|
OPERATING MANUAL
PSP 601A-6
SECTION 17
POWER PLANT
TABLE OF CONTENTS
Page
1. GENERAL 1
2. ENGINE FUEL SYSTEM 1
A. Firewall Fuel Shutoff Valve 2
B. Fuel Pressure Sensor 2
C. Engine Fuel Pump 2
D. Fuel Heater (Aircraft 5001 to 5134) 2
E. Heat Exchanger (Aircraft 5135 and subsequent) 2
F. Fuel Temperature Sensor 2
G. Fuel Filter 2
H. Fuel Control Unit (FCU) 3
I . Fuel Flow Transmitter 3
J. Oil Cooler 3
K. Fuel-Row Distributor and Injectors (Aircraft 5001 to 5134) 3
L. Fuel Manifold (Aircraft 5135 and subsequent) 3
M. Fuel Injectors (Aircraft 5135 and subsequent) 3
N. Ecological Drain System (Aircraft 5001 to 5134) 3
3. AUTOMATIC PERFORMANCE RESERVE (APR) SYSTEM 4
4. ENGINE OIL SYSTEM 4
A. Oil Replenishment System 4
B. Oil Storage Tank 5
C. Oil Temperature Sensor 5
D. Oil Circulation 5
E. Oil Filter 5
F. Oil Cooler 5
G. Oil Pressure Sensor and Low Oil Pressure Switch 5
5 . ENGINE CONTROLS 6
6. THRUST REVERSER 6
6
7
7
7
7
8
8
17 - CONTENTS
Page 1
Apr 10/95
A.
B.
Operation
Safety Features
(1) Throttle Retarder System
(2) Throttle Lockout System
(3) Auto Slow System
(4) Emergency Stow System
(5) Safety relay
OPERATING MANUAL
PSP 601A-6
Page
7. ENGINE INSTRUMENTS 8
A. Signal Data Converter (SDC) 8
B. Engine Instruments 8
8. EN6INE BLEED AIR 8
A. Tenth Stage Bleed Air 9
B. Fourteenth Stage Bleed Air 9
C. Bleed Air Leak Detection and Warning System 9
9. ENGINE STARTING AND IGNITION SYSTEMS 10
A. Ground Starting 10
B. In-Right Starts 10
C. Continuous Ignition 11
10. ENGINE VIBRATION MONITORING SYSTEM 11
LIST OF ILLUSTRATIONS
Figure Title Page
Number
1 Power Plant - Schematic 12
2 Engine Fuel System - Schematic (2 Sheets) 13
3 Fuel Control Panel - Engine Fuel System Monitoring 14A
4 APR and Engine Speed Control Panel 15
5 Engine Oil System - Schematic 16
6 Oil Temperature and Pressure Indicators 17
7 Throttle Quadrant and Thrust Reverser Controls and Indicators 18
8 Thrust Reverser Stowed and Deployed Positions 19
9 Engine Instruments and Control Panel 20
10 Tenth Stage Engine Bleed Air - Schematic 21
11 Bleed Air Control Panel 22
12 Fourteenth Stage Engine Bleed Air - Schematic 23
17 - CONTENTS
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OPERATING MANUAL
PSP 601A-6
Figure
Number Title Page
13 Bleed Air Leak Warning and Testing 24
14 Engine Start and Ignition Controls 25
15 Engine Vibration Monitor Panel 26
17-CONTENTS
Page 3
Apr 02/87
OPERATING MANUAL
PSP 601A-6
SECTION 17
POWER PLANT
. GENERAL (Figure 1)
The aircraft is powered by two General Electric CF34 turbofan engines. The
engine is a dual-rotor, front-fan configuration with a bypass ratio of:
6.2:1 (aircraft 5001 to 5134)
6.26:1 (aircraft 5135 and subsequent).
The low pressure (or NJ rotor consists of a single-stage fan driven by a
four-stage low-pressure turbine. A high-pressure (or N2) rotor consists of a
fourteen-stage axial-flow compressor driven by a two-stage turbine. For a
two-engine operation under standard sea-level conditions, the engine is rated
at a take-off thrust of 8,729 pounds. An automatic performance reserve (APR)
system increases the standard take-off thrust rating to 9,220 pounds if an
engine failure occurs.
The engine airflow passes through the fan assembly and is divided into two
airflow systems. The main airflow, bypass air, is routed around the core
cowls and exhausts through the thrust reverser assembly, over the tailpipe
fairing. The remaining airflow passes through the engine core consisting of
the compressor, a combustion chamber, the high-pressure turbine and the
low-pressure turbine. The hot gas is then exhausted through an exhaust
nozzle.
The compressor has a variable geometry system that varies the position of the
compressor inlet guide vanes and the first five stages of the stator vanes.
