inputs for thrust management are the TLA (Thrust
Lever Angle), provided by the TCQ (Thrust Control Quadrant) (hardwired),

EFFECTIVITY: ALL
POWER CONTROL 76-10
− ELECTRICAL POWER SUPPLY − AMB. DATA

− WOW AIRFRAME AIR DATA (MACH, PO)
− TTO HEATER
− ENGINE POS
− AIRCRAFT ID
ARINC 429
ARINC 429

FADEC CHANNEL A
CAN BUS
(X TALK)
ARINC 429 TLA (RVDTs)
− PMA ELECTRICAL POWER SUPPLY

− FUEL FLOW COMMAND
− BLEED VALVE COMMAND
− ENG. SENSORS (N1, N2, T6) FADEC CHANNEL B
− TTO
TCQ
POWER CONTROL - THRUST CONTROL SCHEMATIC
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
MECHANICAL CONTROL SYSTEM 76-11
Introduction
The thrust control quadrant sends TLA (Thrust Lever Angle) signal to the
FADEC (Full Authority Digital Engine Control) for thrust management
purposes, and for other aircraft systems, for control purposes.
The figure MECHANICAL CONTROL SYSTEM - TCQ LOCATION provides


EFFECTIVITY: ALL

A
THRUST CONTROL
QUADRANT
MECHANICAL CONTROL SYSTEM - TCQ LOCATION
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
General Description
The TCQ (Thrust Control Quadrant) provides lever position (TLA) to the
FADEC via a RVDT (Rotary Variable Differential Transducer). Each RVDT
has two electrically independent channels, one for each of the two FADEC
channels of a given engine. The FADEC provides excitation and
demodulation of the RVDT, providing the pilot with full and progressive
modulation of thrust in response to movements of the TLA, together with
accurate thrust setting to meet engine thrust ratings.

Engine thrust is directly related to N1 (Fan Rotor Speed) speed. The FADEC
calculates a N1 speed setting corresponding to the TLA position selected,
and compensates this setting for ambient temperatures and pressures,
aircraft bleed off-takes and operating modes. The FADEC then governs the
engine to this N1 value.
The TLA is sent to the necessary systems by discrete switches.
The status message related to the refueling sub-subsystem The CAS (Crew
Alerting System) messages related to the mechanical control system are
shown on PFD (Primary Flight Display), in the CAS display. The messages
are given in the table below:
The dual TLARVDT

corresponding to LH
E1 TLA FAIL Caution (Amber)
(Left-Hand) engine is
failed.
The dual TLARVDT
corresponding to RH
E2 TLA FAIL Caution (Amber)
(Right-Hand) engine
is failed.
The figure MECHANICAL CONTROL SYSTEM - TCQ SCHEMATIC provides


EFFECTIVITY: ALL
TCQ − THRUST CONTROL QUADRANT
THRUST LEVERS
MICROSWITCH S1 MICROSWITCH S5

LH TO/GA SWITCH RH TO/GA SWITCH
FADEC 1A RVDT FADEC 1A RVDT FADEC 2A FADEC 2A
FADEC 1B RVDT FADEC 1B RVDT FADEC 2B FADEC 2B
ENGINE
AUTOMATIC SPEED ENGINE
PNEUMATIC
FLIGHT CONTROL BRAKE PNEUMATIC
BLEED
SYSTEM − AFCS (OPTIONAL) BLEED (CABIN)
(COCKPIT)
MECHANICAL CONTROL SYSTEM - TCQ SCHEMATIC
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
Components
The TCQ comprises the external components that follow:
• Throttle levers (one per engine).
• TOGA (Take off / Go Around) switches (one per engine).
• Edge light panel.


Furthermore, the thrust control quadrant comprises the internal components
that follow:
• Thrust lever angle RVDTs (one per engine/dual channel).
• Aircraft systems microswitches set (4 per lever).
The figure MECHANICAL CONTROL SYSTEM - TCQ EXTERNAL

COMPONENTS provides further data on the preceding text.

EFFECTIVITY: ALL
EDGE LIGHT
PANEL

TO/GA
SWITCHES (2)
A THROTTLE
LEVERS (2)
THRUST CONTROL
QUADRANT
MECHANICAL CONTROL SYSTEM - TCQ EXTERNAL COMPONENTS
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
Operation
The lever TCQ combines two lever assemblies, one for each engine, which
may be operated independently or in unison. Each lever assembly comprises
a thrust lever for forward thrust control. The levers travel 73 degrees from Idle
to Max, sensed by the RVDT, which provides the position to the FADEC. The
full travel of the thrust levers provides RVDT output distributed as:
Idle 0.0° to 4.0°


Max Cruise 40.0° ± 2.0°
Max Continuous/Climb 51.0° ± 2.0°
Max Takeoff 61.0° ± 2.0°
Max 71.0° ± 2.0°
For redundancy, TLA is input to both channels A and B, and also cross
communicated between channels. In the very unlikely event that the TLA
signal becomes unavailable on either channels, the engine thrust is switched
to idle and the caution message ENG 1 (2) TLA FAIL is displayed on the CAS
display.
Each thrust lever actuates 4 discrete switches designed to automatically

provide electrical signals to the system that follows:
• AFCS (Automatic Flight Control System) (AMM SDS 22-00-00/1)
• Engine pneumatic bleed (AMM SDS 36-11-00/1)
• Speed brake (AMM SDS 27-60-00/1) - Optional
The TCQ also features TOGA disconnect lever switches to enable the pilot
to manually generate TOGA signal.
The figure MECHANICAL CONTROL SYSTEM - TCQ LEVER TRAVEL


EFFECTIVITY: ALL
69° − 73°
59° − 63°
49° − 5
3°
38° −
42°
MAX
MAX TO/GA MAX
CON/CLB 0°
MAX −4
°
CRZ

IDLE
B
THRUST CONTROL
QUADRANT
A B
MECHANICAL CONTROL SYSTEM - TCQ LEVER TRAVEL
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
ELECTRONIC CONTROL SYSTEM 76-12
Introduction Beyond thrust management, the FADEC provides engine limits protection,
controlled transient engine operation, fault detection, and messages to the
The electronic control system of the PW617F engine is a computer-based aircraft.
electronic control system (FADEC (Full Authority Digital Engine Control)
based), which provides full range of engine control under all conditions and The CAS (Crew Alerting System) messages related to the electronic control
provides information for cockpit indication, maintenance reporting and engine system are shown on PFD (Primary Flight Display), in the CAS display. The
condition monitoring. messages are given in the table below:
General Description
The take off data has

The FADEC has two identical, isolated channels due to the criticality of proper not been successfully
control system operation. During engine operation, one channel is in active ENG NO TO DATA Caution (Amber) entered. This mes-
mode and the other channel is in standby mode. Each channel receives sage is set on ground
identical but separate inputs from the engine sensors which are also only.
electrically dual redundant. After signal conditioning, the two channels share An uncommanded en-
data via a cross channel data link. gine shutdown was
E1 FAIL Caution (Amber)
detected on LH (Left-
The FADEC is powered by the PMA (Permanent Magnet Alternator), which
Hand) engine.
also provides N2 (Core Rotor Speed) signal.
An uncommanded en-
In order to ensure that all engines have the same thrust at a fan speed rating gine shutdown was
and that there is a consistent temperature uptrim margin for each engine, the E2 FAIL Caution (Amber)
detected on RH
FADEC uses trimmed values of N1 (Fan Rotor Speed) and ITT (Interstage (Right-Hand) engine.
Turbine Temperature) for control and indication purposes. The trim data is
located on the engine data plate and is loaded into the EDCU (Engine Data The pilot may is un-
Collector Unit). able to modulate
E1 CTRL FAULT Caution (Amber) thrust on LH engine or
The FADEC controls the operation, performance and efficiency the LH engine re-
characteristics of the engine as follows: The FADEC monitors inputs from the spond it slowly.
aircraft (TLA (Thrust Lever Angle), discrete signals and ARINC (Aeronautical
Radio Incorporated) data) from the engine, and modulates the fuel flow by The pilot may is un-
means of a torque motor in the FMU (Fuel Metering Unit) to vary engine speed able to modulate
(N1 or N2) to achieve the required thrust. The FADEC also modulates by E2 CTRL FAULT Caution (Amber) thrust on RH engine or
means of a torque motor in the bleed valve (compressor pressure control) the the RH engine re-
engine operating condition. spond it slowly.

EFFECTIVITY: ALL
NOTES: NOTES:

21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
An operational limit
was exceeded in at
ENG EXCEEDANCE Caution (Amber)
least one of the two
engines during flight.
Components
The electronic control system comprises the following components of the

engine:

• A two channel FADEC;
• Engine sensors (AMM SDS 77-00-00/1);
• PMA (AMM SDS 73-20-00/1);
• Bleed Valve and Bleed Valve Actuator (AMM SDS 75-30-00/1).
The electronic control system utilizes the following components for control
purposes from the aircraft:
• Display Units (AMM SDS 31-61-00/1);
• EDCU (AMM SDS 73-21-00/1);
• TLA (see AMM SDS 76-11-00/1).
The figure ELECTRONIC CONTROL SYSTEM - LOCATION provides further


EFFECTIVITY: ALL
ZONES CENTER
434 COMPARTMENT
(REF.)
ZONES 444
241
242 D
A

B
D
C E
PAMB
PRESSURE
SENSOR PORT
FADEC 2 ENGINE
DATA PLATE
B
ACCESSORY
PAMB GEAR BOX
FADEC 1 PRESSURE (REF.)
C
SENSOR PORT
E
ELECTRONIC CONTROL SYSTEM - LOCATION
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
Operation • No engine limits shall be exceeded due to the application of power

START reserve.
The FADEC schedules fuel flow during starting based on N2. As the engine The display indicates an ATR icon when it is enabled or armed. This indication
accelerates, the FADEC monitors ITT to ensure that the engine accelerates is active in takeoff mode only. The icon is positioned below the thrust mode
to idle without exceeding defined limits. FADEC incorporates automatic icon. In case the ATR becomes enable, a white indication of ATR appears
engine cool down motoring prior to auto start. The pilot can also abort any just below the thrust mode. If the ATR is armed (TLA at TOGA (Take off / Go
start attempt at any time by moving the starter switch to STOP. The FADEC Around) position for takeoff) then the ATR indication is green. In case of an

only aborts the start in the event of detecting an unsatisfactory operating engine failure and ATR being triggered, the ATR indication disappears and
condition during a ground start. the thrust mode changes to TO-RSV.
TAKEOFF DATA SET ENGINE SPEED GOVERNING
For takeoff procedures, the pilot is requested to enter the OAT (Outside Air During operation above idle, the N1 loop is usually in control, providing
Temperature). This method provides reliable temperature information to the governed N1 speed at the reference value. At idle, the N2 loop is in control
FADEC for thrust computation during takeoff phase. The OAT value has to providing N2 governed idle speed. During the start process, fuel flow is
be selected before takeoff. To change the OAT value, the pilot shall use the scheduled to a MAX value for a smooth engine start based on corrected N2.
soft keys OAT! and OAT", and accept it through ACCEPT button on the MFD At idle, the engine is governed to the N2 idle speed reference, as N1 is quite
(Multi-Function Display). low and slow to respond to fuel flow changes. At any condition above idle,
the engine is governed by a proportional plus loop integral to the N1 reference
ATR (AUTOMATIC THRUST RESERVE) speed as selected by the TLA position. The FADEC software provides many
separate control loops. The loop in control at any time depends on the pilot's
In addition to the OAT selection, the pilot is requested to select the ATR demand and on the engine conditions. Most of the loops provide transient or
(Automatic Thrust Reserve) for takeoff. Before takeoff, ATR is selected ON protective functions.
as default, but can be disabled via the soft keys ON and OFF on MFD takeoff
data set page. The power reserve improves aircraft performance. The ATR CRUISE SPEED CONTROL
increases thrust in case of OEI (One Engine Inoperative), only during takeoff
phase: During operation between flight idle and cruise, under certain conditions, it is
possible to set an aircraft constant speed controlled by the FADEC. The pilot
• The FADEC detects OEI based on N1 mismatch between both engines is able to set the cruise speed control to ON through the CSC switch on the
or loss of engine-to-engine communication, or; main instrument panel, when the following conditions are true:
• The FADEC detects the TLA to MAX position during TO (Takeoff) phase • CSC (Cruise Speed Control) request is received.
and ATR OFF false signal (ATR is selected ON);
• More than 5 s have elapsed since the last disengagement of the CSC
• The bleed valve for pressurization is commanded to close through the function.
FADEC in case OEI condition is detected and the aircraft is at takeoff
mode;

EFFECTIVITY: ALL
NOTES: NOTES:

21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
• Calibrated air speed is above 105 kts when initial CSC activation request ENGINE TRANSIENT CONTROL
is accepted. Subsequently, calibrated air speed is above 100 kts.
The FADEC software contains several features to provide satisfactory
• Mach number is below or equal to 0.78 when initial CSC activation request operation of the engine across its thrust and operating envelope. Acceleration
is accepted. Subsequently, Mach number is below 0.72. and deceleration maneuvers, in response to rapid TLA movements, are
controlled based on the rate of change of N2 and fuel flow. N2 schedules are
• WOW (Weight-on-Wheels) = false. set to ensure the avoidance of surge during normal operation. Fuel flow limits
are set to prevent surge and flameout during the initial portion of the
• Engine TLA position is above idle and below max. cruise.
acceleration. Transitions between the various controlling loops during

• TLA is not moved more than 2 degrees since the CSC was engaged. acceleration and deceleration are not perceptible.
• Engine data is valid. The figure ELECTRONIC CONTROL SYSTEM - SCHEMATIC DIAGRAM
• Smart probe data is valid.
• AHRS (Attitude and Heading Reference System) data is valid.
• Flap data is valid.
• The FD (Flight Director)/autopilot vertical mode is either altitude hold or

glideslope capture.
• The control wheel steering FD mode is not active.
• Both engines in operation and above idle.
• The absolute difference of the Cruise Speed Control N1 command

between the two engines is less than 1%.
• No E1 (2) CONTROL FAULT CAS message is active.
• Flap angle is below 5 degrees or Flap is above 34.8 degrees.
When the cruise speed control is engaged, the FADEC controls N1 ensuring
that it stays as close as possible to the N1 selected. The pilot can disengage
the functionality at any time by moving the TLA more than 10 degrees. When
disengaged, the FADEC ensures a gradual transition from the N1 cruise
speed control to the N1 speed selected through the TLA.

EFFECTIVITY: ALL
EICAS
87.5 TO 87.5
ATR AIRFRAME AIR DATA REMOTE ENGINE
− WOW
(AVIONICS) FADECs
− FCV CLOSE − ENGINE POSITION
ELECTRICAL AMB. DATA
77.5 N1% 27.4 COMMAND − AIRCRAFT ID
POWER SUPPLY (MACH, P o )
− IGN A ON − MAINT. FAULT RESET
(FADEC, IGNITION,
− IGN B ON − TEST MODE ENABLE
TT0 HEATERS)
−TT0 HEATER INTERNACELLE X TALK (CAN BUS)
IGN IGN
544 ITT C 350
__ __ 28VDC
ENG START/STOP
N2%
RUN RUN

OIL PRES PSI STOP START STOP START
OIL TEMP C
− N1 RED LINE
FADEC CHANNEL A START/ STOP/ IGN 1 2
ENG IGNITION
− N2 RED LINE ON
− ITT RED LINE AUTO
ARINC 429 − N1 MAX OFF

1 2
− THRUST RATING
X TALK
INTEGRATED ARINC 429
AVIONICS UNIT − N1 TARGET
ARINC 429 − N1 REQUEST
STOP
GSD FOR CHANNELS A
GIA FOR CHANNELS B ARINC 429 − N1
− N2 TLA
− ITT RVDTs
− CMC
TT0 HEATER 28VDC
IGNITION 28 VDC
− FUEL FILTER IMP. BYPASS

FADEC CHANNEL B
ENGINE/AIRFRAME ENGINE/AIRFRAME
− OIL PRESSURE UNIT
UNIT
− OIL TEMPERATURE (GEA)
GEA 1 − ENGINE LH
GEA 2 − ENGINE RH
− OIL FILTER IMP. BYPASS
− CHIP DETECTOR
FUEL FLOW COMMAND
BLEED VALVE COMMAND (BVA)

STATIC PRESSURE
ELEC. POWER SUPPLY (PMA) EDCU
SENSOR
SHUTDOWN SOLENOID
ENG SENSORS (N1, N2, T6)
TT0
ELECTRONIC CONTROL SYSTEM - SCHEMATIC DIAGRAM
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
EMERGENCY SHUTDOWN 76-20
Introduction that composes the Emergency Fuel Shutoff Valve (ESOV) Mechanism. When
the disk strikes the plunger it pulls on the mechanism and trips the valve,
The emergency shutdown subsystem furnishes means of controlling the flow causing it to move to the cutoff position. The valve is pressure loaded and will
of fluids to and from the engine during emergency procedures. remain in the cutoff position until a manual reset is performed.
General Description The figure EMERGENCY SHUTDOWN - EMERGENCY FUEL SHUTOFF

VALVE MECHANISM - LOCATION AND OPERATION provides further data
In an emergency situation, the pilot may stop the engine immediately by on the preceding text.
pushing the fire system ENG 1/2 SHUTOFF switches. This action stops the

fuel flow and also stops the bleed air from the engine.
The shaft shear protection is an independent mean of engine shutdown via

emergency shutoff mechanical linkage to an independent emergency fuel
shutoff valve.
Components
The emergency shutdown comprises the components that follow:
• Emergency Fuel Shutoff Valve (ESOV (Emergency Fuel Shutoff Valve))

Mechanism (Piston, Rotate Lever and Cable);
• FMU (Fuel Metering Unit) Integrated Emergency Fuel Shutoff Valve

(ESOV).
Operation
To stop the engine in emergencies, the pilot must push the fire system ENG
1/2 SHUTOFF switch, which commands the valves that follow to close
directly, by energizing their torque motors with DC (Direct Current) power
from the hot busses:
• Engine 1(2) Fuel SOV (Shutoff Valve)
• Engine 1(2) PRSOV (Pressure Regulating and Shutoff Valve) (AMM SDS
36-11-00/1)
In the event of an LP (Low Pressure) shaft failure, the LP turbine moves

rearward and trips a plunger mounted in the exhaust cone. The plunger is
connected through a cable and rod system to the cutoff valve in the FMU,

EFFECTIVITY: ALL
XFR
XXX LB XXX LB

LH ENGINE RH ENGINE
SHUTOFF VALVE SHUTOFF VALVE
(SOV) TOTAL (SOV)
XXXX LB
B USED
XXXX LB

BOTTLE YAW
SHUTOFF 1 SHUTOFF 2
LEFT RIGHT
DISCH
ROLL
OFF LWD RWD MFD
(SYNOPTIC)
ENG START/STOP
RUN RUN
PITCH BKP A
DN
UP
1 2
ENG IGNITION MODE
+
ON BKP
AUTO
OFF OFF
1 2
LEGEND:
FIRE CONTROL PANEL SOV IS OPEN (COLOUR FOLLOWS LINE DOWNSTREAM)
B SOV IS CLOSED (WHITE)
EMERGENCY SHUTDOWN - FIRE SYSTEM SHUTOFF SWITCHES - LOCATION AND INTERFACES
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
NOTES: NOTES:

EFFECTIVITY: ALL
A

PISTON
ROTATE LEVER
SHAFT
SHAFT MOVES BACKWARDS

THE ROTATE LEVER PULLS
DURING SHAFT SHEAR,
THE CABLE CONNECTED
PUSHING THE PISTON
TO ESOV TO SHUT OFF
TO ROTATE LEVER
FUEL FLOW
EMERGENCY FUEL
SHUT OFF MECHANISM
EMERGENCY SHUTDOWN - EMERGENCY FUEL SHUTOFF VALVE MECHANISM - LOCATION AND OPERATION
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80



CHAPTER 77 - ENGINE INDICATING
SECTION TITLE PAGE

77-00 ENGINE INDICATING 256

77-10 POWER 262
77-11 N1 INDICATION 266
77-12 N2 INDICATION 272
77-20 TEMPERATURE 276
77-21 TEMPERATURE INDICATION 278
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
ENGINE INDICATING 77-00
Introduction • Oil pressure
The engine indicating system provides cockpit indications for the flight crew • Oil temperature
and engine operational data for the maintenance crew.
• Engine thrust rating
General Description
• ATR (Automatic Thrust Reserve) status
The ENGINE INDICATING includes these subsystems:
• Cruise speed control status

• POWER (AMM SDS 77-10-00/1)
• TEMPERATURE (AMM SDS 77-20-00/1) • Ignition indication
The rotor speed is monitored and protected by the FADEC (Full Authority
The powerplant indications are displayed on the EICAS (Engine Indication Digital Engine Control) to avoid overspeed both on the ground and in flight.
Crew Alert System), on the left stripe of the center MFD (Multi-Function
The ITT is monitored and protected by the FADEC to avoid overheat during
Display) unit of the cockpit panel. The EICAS provides analog and digital
ground start. When the ITT exceeds the in-flight limits, the information shows
engine indications and icons. The powerplant indications can also be shown
on the EICAS, alerting the flight crew to take action.
on the PFD (Primary Flight Display) in reversionary mode. The CAS (Crew
Alerting System) messages are shown in the CAS window on the PFD and Under normal operating conditions, the pointer and digits are green for each
on the MFD in reversionary mode. parameter. Under abnormal conditions, the pointer and digits change color
accordingly.
The powerplant instruments are closely grouped on the instrument panel. The
location of identical powerplant instruments is so designed as to prevent The engine thrust rating indication is provided by a cyan icon at the top of the
confusion as to which engine each instrument relates. The left engine EICAS. The possible thrust modes are:
indications are shown on the left side of the engine section of the EICAS and
the right engine parameters are shown on the right side. Based on the location • TO - Takeoff
of the instruments referred to above, the powerplant instruments, which are
vital for the safe operation of the airplane, are clearly visible to the crew • GA - Go-around
members.
• CLB - Climb
The EICAS provides the following engine indications:
• CON - Continuous
• N1 (Fan Rotor Speed) (AMM SDS 77-11-00/1)
• CRZ - Cruise
• N2 (Core Rotor Speed) (AMM SDS 77-12-00/1)
The ATR display shows an ATR icon when the ATR is enabled or armed. This
• ITT (Interstage Turbine Temperature) (AMM SDS 77-20-00/1) indication is active in takeoff mode only. The ATR display is positioned below
the thrust rating indication display. When the ATR is enabled, a white
• Fuel flow indication "ATR" shows just below the thrust rating indication display. When

EFFECTIVITY: ALL
NOTES: NOTES:

21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
the ATR is armed (TLA (Thrust Lever Angle) at TOGA (Take off / Go Around) The function of the power indicating system is to provide engine power data
position for takeoff procedure), the "ATR" indication becomes green. In case to the FADEC (Full Authority Digital Engine Control) to perform engine
of an engine failure with the ATR triggered, the ATR indication disappears electronic control and to the EICAS (Engine Indication Crew Alert System)
and the "TO-RSV" indication shows on the thrust rating indication display. for crew information.
The color scheme adopted for the propulsion system warning, caution, and TEMPERATURE (77-20)
advisory indications is shown below:
The function of the temperature indicating system is to monitor the engine
• Red for warning - conditions which require immediate crew awareness temperatures and send the values to the FADEC (Full Authority Digital Engine

and immediate crew action. Control) and the EICAS (Engine Indication Crew Alert System).
• Yellow for caution - conditions which require immediate crew awareness
and subsequent crew action. The figure ENGINE INDICATING - INDICATION provides further data on the
preceding text.
• White for advisory lights - conditions which require crew awareness and
may require subsequent crew action.
The CAS messages related to the engine indicating system are listed in the
table below:
ENGINE INDICATING - CAS MESSAGES (Continued)

INDICATION LEVEL (COLOR) DESCRIPTION
Indicates that a failure
in the TT0 (Inlet Total
E1 TTO HTR FAIL Caution (Yellow) Temperature) sensor
heating circuit for en-
gine 1 was detected.
Indicates that a failure
in the TT0 sensor
E2 TTO HTR FAIL Caution (Yellow)
heating circuit for en-
gine 2 was detected.
Components
POWER (77-10)

EFFECTIVITY: ALL
PARAMETER DISPLAY
DC BUS 1 CAS MESSAGES EXCEEDANCE INDICATIONS CAS MESSAGES EMERGENCY BUS DC BUS 2
CAS WINDOW EICAS CAS WINDOW

(PFD 1) (MFD) (PFD 2)

ARINC 429 DATA
TT0 INPUT FADEC CHANNEL A CONCENTRATOR
UNIT 1 (DCU 1)
N1 INPUT
CAN BUS
(X TALK)
N2 INPUT
T6 INPUT INTEGRATED
ARINC 429
FADEC CHANNEL B AVIONICS UNIT
− N1 RED LINE (GIA 2)
− N2 RED LINE
− ITT RED LINE
− N1 MAX
TT0 N1 − THRUST RATING
N2 T6
ARINC 429
− N1 TARGET
− N1 REQUEST
− N1
− N2
− ITT
− FAULT LOG
− TREND DATA
CMC − EXCEEDANCE DATA
− DISPATCH BITS
ENGINE INDICATING - BLOCK DIAGRAM
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
NOTES: NOTES:

EFFECTIVITY: ALL
B
A
AUTOMATIC THRUST RESERVE INDICATION:
ATR, TO−RSV OR NO INDICATION
THRUST RATING SELECTED INDICATION:
TO, GA, CLB, CON OR CRZ

87.5 TO 87.5
ATR
77.5 N1% 27.4

IGNITION INDICATING:
A, B, AB OR OFF
IGN IGN
__
544 ITT C 350 __
CAS
E1(2) OIL LO PRESS
ENG NO TO DATA 89.6 N2% 45.7
E1(2) FAIL OIL PRES PSI
E1(2) CONTROL FAULT
ENG EXCEEDANCE OIL TEMP C
E1(2) IMP BYPASS FUEL
ENG NO DISPATCH
E1(2) TT0 HTR FAIL
E1(2) TLA FAIL
ENG SHORT DSPTCH
E1(2) FADEC FAULT
EICAS
CAS WINDOW
B
A
ENGINE INDICATING - INDICATION
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
POWER 77-10
Introduction
The function of the power indicating system is to provide engine power data
to the FADEC (Full Authority Digital Engine Control) to perform engine
electronic control and to the EICAS (Engine Indication Crew Alert System)
for crew information.
General Description

The POWER includes these subsystems:
• N1 INDICATION (AMM SDS 77-11-00/1)

• N2 INDICATION (AMM SDS 77-12-00/1)
The power indicating system comprises the N1 (Fan Rotor Speed) sensor
and the N2 (Core Rotor Speed) sensor.
Components
N1 INDICATION (77-11)
The function of the N1 (Fan Rotor Speed) indicating system is to provide

engine thrust data.
N2 INDICATION (77-12)
The function of the N2 (Core Rotor Speed) indicating system is to provide the
engine core rotor speed.
The figure POWER - COMPONENT LOCATION provides further data on the

preceding text.

EFFECTIVITY: ALL
POWER 77-10
A

N1 SENSOR
A
POWER - COMPONENT LOCATION

Sheet 1
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
POWER 77-10
NOTES: NOTES:

EFFECTIVITY: ALL
POWER 77-10

PMA
(N2 SIGNAL)
ACCESSORY
GEAR BOX
(REF.)
FMU
(REF.)
POWER - COMPONENT LOCATION

Sheet 2
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
N1 INDICATION 77-11
Introduction The N1 indication modes are shown below:
The function of the N1 (Fan Rotor Speed) indicating system is to provide • Physical N1 (Analog Indication N1 Trimmed): There is an arc and a pointer
engine thrust data. display representing mechanical N1 speed in %. The pointer is configured
as a green needle and the actual N1 value lower speed quadrant is filled
General Description with grey color. The N1 indication display shows speed values up to 101%
N1. If the FADEC detects an exceedance, the grey portion of the quadrant
The N1 indicating system provides the indication of the engine thrust. It also
will become red. The speed signal is not accurate below 10%. In the event
indicates the target thrust and the maximum thrust available in any given
of loss of the N1 signal, the EICAS removes the pointer from the display

mode of operation. The N1 data is displayed in both analog and digital forms
until a valid signal is received.
and is supplemented with reference bugs for operability.
There is also a digital display representing mechanical N1 speed in %.
The FADEC (Full Authority Digital Engine Control) uses the N1 signal to This is the digital representation of the same data displayed by the analog
control the engine thrust. gauge. The value is displayed with one decimal place. In normal
conditions, the display is green and is reconfigured to show dashes if the
To ensure that both engines supply the same thrust at a fan speed rating, the data is invalid.
FADEC uses a N1 trimmed value for control and indication purposes. The N1
trimmed value is loaded in the EDCU (Engine Data Collector Unit) and shows • N1 Rating (Thrust Rating Max Speed): Is the maximum N1 speed value
in the N1 field of the EICAS (Engine Indication Crew Alert System) for 5 for the current thrust mode. The FADEC provides this information to the
seconds after power cycling. EICAS. The N1 Rating bug is displayed as a T-shaped cyan bug on the
analog N1 gauge.
Components A cyan digital display is provided to indicate the maximum N1 value for
the active thrust rating. This is the digital display of the T-shaped N1 rating
N1 SENSOR bug. The display is positioned above the N1 gauge for each engine.
The N1 sensor is a single magnetic reluctance pickup probe, installed on the • N1 Request (N1 Rating Commanded): N1 Request is the N1 speed value
engine monocase at the 12:00 o'clock position. There are two output coils, requested, based on the current TLA position. The FADEC may limit the
one for each FADEC channel. N1 Request value for some conditions. The difference between the
Physical N1 speed and N1 Request is presented as a white arc and is
The N1 value is trimmed to ensure that both engines supply rated thrust at
shown only during a thrust transient or if the Physical N1 speed cannot
the same TLA (Thrust Lever Angle) position.
reach the N1 Request.
The N1 indication is transmitted to the cockpit display from each FADEC
• N1 Cruise Speed Control: When cruise speed control is engaged, the cyan
channel via serial digital communication bus.
band in the analog N1 gauge will appear. This cyan band represents the
Each N1 signal is shared with the other FADEC channel. Thus, each channel bug indicating N1 authority and system status engaged and active.
receives two independent electrical fan speed inputs.
• N1 Red Line (N1 Transient Red Line): N1 Red Line is the maximum
Operation allowable value for N1, which is the engine operating limit. The display is

EFFECTIVITY: ALL
N1 INDICATION 77-11
NOTES: NOTES:

21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
N1 INDICATION 77-11
a red mark in the N1 gauge. If the limit is exceeded, this value triggers a
color change in both the dial and digital readouts.
• Engine OFF Indication: An indication is provided on the EICAS when an

engine has been shut down by pilot action in flight or on the ground. The
indication comprises the icon "OFF" in black letters in a cyan rectangle in
the center of the associated engine N1 dial.
• Engine Fail Indication: An indication is provided on the EICAS to indicate


when an engine is flamed out or shut down without pilot action. The
indication comprises the icon "FAIL" in black letters in a yellow rectangle
in the center of the associated engine N1 dial. In addition, there is an
associated CAS (Crew Alerting System) "E1(2) FAIL" message on the
CAS window.
The figure N1 INDICATION - DISPLAYS provides further data on the

preceding text.

EFFECTIVITY: ALL
N1 INDICATION 77-11
PARAMETER DISPLAY


ARINC 429 DATA
UNIT 1 (DCU 1)
N1 INPUT
CAN BUS
(X TALK)
N2 INPUT
T6 INPUT INTEGRATED
ARINC 429
− N2 RED LINE
− ITT RED LINE
− N1 MAX
N2 T6
ARINC 429
− N1 TARGET
− N1 REQUEST
− N1
− N2
− ITT
− FAULT LOG
− TREND DATA
− DISPATCH BITS
N1 INDICATION - BLOCK DIAGRAM
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
N1 INDICATION 77-11
NOTES: NOTES:

EFFECTIVITY: ALL
N1 INDICATION 77-11
A B
N1 THRUST
RATING MAX SPEED
N1 CRUISE
SPEED CONTROL
N1 TRANSIENT

87.5 TO 87.5
RED LINE ATR N1 ANALOG DISPLAY
77.5 N1% 27.4 N1 REQUEST ARC
N1 DIGITAL DISPLAY
IGN IGN
__
544 ITT C 350 __
CAS
E1(2) FAIL 89.6 N2% 45.7
OIL PRES PSI
OIL TEMP C
FUEL
EICAS
B
CAS WINDOW
N1 INDICATION - DISPLAYS
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
N2 INDICATION 77-12
Introduction
The function of the N2 (Core Rotor Speed) indicating system is to provide the
engine core rotor speed.
General Description
The N2 indicating system provides indication of the engine core rotor speed
via digital display on the EICAS (Engine Indication Crew Alert System).

The FADEC (Full Authority Digital Engine Control) uses the N2 signal to
control the engine for transient purposes and for idle speed governing.
Components
N2 SPEED SENSOR
The N2 speed signal is provided from a speed sensor installed on the PMA
(Permanent Magnet Alternator). The speed signal is generated by a
frequency output that is proportional to the rotational speed of the PMA. This
speed signal is sent to the FADEC.
Each N2 signal is shared with the other FADEC channel. Thus, each channel
receives two independent electrical core speed inputs.
Operation
The N2 indicating modes are shown as described below:
• Digital Display: The N2 speed indication provides a digital display in %. If

the N2 signal becomes invalid, the display is reconfigured to dashes using
the sign status matrix of the ARINC (Aeronautical Radio Incorporated)
data to indicate faulty data.
• N2 Red Line (Transient Limit): If the N2 transient limit value is exceeded,

a color change in the digital readout is triggered.
The figure N2 INDICATION - N2 COMPONENT LOCATION provides further


EFFECTIVITY: ALL
N2 INDICATION 77-12
PARAMETER DISPLAY


ARINC 429 DATA
UNIT 1 (DCU 1)
N1 INPUT
CAN BUS
(X TALK)
N2 INPUT
T6 INPUT INTEGRATED
ARINC 429
− N2 RED LINE
− ITT RED LINE
− N1 MAX
N2 T6
ARINC 429
− N1 TARGET
− N1 REQUEST
− N1
− N2
− ITT
− FAULT LOG
− TREND DATA
− DISPATCH BITS
N2 INDICATION - BLOCK DIAGRAM
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
N2 INDICATION 77-12
NOTES: NOTES:

EFFECTIVITY: ALL
N2 INDICATION 77-12

PMA
(N2 SIGNAL)
ACCESSORY
GEAR BOX
(REF.)
FMU
(REF.)
N2 INDICATION - N2 COMPONENT LOCATION
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
TEMPERATURE 77-20
Introduction to generate a reference temperature for EGT (T6) thermocouples for use in
several of the FADEC control calculations.
The function of the temperature indicating system is to monitor the engine
temperatures and send the values to the FADEC (Full Authority Digital Engine Components
Control) and the EICAS (Engine Indication Crew Alert System). TEMPERATURE INDICATION (77-21)
General Description The temperature indicating system sensors measure the temperature of
The TEMPERATURE includes this subsystem: many parts of the engine for use in several of the FADEC (Full Authority
Digital Engine Control) control calculations.

• TEMPERATURE INDICATION (AMM SDS 77-21-00/1)
The temperature indicating system comprises the following components:
The temperature indicating system comprises the following sensors for each • TT0 Sensor
engine:
• EGT (T6) Sensors
• TT0 (Inlet Total Temperature) Sensor
• CJC Sensor
• EGT (Exhaust Gas Temperature) (T6) Sensor
The figure TEMPERATURE - COMPONENT LOCATION provides further
• CJC (Cold Junction Compensation) Sensor data on the preceding text.
TT0 SENSOR
The TT0 consists of a single total temperature probe located in the engine
inlet duct and measures the engine inlet air temperature for use in several of
the FADEC control calculations. The PW617F Engine Maintenance Manual
PN 3072162 refers to this sensor as T1 sensor.
EGT (T6) SENSOR
The EGT (T6) sensor consists of a set of six thermocouple temperature

probes extended into the engine gas stream to generate the EGT signals for
use in several of the FADEC control calculations.
CJC SENSOR
The CJC sensor consists of a RTD (Resistance Temperature Detector)

mounted at the end of the engine bypass duct at the 6 o'clock position in order

EFFECTIVITY: ALL
TEMPERATURE 77-20
C

D
B
A B
TERMINAL 2
TERMINAL 1
TTO SENSOR
TERMINAL 3
CJC SENSOR
C D
TEMPERATURE - COMPONENT LOCATION
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
TEMPERATURE INDICATION 77-21
Introduction The TT0 single total temperature probe is located in the engine inlet duct and
measures the engine inlet air temperature for use in several of the FADEC
The temperature indicating system sensors measure the temperature of control calculations.
many parts of the engine for use in several of the FADEC (Full Authority
Digital Engine Control) control calculations. The TT0 temperature is measured by two independent resistance
temperature devices located in a single total temperature probe mounted in
General Description the engine inlet duct. The TT0 temperature signal is input to each FADEC
channel.
The temperature indicating system sensors provide EGT (Exhaust Gas

Temperature) (T6) and inlet total temperature (TT0 (Inlet Total EGT (T6) SENSOR
Temperature)), which are associated with the N1 (Fan Rotor Speed) (AMM
SDS 77-11-00/1) to calculate a value corresponding to the ITT (Interstage The EGT (T6) sensor is a set of probes extended into the engine gas stream
Turbine Temperature). This calculated parameter is used for temperature to generate the EGT signals for use in several of the FADEC control
limiting purposes in the FADEC and for cockpit indication. The PW617F calculations.
Engine Maintenance Manual PN 3072162 refers to TT0 sensor as T1 sensor.
The EGT (T6) is detected by six thermocouple temperature probes that are
The ITT data is displayed in both analog and digital formats. The indication connected in a parallel arrangement prior to being connected to both FADEC
provides a method of detecting engine deterioration or failure conditions. channels.
To ensure that both engines have the same consistent temperature uptrim CJC SENSOR
margin, the FADEC uses a ITT trimmed value for control and indication
purposes. The ITT trimmed value is loaded in the EDCU (Engine Data The CJC sensor is an RTD (Resistance Temperature Detector) mounted in
Collector Unit) and shows in the ITT field of the EICAS (Engine Indication the end of the engine bypass duct at the 6 o'clock position in order to generate
Crew Alert System) for 5 seconds, after power cycling. a reference temperature for EGT (T6) thermocouples for use in several of the
FADEC control calculations.
Components
The FADEC converts the analog electrical signals from EGT (T6)
The components of the temperature indicating system are: thermocouples and from CJC sensor into digital signals and computes the T6
value. This resulting digital signal is cross-communicated to the opposite
• TT0 sensor FADEC via internal CAN bus.
• EGT (T6) sensor Operation
• CJC (Cold Junction Compensation) sensor The ITT indication modes are shown below:
TT0 SENSOR ANALOG INDICATOR
The analog indicator consists of an arc and pointer display representing the
ITT in °C. In case of invalid ITT data, the pointer is removed from the display.

EFFECTIVITY: ALL
NOTES: NOTES:

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EFFECTIVITY: ALL
DIGITAL DISPLAY
The ITT digital display uses the same data source as the analog display and
re-configures the indication to dashes if the data is invalid.
ITT RED LINE (Transient Red Line)
The ITT red line is visible as a red tick mark at the exceedance limit on the
indicator arc. Exceedance of this value triggers a color change to both dial

and digital readouts.
The ITT red line function is to protect the engine capability to achieve
maximum rated thrust.
When the engines are not running and during the restart process, the ITT
start transient limit is displayed.
The EGT (T6) probes are mounted on the turbine case and indicate the
temperature of the combustor gases at the T6 location. Six probes are
connected in parallel and provide an electronic signal that is the average of
the thermocouple probe outputs. The electrical signal is transferred from the
probes to the outside of the engine by a flexible cable.
The figure TEMPERATURE INDICATION - ITT DISPLAY provides further


EFFECTIVITY: ALL
A

87.5 TO 87.5
ATR
ITT ANALOG
ITT TRANSIENT 77.5 N1% 27.4 DISPLAY
RED LINE
ITT DIGITAL
DISPLAY
IGN IGN
__
544 ITT C 350 __
89.6 N2% 45.7

OIL PRES PSI
OIL TEMP C
FUEL
EICAS
TEMPERATURE INDICATION - ITT DISPLAY
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80



CHAPTER 78 - EXHAUST
SECTION TITLE PAGE

78-00 EXHAUST 284

78-10 EXHAUST NOZZLE 286
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
EXHAUST 78-00
Introduction
The exhaust system consists of those components which direct the exhaust
gases overboard.
General Description
The EXHAUST includes this subsystem:


• EXHAUST NOZZLE (AMM SDS 78-10-00/1)
The exhaust system is in the aft region of the nacelle and includes the
collector/nozzle subsystem. This subsystem has an exhaust nozzle, a
centerbody and an aft body. The exhaust nozzle and the centerbody are
located in the aft body compartment that is in the aft region of the nacelle.
Components
EXHAUST NOZZLE (78-10)
The collector/nozzle subsystem has the function of directing the flow of gases
overboard as efficiently as possible.
The figure EXHAUST - COMPONENT LOCATION provides further data on

the preceding text.

EFFECTIVITY: ALL
EXHAUST 78-00
CENTERBODY
EXHAUST NOZZLE
AFT BODY

EXHAUST - COMPONENT LOCATION
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
EXHAUST NOZZLE 78-10
Introduction Exhaust Case). The hot air flow that comes from engine exhaust is mixed
with a cold bypassed air flow in the aft body.
The collector/nozzle subsystem has the function of directing the flow of gases
overboard as efficiently as possible. The aft body compartment is not considered a fire zone, due to the fact that
there are no flammable fluid line and components in these compartments and
General Description in the outer aft body area that can create a hazard condition. This
compartment is also ventilated to prevent points of excessively high
The exhaust nozzle primary function is to expand the exhaust gases of the temperature. However, the aft body is fire resistant, with an aft bulkhead that
engine to provide as much thrust as possible. provides segregation from fire zone.

The centerbody directs the exhaust gases that flow to the aft body in a The figure EXHAUST NOZZLE - COMPONENT LOCATION provides further
controlled and shielded environment before going out to the atmosphere. data on the preceding text.
The main function of aft body is to incorporate an end-nozzle to meet the
performance requirements. Its secondary function is to aerodynamically flare
ambient air around the engine. The aft body also allows the relative engine
movement due to thermal expansion.
Components
EXHAUST NOZZLE
The exhaust nozzle, which has a chevron design, reduces the noise levels
from the engine by mixing core exhaust gases with bypass air.
CENTERBODY
The centerbody assembly consists of a forward flange and a sheet metal

conical section. The conical section has internal stiffeners that are attached
to the forward flange. The forward flange is attached to the hub flange of the
engine turbine frame.
AFT BODY
The aft body was developed with the inner and outer nozzle walls, the
attachment flange, L and Z sections and the aft bulkhead.
The aft body is not submitted to high air temperatures because it is not directly
subjected to the hot gases that pass through the engine TEC (Turbine

EFFECTIVITY: ALL
CENTERBODY
EXHAUST NOZZLE
A

OUTER NOZZLE WALL
INNER NOZZLE WALL
ATTACHMENT
FLANGE
SECTIONS
AFT BULKHEAD A
EXHAUST NOZZLE - COMPONENT LOCATION
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
The inner nozzle wall is composed by the upper and lower walls,
reinforcement, and splices. The upper and lower walls and reinforcement are
formed from ALCLAD 6061-T62 sheet. The reinforcement is attached to the
lower wall by the solid rivets. The splices are formed from ALCLAD 2024-T42
sheet. The upper and lower walls are joined by the splices with solid rivets.
The outer nozzle wall is composed of the right and left walls and splices.
These components are formed from ALCLAD 2024-T3 sheet. The right and
left walls are joined by the splices with solid rivets.

The attachment flange is manufactured from AL 7050-T7541 plate and it is
attached to the inner and outer nozzle walls with the solid rivets.
The figure EXHAUST NOZZLE - COMPONENT LOCATION provides further


EFFECTIVITY: ALL
A B

SPLICE
SPLICE
UPPER INNER
WALL C
D
ATTACHMENT
FLANGE
E SPLICE
LOWER INNER E LEFT OUTER
WALL
RIGHT OUTER WALL
WALL
REINFORCEMENT
SPLICE
B
C
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
The aft bulkhead and L sections are formed from Titanium sheet. The L
sections join the upper and lower walls to the aft bulkhead by the solid rivets.
The Z sections are formed from ALCLAD 2024-T42 sheet. The Z sections
join the inner walls to the outer walls by solid rivets.
The figure EXHAUST NOZZLE - COMPONENT LOCATION provides further


EFFECTIVITY: ALL
A B
C

BULKHEAD
A
Z SECTION
L SECTION
B C
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80



CHAPTER 79 - OIL
SECTION TITLE PAGE

79-00 OIL 294

79-10 STORAGE 302
79-11 OIL TANK SYSTEM 304
79-20 DISTRIBUTION 306
79-21 OIL FILTER 312
79-30 INDICATING 314
79-31 OIL TEMPERATURE/PRESSURE INDICATION 322
79-34 OIL FILTER BYPASS WARNING INDICATION 326
79-35 CHIP DETECTOR INDICATION 328
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
OIL 79-00
Introduction scavenge elements to remove oil from the bearing chambers and return it to
the tank. The oil filter and electrical monitoring sensors are combined in an
The function of the engine oil system is to provide lubrication and cooling of oil filter module, mounted on the left side of the oil tank. The electrical chip
the engine turbine main shaft bearings and AGB (Accessory Gearbox) detector/collector also mounts on the bottom of the AGB. The FOHE (Fuel-
internal components and bearings. Oil Heat Exchanger) is separately mounted on its own brackets and cools the
oil from the supply pump before it is routed to the bearing chambers and AGB.
General Description
For further information about the engine, refer to the latest revision of the Pratt
The OIL includes these subsystems: & Whitney Engine Manual.

• STORAGE (AMM SDS 79-10-00/1) The CAS (Crew Alerting System) messages related to the engine oil system
• DISTRIBUTION (AMM SDS 79-20-00/1) are listed in the table below:
• INDICATING (AMM SDS 79-30-00/1)
OIL - CAS MESSAGES (Continued)
Each PW617F engine has an independent lubrication supply system which INDICATION LEVEL (COLOR) DESCRIPTION
uses an engine-driven positive displacement vane type pump element to Indicates that low oil
supply oil to the different engine components requiring cooling and E1 OIL LO PRESS Warning (Red) pressure is detected
lubrication. The lubrication system is a self contained pressurized full flow on the engine 1.
system. There are three independent bearing chambers in the engine that
require lubrication: Indicates that low oil
E2 OIL LO PRESS Warning (Red) pressure is detected
• The number 1 and 2 LP (Low Pressure) fan thrust bearings, and the on the engine 2.
number 3 HP (High Pressure) roller bearing are located in the same
chamber that is sealed at the cold end by labyrinth seals and is externally Components
pressurized by compressor discharge air.
STORAGE (79-10)
• The number 4 HP turbine roller bearing is located in a chamber in front of
the HP turbine. The chamber is sealed by two carbon seals externally The storage system supplies the oil from the oil tank to the inlet of oil pump.
pressurized by compressor discharge air.
DISTRIBUTION (79-20)
• The number 5 LP turbine roller bearing is located in a chamber aft of the
LP turbine. It is housed within the turbine exhaust case, in the engine LP The oil distribution system supplies oil for engine bearing lubrication and
turbine module. This chamber is sealed by a single carbon seal, and is cooling. Lubricating oil is filtered, cooled and then sent to the bearing
externally pressurized by compressor discharge air. chambers for bearing lubrication. The system also removes the oil from the
bearing chambers, checks for particle contamination and removes the air
The AGB holds and provides drive pads for the engine oil pump assembly, before returning the oil to the tank.
fuel pump assembly and a starter/generator. The lubrication and scavenge
pump supplies oil to all bearings and gears as required, and includes INDICATING (79-30)

EFFECTIVITY: ALL
OIL 79-00
NOTES: NOTES:

21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
OIL 79-00
The oil indicating system gives an indication of oil level, pressure and Basically, the system pulls oil from the oil tank, pressurized by the oil pressure
temperature, and metal debris presence in the oil. pump, and sends this oil to the filter, to the heat exchanger for cooling, and
then to the engine bearings.
The PW617F engine lubrication system has the following components: The scavenge oil is removed from the bearing chambers to the AGB by the
scavenge elements of the oil pump. Afterwards the oil flows through the chip
• Oil tank with a filler neck and a sight glass oil level indicator.
detector/collector and then it is scavenged by the AGB scavenge pump to the
• ACOC (Air-Cooled Oil Cooler) with a pressure and a thermal bypass tank.
valves.

Vent air is removed from the sumps along with the bearing chambers
• MOPT (Main Oil Pressure and Temperature) sensor. scavenge oil, sent to the air/oil separator on the AGB and vented overboard
through the engine exhaust.
• Breather system.
• Oil Pump.
Access to the engine oil system is provided by opening the oil inspection/
• Oil PAV (Pressure Adjusting Valve)/CSV (Cold Start Valve) assembly. servicing door, located in the engine lower mid cowl, which can also be
opened. Opening the door allows access to the oil filler neck and sight glass
• Electrical chip detector/collector. oil level indicator. There is an oil impending bypass popup door to check the
status of the oil impending bypass indication.
• Oil filter module with a bypass valve and an impending bypass indicator.
Each of the scavenge elements of the scavenge pump includes a wire-mesh
• FOHE.
screen located in the AGB that is accessible from the front side of the AGB.
• Restrictor.
An electric master chip detector and a self-closing valve with debris screen
• Strainers. are located in the scavenge return line on the AGB.
Operation The oil tank fill system includes the following features:
The functions of the lubrication system are given below: • Filler port accessible through dedicated door on lower mid cowl.
• Oil storage and supply. • Oil filler and oil tank level indication accessible through the oil servicing
door.
• Pressurization and vent.
The figure OIL - COMPONENT LOCATION provides further data on the
• Heat and contamination removal. preceding text.
• Lubrication/protective barrier against wear and corrosion of internal

components.

EFFECTIVITY: ALL
OIL 79-00
STRAINER STRAINER
STRAINER
#1 BRG #2 BRG #3 BRG #4 BRG #5 BRG


#5 SCAVENGE
IMPENDING PUMP
THERMAL
MOPT BYPASS
BYPASS
SENSOR POP−UP
VALVE
INDICATOR
PRESSURE PAV / CSV

BYPASS BYPASS
VALVE AGB
OIL
FOHE ACOC AIR−OIL
RESTRICTOR OIL FILTER PRESSURE STRAINER OIL TANK
SEPARATOR
PUMP
AGB SCAVENGE PUMP STRAINER CHIP DETECTOR
LEGEND:
SUPPLY LINE BYPASS AIR LINE
SCAVENGE LINE FUEL LINE
SUMP VENT LINE
OIL - SCHEMATIC DIAGRAM
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EFFECTIVITY: ALL
OIL 79-00
NOTES: NOTES:

EFFECTIVITY: ALL
OIL 79-00

IMPENDING
BYPASS
BYPASS
VALVE
INDICATOR OIL FILTER
OIL TANK PAV/CSV

OIL SIGHT ASSEMBLY
GLASS
BREATHER
LINE
OIL FILLER
A OIL
PUMP
CHIP DETECTOR/
COLLECTOR
OIL - COMPONENT LOCATION

Sheet 1
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EFFECTIVITY: ALL
OIL 79-00
NOTES: NOTES:

EFFECTIVITY: ALL
OIL 79-00
ACOC

ACESSORY
GEAR BOX
MOPT SENSOR FOHE
OIL - COMPONENT LOCATION

Sheet 2
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EFFECTIVITY: ALL
STORAGE 79-10
Introduction oil also flows through the FOHE (Fuel-Oil Heat Exchanger), which basically
is used for fuel heating and oil cooling.
The storage system supplies the oil from the oil tank to the inlet of oil pump.
The oil, including AGB lubrication oil, is then drawn by the AGB scavenge
General Description pump and returned to oil tank. The air mixed with the oil in the AGB is
separated by an air/oil separator which is vented to the engine exhaust duct,
The STORAGE includes this subsystem: through the breather tube.
• OIL TANK SYSTEM (AMM SDS 79-11-00/1) With the engine inoperative, all the oil from system returns to the oil tank,

what allows a trustful check of oil level through the oil sight glass.
The storage system holds the oil for the engine and includes the oil tank that The oil tank with integral AGB and lubrication system components are
is integral to the AGB (Accessory Gearbox). fireproof.
Components The figure STORAGE - BLOCK DIAGRAM provides further data on the
preceding text.
OIL TANK SYSTEM (79-11)
The oil tank is a cast housing that is an integral part of the AGB (Accessory
Gearbox), providing storage of the lubricating oil.
OIL TANK SYSTEM (79-11)
The oil tank system provides storage of the lubricating oil.
The storage system includes the components listed below:
• Oil filler cap fitted with an oil tight seal and locked down by an over center
lever.
• Oil filler neck with a piston valve to limit oil exiting the tank.
• Sight glass for visual oil level indication.
• Drain plug.
Operation
The oil that circulates through the engine, pumped by the oil pressure pump,
is mixed with the air existing in the system, deriving from the sealing of the
bearing chambers, which are pressurized by a compressor discharge air. This

EFFECTIVITY: ALL
STORAGE 79-10
AGB SCAVENGE
PUMP

#1, #2, #3
BEARINGS
OIL
AGB PRESSURE
PUMP
OIL
#4 BEARING TANK
AIR OIL
SEPARATOR OIL
FILLER
#5 BEARING
ENGINE
EXAUST FOHE
LEGEND
SCAVENGE LINE
SUPPLY LINE
SUMP VENT LINE
STORAGE - BLOCK DIAGRAM
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
OIL TANK SYSTEM 79-11
Introduction
The oil tank is a cast housing that is an integral part of the AGB (Accessory
Gearbox), providing storage of the lubricating oil.
General Description
The oil tank maximum capacity is 3.92 . The minimum usable oil quantity
allowable without adversely affecting the operation of the engine is 2.98 .

These values are for the worst allowable aircraft attitude of 2 degrees on the
ground.
The tank has sufficient oil to provide operation during a 10 h mission at the
maximum oil consumption of 0.068 /h. For the oil level at the minimum
servicing level, the oil is sufficient for a 5 h mission, considering the maximum
oil consumption.
The oil tank also incorporates a drain plug located in the bottom of the oil tank
to provide the oil drainage. The drain plug is accessible with standard tools
and incorporates a safety cable. The oil tank draining is also provided through
the removal of the chip detector/collector self-closing valve (AMM SDS
79-35-00/1), located in the bottom of the AGB.
The figure OIL TANK SYSTEM - COMPONENT LOCATION provides further


EFFECTIVITY: ALL
OIL TANK SYSTEM 79-11
B

A
A B
VENT LINE C
AGB
DRAIN
PLUG
OIL TANK
C
B−B
OIL TANK SYSTEM - COMPONENT LOCATION
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
DISTRIBUTION 79-20
Introduction OIL BREATHER SYSTEM
The oil distribution system supplies oil for engine bearing lubrication and The oil breather system includes the air/oil separator and the breather line
cooling. Lubricating oil is filtered, cooled and then sent to the bearing and its function is to prevent excessive air pressure in the bearing
chambers for bearing lubrication. The system also removes the oil from the compartment so that the flow of oil to the bearings and the operation of the
bearing chambers, checks for particle contamination and removes the air scavenge system is not impaired. The air/oil separator is a centrifugal type
before returning the oil to the tank. breather that has the function of separating the air from the aerated scavenge
oil. It is located in the AGB. The breather line is a tube that vents to the bypass
General Description duct just upstream of the mixer. It is arranged so that condensed water vapor

or oil that might freeze and obstruct the line cannot accumulate at any point.
The DISTRIBUTION includes this subsystem:
Components
• OIL FILTER (AMM SDS 79-21-00/1)
OIL FILTER (79-21)
The PW617-F engine oil distribution system has the following topics: The oil filter module contains a filter through which all lubrication oil must
pass. The module is also equipped with an impending bypass valve and a
OIL SUPPLY SYSTEM
mechanical popup impending bypass indicator.
The oil supply system supplies pressurized clean cooled oil to the engine
bearings and accessory drives. Oil from the tank passes through a strainer The PW617F engine oil distribution system has the following components:
and then is pumped by the oil pressure pump that has a PAV (Pressure
Adjusting Valve)/CSV (Cold Start Valve) assembly. From this pressure • Oil tank, integral with AGB.
element, the oil passes through the filter, ACOC (Air-Cooled Oil Cooler) and
FOHE (Fuel-Oil Heat Exchanger), in this order. Then, the oil flow is divided • MOPT (Main Oil Pressure and Temperature) sensor.
into several circuits and passes through a strainer to finally lubricate the 1st,
• Oil filter module.
2nd and 3rd bearing chambers.
• Oil filter bypass valve.
OIL SCAVENGE SYSTEM
• Oil filter impending bypass indicator.
The oil scavenge system collects oil from the bearing compartments and
scavenges to the AGB (Accessory Gearbox). Oil from the 1st bearing • Electrical chip detector/collector.
chamber, that houses the numbers 1, 2 and 3 bearings, is gravity scavenged.
Oil from the 2nd bearing chamber, that houses the number 4 bearing, is • PAV.
blowdown scavenged. Oil from the 3rd bearing chamber, that houses the
number 5 bearing, is pump scavenged. The oil, including AGB lubrication oil, • CSV.
is then drawn by a dedicated pump past a chip detector/collector, returning
• Oil pump.
to the tank.

EFFECTIVITY: ALL
DISTRIBUTION 79-20
NOTES: NOTES:

21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
DISTRIBUTION 79-20
• FOHE. The FOHE is a heat transfer matrix mounted on the front of the AGB used for
fuel heating and oil cooling. The fuel side of the FOHE is positioned in the
• ACOC. fuel system between low and high pressure pumps, upstream the fuel filter
and the oil side is in the oil feed line of the lubrication system, downstream
For more information about oil tank, refer to AMM SDS 79-11-00/1.
the oil filter.
For more information about MOPT sensor, refer to AMM SDS 79-31-00/1.
The FOHE is equipped with a fuel filter housing, so designed that the fuel
For more information about oil filter module, bypass valve and impending flows from the outer to the inner diameter of the filter. There are also
bypass indicator, refer to AMM SDS 79-21-00/1. provisions for the mounting of an impending bypass switch and bypass

indicator that are so mounted that the pressure drop and peak pressure
For more information about chip detector/collector, refer to AMM SDS across the filter can be monitored. A fuel bypass valve allows fuel to bypass
79-35-00/1. the filter if the pressure drop across the filter exceeds the cracking pressure
of the valve.
PAV
ACOC
The PAV is located on the oil pressure pump housing and composes an
assembly with the CSV. The PAV is used to set the oil system pressure to a The ACOC is a plates and fins type of heat transfer matrix used for air heating
preset value for specified compressor speed and oil temperature, bypassing and oil cooling. A temperature actuated valve combined with a pressure relief
oil from the pump outlet back to the pump inlet. valve is used in order to bypass the ACOC matrix under cold oil conditions
or under blocked matrix.
CSV
The figure DISTRIBUTION - COMPONENTS LOCATION provides further
The CSV is located immediately at the oil pressure pump outlet and provides data on the preceding text.
a safeguard against excessive pressure buildup due to downstream system
blockage or high oil viscosity in cold weather operation. When open, it diverts
oil from the pump outlet back to the pump inlet.
OIL PUMP
The oil pump is mounted next to the accessory gearbox, behind the oil filter.
It is a positive displacement pump with a series of pumping elements (one
pressure and two scavenge ones) that are put in series on a common drive
shaft. The oil pressure element supplies the oil to the engine and is fed from
the oil tank. The number 5 scavenge element returns the oil to the AGB and
the AGB scavenge element returns the oil from the AGB back to the oil tank
past the chip detector.
FOHE

EFFECTIVITY: ALL
DISTRIBUTION 79-20

IMPENDING
BYPASS
BYPASS
VALVE
INDICATOR OIL FILTER
OIL TANK PAV/CSV

ASSEMBLY
BREATHER
SYSTEM
A OIL
PUMP
A CHIP DETECTOR /
COLLECTOR
DISTRIBUTION - COMPONENTS LOCATION

Sheet 1
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
DISTRIBUTION 79-20
NOTES: NOTES:

EFFECTIVITY: ALL
DISTRIBUTION 79-20
ACOC

ACESSORY
GEAR BOX
(REF.)
FOHE
DISTRIBUTION - COMPONENTS LOCATION

Sheet 2
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
OIL FILTER 79-21
Introduction The oil pressure pump has the engine lubrication supply element and two
scavenge elements. Oil from the tank enters the supply element of the oil
The oil filter module contains a filter through which all lubrication oil must pressure pump. From this pressure element, the oil passes through the filter
pass. The module is also equipped with an impending bypass valve and a module.
mechanical popup impending bypass indicator.
The oil filter has a bypass valve, which permits oil flow to the engine if the
General Description filter becomes clogged. The filter has also a mechanical popup impending
bypass indicator. For further information about the oil impending bypass
The oil filter is located in the lubrication filter module, on the side of the AGB
indicator, refer to AMM SDS 79-34-00/1.

(Accessory Gearbox), schematically in the lubrication supply line between the
oil pressure pump and ACOC (Air-Cooled Oil Cooler). The figure OIL FILTER - COMPONENT LOCATION provides further data on
the preceding text.
The oil filter cannot be cleaned, and must be replaced at previously defined
intervals. The access for maintenance purposes is by removal of the bottom
filter cover.
The oil filter module contains an impending bypass valve as a safeguard

against filter blockage. The bypass is located above the oil filter and FOHE
(Fuel-Oil Heat Exchanger) assembly so that hot oil flow to the filter encounters
this passage before flowing down into and around the cylindrical oil filter.
In the oil filter module there is also a mechanical popup impending bypass
indicator with a colored button that pops up to give a visual indication that the
filter needs to be replaced. An impending bypass indication is provided prior
to bypass activation.
Components
The oil filter module includes the components listed below:
• Oil filter element.
• Oil filter housing.
• Oil bypass valve.
• Mechanical popup impending bypass indicator.
Operation

EFFECTIVITY: ALL
OIL FILTER 79-21
MECHANICAL
POP−UP IMPENDING
BYPASS INDICATOR
OIL BYPASS VALVE


OIL FILTER
ACCESSORY
GEAR BOX
OIL FILTER - COMPONENT LOCATION
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
INDICATING 79-30
Introduction The mechanical oil filter impending bypass indicator is installed into the oil
filter housing. It is activated by the excessive oil filter pressure across the
The oil indicating system gives an indication of oil level, pressure and filtering element.
temperature, and metal debris presence in the oil.
CHIP DETECTOR INDICATION (79-35)
General Description
The function of the electrical chip detector/collector is to attract and trap
The INDICATING includes these subsystems: magnetic particles that are suspended in the scavenge oil because it may be
an indication of an impending failure. This is achieved with the use of a

• OIL TEMPERATURE/PRESSURE (AMM SDS 79-31-00/1) permanent magnet immersed in the scavenge oil flowing from the AGB
INDICATION (Accessory Gearbox), before it passes through the AGB scavenge pump. The
• OIL FILTER BYPASS WARNING (AMM SDS 79-34-00/1) chip detector/collector can also function as a drain of the oil tank.
INDICATION
• CHIP DETECTOR INDICATION (AMM SDS 79-35-00/1)
The oil indicating system includes the following components:
There is an oil level indicator for each engine mounted externally to each oil • Oil level indicator.
tank with maximum and minimum level indications. Oil temperature and
pressure indications are also provided for each engine and displayed in the • Oil filter impending bypass indicator.
cockpit on the engine indication field on the EICAS (Engine Indication Crew
• Chip detector/collector.
Alert System). A warning message is provided in the CAS (Crew Alerting
System) window on the PFD (Primary Flight Display) in case of low oil • MOPT (Main Oil Pressure and Temperature) sensor.
pressure. An electric master chip detector and a self-closing valve are located
in the scavenge return line in both oil tanks, where ferromagnetic particles Operation
are most likely to be deposited.
The operation of the sensors is described below:
Components
OIL LEVEL INDICATOR
OIL TEMPERATURE/PRESSURE INDICATION (79-31)
The oil tank level indicator is a vertical sight glass that enables to see the
The oil temperature and pressure indications in the cockpit are provided by amount of oil in the tank. The indicator also shows the maximum and
the MOPT (Main Oil Pressure and Temperature) sensor that incorporates the minimum acceptable levels for oil. It is mounted externally to the oil tank to
two functions. This sensor is mounted on the AGB (Accessory Gearbox), make it possible to view the oil level.
downstream the FOHE (Fuel-Oil Heat Exchanger).
OIL FILTER IMPENDING BYPASS INDICATOR
OIL FILTER BYPASS WARNING INDICATION (79-34)
The oil filter impending bypass indicator is installed on the oil filter and is
equipped with a button that pops up to indicate that the oil filter must be
replaced.

EFFECTIVITY: ALL
INDICATING 79-30
NOTES: NOTES:

21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
INDICATING 79-30
For more information about the oil filter impending bypass indicator, refer to
AMM SDS 79-34-00/1.
CHIP DETECTOR/COLLECTOR
The chip detector/collector is installed on the aft face of the AGB (Accessory
Gearbox) and its function is to trap magnetic particles that are suspended in
the scavenge oil. This is accomplished with the use of a permanent magnet
immersed in the scavenge oil flowing from the oil pressure pump, returning

to the oil tank.
The detector also provides a CMC (Central Maintenance Computer)

message when magnetic debris buildup has reached an unacceptable level.
For more information about the oil filter impending bypass indicator, refer to
AMM SDS 79-35-00/1.
MOPT SENSOR
The MOPT sensor is mounted downstream of the FOHE (Fuel-Oil Heat

Exchanger) and its function is to provide electrical outputs for pressure and
temperature values.
The pressure measured by the MOPT sensor is the differential between the
oil filter output and the AGB. This pressure differential deflects a silicon
diaphragm that has resistors on its surface, changing the resistance
proportionally and sending an electrical signal to the aircraft.
The principle of operation for temperature measurement is based on the RTD

(Resistance Temperature Detector) technology.
The sensor sends a signal to the cockpit that displays the current oil pressure
and temperature status in the engine indication field on the EICAS.
The figure INDICATING - SCHEMATIC DIAGRAM provides further data on

the preceding text.

EFFECTIVITY: ALL
INDICATING 79-30

MOPT SENSOR
ACESSORY
GEAR BOX
(REF.)
INDICATING - COMPONENT LOCATION
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
INDICATING 79-30
NOTES: NOTES:

EFFECTIVITY: ALL
INDICATING 79-30
GEA 1 OIL CHIP DETECTOR 1
HI
ENG 1 CHIP DET
INPUT CHIP
LO
SIGNAL GROUND

MOPT SENSOR 1
OIL PRESSURE SENSOR

HI
OIL PRESS EXC
OUTPUT
LO
HI
OIL PRESS SIG
INPUT
LO
HI
ENG 1 OIL TEMP
INPUT OIL TEMP SENSOR
LO
SIGNAL GROUND

Sheet 1
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
INDICATING 79-30
NOTES: NOTES:

EFFECTIVITY: ALL
INDICATING 79-30
GEA 2 OIL CHIP DETECTOR 2
HI
ENG 2 CHIP DET
INPUT CHIP
LO

SIGNAL GROUND
MOPT SENSOR 2
OIL PRESSURE SENSOR

HI
OIL PRESS EXC
OUTPUT
LO
HI
OIL PRESS SIG
INPUT
LO
HI
ENG 2 OIL TEMP
INPUT OIL TEMP SENSOR
LO
SIGNAL GROUND

Sheet 2
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
OIL TEMPERATURE/PRESSURE INDICATION 79-31
Introduction The oil pressure is indicated in psig and the oil temperature in Celsius degrees
on the EICAS, both via a digital display.
The oil temperature and pressure indications in the cockpit are provided by
the MOPT (Main Oil Pressure and Temperature) sensor that incorporates the In normal conditions, the oil temperature and oil pressure will be displayed in
two functions. This sensor is mounted on the AGB (Accessory Gearbox), green. When the oil pressure is out of the normal range but within the steady
downstream the FOHE (Fuel-Oil Heat Exchanger). state limit, the oil indication on EICAS will be displayed in amber inverse
video. If the oil pressure exceeds the transient limit, the oil indication on
General Description EICAS will be displayed in red inverse video.

The purpose of the MOPT sensor is to provide electrical outputs for pressure The oil temperature indication works in the same way. For normal range
and temperature values. within the steady state limit, the oil temperature indication will be displayed
in amber inverse video and if the oil temperature exceeds the transient limit,
The pressure sensor principle of operation is based on SOI (Silicon on
the oil indication on EICAS will be displayed in red inverse video.
Insulator) technology and the temperature sensor principle of operation is
based on RTD (Resistance Temperature Detector) technology. The figure OIL TEMPERATURE/PRESSURE INDICATION - COMPONENT
The sensor sends a signal to the cockpit that displays the current oil pressure
and temperature status in the engine indication field on the EICAS (Engine
Indication Crew Alert System).
For low pressure condition, warning messages are shown in the CAS (Crew
Alerting System) window on the PFD (Primary Flight Display).
E1 OIL LO Warning The oil pressure indication for LH (Left-

PRESS (Red) Hand) engine is lower than the set point.
The oil pressure indication for RH
E2 OIL LO Warning
(Right-Hand) engine is lower than the
PRESS (Red)
set point.
Operation
The MOPT sensor is fed by an oil tapping on the AGB. The oil pressure and
temperature signals from the sensor are sent directly to the airframe avionics
system that supplies this data to the FADEC (Full Authority Digital Engine
Control). The FADEC is responsible for monitoring and comparing the engine
oil temperature and pressure with the transient and steady state limits,
commanding display color changes in case of an exceedance occurrence.

EFFECTIVITY: ALL

MOPT SENSOR
ACESSORY
GEAR BOX
(REF.)
OIL TEMPERATURE/PRESSURE INDICATION - COMPONENT LOCATION
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
NOTES: NOTES:

EFFECTIVITY: ALL
A

87.5 TO 87.5
ATR
77.5 N1% 27.4
IGN IGN
__
544 ITT C 350 __
N2%
56 OIL PRES PSI 57

ENGINE OIL
98 OIL TEMP C 93 PRESSURE
ENGINE OIL FUEL
TEMPERATURE
EICAS
OIL TEMPERATURE/PRESSURE INDICATION - COMPONENT LOCATION
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
OIL FILTER BYPASS WARNING INDICATION 79-34
Introduction
The mechanical oil filter impending bypass indicator is installed into the oil
filter housing. It is activated by the excessive oil filter pressure across the
filtering element.
General Description
The oil filter impending bypass indicator is a mechanical device that has a

colored button that pops up to provide a visual indication that the filter needs
to be replaced.
When the filter becomes clogged, the differential pressure between the filter
inlet and the filter outlet starts to increase. When this pressure reaches 22 ±
2 psid (151 ± 13 kPa), the button pops up. After actuation, the indicator has
to be manually reset.
When the differential pressure reaches 38 ± 2 psid (262 ± 13 kPa), the oil
bypass valve opens and the oil stops passing through the filter.
The indicator incorporates a bimetallic strip that will shrink at oil temperatures
below 100 °F (38 °C), thus preventing inadvertent actuation when the oil is
cold and viscous. At oil temperatures equal to or greater than 140 °F (60 °C),
the bimetallic strip becomes large enough to allow free movement of the
indicator.
The maximum operational temperature limit for oil filter impending bypass
indicator is 149 °C (300 °F).
The figure OIL FILTER BYPASS WARNING INDICATION - COMPONENT


EFFECTIVITY: ALL
OIL FILTER BYPASS WARNING INDICATION 79-34
MECHANICAL
POP−UP IMPENDING
BYPASS INDICATOR
OIL BYPASS VALVE


OIL FILTER
ACCESSORY
GEAR BOX
OIL FILTER BYPASS WARNING INDICATION - COMPONENT LOCATION
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
CHIP DETECTOR INDICATION 79-35
Introduction
The function of the electrical chip detector/collector is to attract and trap

magnetic particles that are suspended in the scavenge oil because it may be
an indication of an impending failure. This is achieved with the use of a
permanent magnet immersed in the scavenge oil flowing from the AGB
(Accessory Gearbox), before it passes through the AGB scavenge pump. The
chip detector/collector can also function as a drain of the oil tank.

General Description
The chip detector/collector is located in the bottom of the AGB. It can be easily
removed for inspection or to drain the oil without any other disassembly. A
self-closing valve seals the scavenge return line when the chip detector/
collector is removed to prevent oil losses and low oil pressure.
Operation
The basic operation of the electrical chip detector/collector is the following:

the scavenge return oil is directed through the chip detector/collector which
has a powerful magnet that separates two electrical terminals. This magnet
attracts and holds ferrous metal pieces and keeps them from going back to
the oil system. When the metal pieces bridge the gap between the two
terminals, an electrical circuit is closed and a signal is sent to the CMC
(Central Maintenance Computer), alerting against impending distress of a
metallic component in the engine.
The figure CHIP DETECTOR INDICATION - COMPONENT LOCATION


EFFECTIVITY: ALL
CHIP DETECTOR INDICATION 79-35

B
CHIP DETECTOR/COLLECTOR ACCESSORY GEARBOX
B A
CHIP DETECTOR INDICATION - COMPONENT LOCATION
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80



CHAPTER 80 - STARTING
SECTION TITLE PAGE

80-00 STARTING 332

80-10 CRANKING 338
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
STARTING 80-00
Introduction FLAMEOUT DETECTION / AUTO RELIGHT
The starting system function is to initiate the engine operation. In a flameout situation, both igniters are automatically sequenced ON by the
FADEC when the N2 speed drops and the requested fuel flow increases. If
General Description the engine does not relight, then the igniters and fuel flow remain ON until the
pilot sets the ENG START/STOP switch to the STOP position.
The STARTING includes this subsystem:
WET and DRY MOTORING
• CRANKING (AMM SDS 80-10-00/1)

Wet and dry motoring procedures required for maintenance and purging
The control system provides automatic control of fuel flow, ignition and purposes can be achieved by the correct selection of input switches to the
protection of the engine during the starting phase. FADEC.
During engine starting phase, the starter drives the engine by rotating the The wet motoring procedure is performed by setting the ENG IGNITION
high pressure shaft up to 44% N2 (Core Rotor Speed). At this point, the switch to the AUTO/ON position while the ENG START/STOP switch is at
FADEC (Full Authority Digital Engine Control) sends the cut-off signal to the START and the Ignition Circuit Breakers is open.
GCU (Generator Control Unit), which disconnects the starter from the AGB
(Accessory Gearbox) and connects the generator to the DC Bus. The dry motoring procedure is performed by setting the ENG IGNITION
switch to the OFF position, while the engine is in shutdown state, and by
There are four distinct starting modes: ground start, air start, wet and dry engaging the starter. The motoring procedure may be aborted at any time by
motoring. The starting modes are based on the inputs from the ENG START/ setting the ENG START/STOP switch to the OFF position.
STOP and ENG IGNITION switches. For normal operation, the ENG
IGNITION switch must be set to the AUTO position for the FADEC to have The STARTING includes this (these) subsystem(s):
control of the igniters. See the various starting modes in the table below:
• CRANKING (AMM SDS 80-10-00/1)
Normal Components
IDLE
Start
START AUTO/ON or N/A
(ground CRANKING (80-10)
Above
or in flight
Cranking is the system function utilized to perform the starting operation,
Ignition cir- basically consisting of starter-generator, SC (Start Contactor) and ENG
Wet Mo-
START AUTO/ON IDLE cuit break- START/STOP switch.
toring
ers open
Electrical Operation
Dry Mo-
START OFF IDLE Pumps -
toring
OFF

EFFECTIVITY: ALL
STARTING 80-00
NOTES: NOTES:

21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
STARTING 80-00
The starting cycle is initiated by setting the ENG START/STOP switch to

START on the ENG control panel, and is automatically interrupted as the
predetermined N2 value is reached.
During the starting, the starter-generator actuates as a starter motor that

drives the engine gas generator section.


It is very important, for the preservation of the start-generator service life, that
the specified operational limits not be exceeded.
The figure STARTING - RH ENGINE SCHEMATIC DIAGRAM provides


EFFECTIVITY: ALL
STARTING 80-00
CONTROL PEDESTAL GEA 1 GCU 1 GCU 2
ENG / START / STOP
RUN
STOP START
1

CENTRAL BUS
FADEC 1 ELEC RACK STARTER/
GENERATOR
CHANNEL A
START IN TO QSC
SHUTDOWN
SC1
AUX
START CMD
TO ELECTRICAL
SC1 GENERATION
LPDU
CHANNEL B
START IN
SHUTDOWN
START CMD
GCU 1
STARTING - LH ENGINE SCHEMATIC DIAGRAM
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
STARTING 80-00
NOTES: NOTES:

EFFECTIVITY: ALL
STARTING 80-00
CONTROL PEDESTAL GEA 2 GCU 1 GCU 2
ENG / START / STOP
RUN
STOP START
2

CENTRAL BUS
FADEC 2 ELEC RACK STARTER/
GENERATOR
CHANNEL A
START IN TO QSC
SHUTDOWN
SC1
AUX
START CMD
TO ELECTRICAL
SC1 GENERATION
RPDU
CHANNEL B
START IN
SHUTDOWN
START CMD
GCU 2
STARTING - RH ENGINE SCHEMATIC DIAGRAM
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80
EFFECTIVITY: ALL
CRANKING 80-10
Introduction percent to function as a starter motor. The starting cycle can also be
interrupted when the switch is set to the STOP position.
Cranking is the system function utilized to perform the starting operation,
basically consisting of starter-generator, SC (Start Contactor) and ENG Information Training Points
START/STOP switch. The starter-generator must be visually checked whenever the engine lower
General Description mid cowl is opened. It must be checked for general condition, signs of oil
leakage on its installation mount, security, and condition of the cooling
The engine starting is a semiautomatic process, referred to as starting cycle, system.

in which the starting and ignition systems (AMM SDS 74-00-00/1) are
actuated simultaneously. The ENG START/STOP switch has two positions The figure CRANKING - COMPONENT LOCATION provides further data on
as follows: the preceding text.
START
Starts the starting cycle that is interrupted when the N2 (Core Rotor Speed)
value is reachead.
STOP
This position allows the interruption of the starting cycle by cutting off the
electrical power supply to the starter-generator.
Components
STARTER/GENERATOR (AMM SDS 24-00-00/1)
Operation
When the starting is commanded, the starter-generator actuates the gas

generator section, accelerating it so that it reaches the sufficient rotation to
produce the beginning of the air/fuel mixture burning in the combustion
chamber.
When the ENG START/STOP switch is selected to START, the voltage is

supplied to the GCU (Generator Control Unit) which energizes the starting
relays. The starting cycle is initiated when the switch is set to the START
position and it is finished when the N2 indication reaches 44%, a preset

EFFECTIVITY: ALL
CRANKING 80-10

BOTTLE YAW
SHUTOFF 1 SHUTOFF 2
LEFT RIGHT
DISCH
ROLL
A OFF LWD RWD
ENG START/STOP
RUN RUN
PITCH BKP
DN

UP
1 2
ENG IGNITION MODE
+
ON BKP
AUTO
OFF OFF
1 2
FIRE/ENG/TRIM
CONTROL PANEL
A
ZONES
434
444
B GCU 1
GCU 2
RPDU
C
B
ZONES
241
C LPDU
242 STARTER/GENERATOR
CRANKING - COMPONENT LOCATION
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 38 44 45 50 51 52 53 54 55 56 57 71 72 73 74 75 76 77 78 79 80

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FRONT MATTER - page 6

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21 ...........................................................AIR CONDITIONING 55 .....................................................................STABILIZERS

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FRONT MATTER - page 8

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CHAPTER 26 - FIRE PROTECTION

SECTION TITLE PAGE

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22-Aug-2008 CHAPTER 26 - page 9

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Introduction • One portable fire extinguisher.

General Description • GIA (Garmin Integrated Avionics unit)

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FIRE/SMOKE DETECTION SYSTEM (26-10)

FIRE EXTINGUISHING (26-20)

The function of the fire extinguishing system is to discharge fire extinguishing

The fire protection system has these components:

• One fire detector per engine, in a single loop configuration.

• One ENG FIRE EXTINGUISHER and one TEST control panels.

• One engine fire extinguishing bottle.

22-Aug-2008 CHAPTER 26 - page 10

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Developed for Training Purposes Only

ENGINE FIRE EXTINGUISHING BOTTLE

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ENGINE FIRE EXTINGUISHING BOTTLE

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The FIRE/SMOKE DETECTION SYSTEM includes this subsystem:

Developed for Training Purposes Only

ENGINE FIRE/OVERHEAT DETECTION SYSTEM (26-11)

The figure FIRE/SMOKE DETECTION SYSTEM - BLOCK DIAGRAM

22-Aug-2008 CHAPTER 26 - page 14

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Developed for Training Purposes Only

The system is able to detect either overheat (average temperature) or fire

Each engine fire detector is electrically connected to supply:

– Warning indication by means of a red light on the engine shutoff

– Warning indication by means of a red FIRE message in the ITT (Interstage

– Voice message FIRE.

– Engine 1(2) - Hot Bus 1.

22-Aug-2008 CHAPTER 26 - page 18

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99.9 N2% STALL PROT

OIL PRES PSI