The system operates throughout the operating range of the engine to improve
compressor efficiency and prevent stalling and surging.
An accessory gearbox, driven by the N2 rotor, drives the engine lubrication
pumps and fuel pump as well as an aircraft hydraulic pump and ac generator.
The engine starter drives the N2 rotor through this accessory gear box.
Bleed air is taken from the 10th and 14th stages of the compressor for air
conditioning/pressurization, engine crossbleed starting, anti-icing and
thrust reverser operation.
. ENGINE FUEL SYSTEM (Figures 2 and 3)
Each engine has a self-contained fuel system for the controlled distribution
of fuel to the combustion chamber. Secondary functions of the system are
control of the compressor variable geometry system, cooling of engine oil,
and motive fuel supply to the aircraft fuel system ejector pumps (refer to
Section 12).
Sensors are installed at suitable locations in the system to provide the
required inputs to the flight compartment controls and indicators.
On aircraft 5001 to 5134,
the principal components that make up the engine fuel system are described in
a fuel flow sequence, starting at the firewall fuel shutoff valve through to
the combustion chamber and the ecological drain system.
On aircraft 5135 and subsequent,
the principal components that make up the engine fuel system are described in
a fuel flow sequence* starting at the firewall fuel shutoff valve through to
the combustion chamber.
SECTION 17
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OPERATING MANUAL
PSP 601A-6
A. Firewall Fuel Shutoff Valve
This valve isolates the engine fuel system from the engine feed line
(refer to Sections 9 and 12).
B. Fuel Pressure Sensor
This sensor, connected to a warning light in the flight compartment,
allows monitoring of fuel pressure from associated ejector and electric
fuel pumps in the aircraft fuel system,
C. Engine Fuel Pump
The engine-driven fuel pump contains a low-pressure section and a
high-pressure section. The low-pressure section supplies fuel through
the fuel heater to the fuel filter, and back to the high-pressure section
of the pump.
The high-pressure section is divided into two elements, designated the
primary element and the secondary element. The primary element pumps
fuel from the fuel filter to the fuel control unit and the secondary
I element supplies high-pressure fuel to the aircraft tanks for motive
I flow.
I D. Fuel Heater (Aircraft 5001 to 5134)
The fuel heater is an air-to-liquid heat exchanger which uses hot
compressor bleed air to heat the fuel. The fuel temperature is
maintained above 5*C, by a thermal sensor and an air modulating valve, to
prevent icing in the fuel filter. A fuel bypass valve allows fuel to
bypass the fuel heater should i t become clogged.
| E. Heat Exchanger (Aircraft 5135 and subsequent)
I Fuel is heated by the liquid to liquid (oil to fuel) heat exchanger.
I F. Fuel Temperature Sensor
This sensor, connected to an indicator in the flight compartment, allows
monitoring of the fuel temperature and the fuel heater operation.
| 6. Fuel Filter
The fuel filter, located downstream of the fuel heater, contains a
disposable filter element. A bypass valve allows the fuel to bypass the
filter element should it become clogged. A differential pressure switch
connected to a warning light in the flight compartment warns of an
impending bypass condition. Should a bypass occur, a red button on the
housing rises to indicate the condition.
SECTION 17
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Apr 10/95
OPERATING MANUAL
PSP 601A-6
J H. Fuel Control Unit (FCU)
The FCU i s a hydro-mechanical/electrical device consisting of two
sections: a fuel metering section and a computer section.
The flight compartment throttle lever movement is transmitted to the FCU
which in turn controls engine speed in one of the following two modes:
- At relatively low power settings, the FCU hydro-mechanically meters the
fuel to the injectors to control engine N2 speed. In this mode,
matched movement of the throttle levers produces matched N2 speeds, but
Ni speeds and thrust may be mismatched between the engines.
- At take-off, climb and cruise power settings, the engine is Ni speed
controlled. In this mode, the FCU electrically responds to Ni speed
references. (Electrical power for operation in this mode is supplied
by an Ni driven alternator, completely independent of the aircraft
electrical system). Matched movement of the throttle levers produces
matched Ni speeds, hence matched thrust between the engines. This
engine speed control can be selected on or off by a switch in the
flight compartment. If selected off, the engine speed control reverts
to the N2 mode described above for all engine speeds.
The FCU also meters pressurized fuel from the engine-driven fuel pump to
the two actuators for the compressor variable geometry system. The
actuators move the compressor inlet guide vanes, and the affected stator
vanes open as engine speed increases and close as speed decreases.
j I. Fuel Flow Transmitter
This transmitter sends fuel flow information to be displayed on the
associated fuel flow indicator in the flight compartment.
I J. Oil Cooler
The oil cooler heats the fuel before the fuel enters the combustion
chamber while cooling the engine oil. (Refer to paragraph 4.F.)
I K. Fuel-Flow Distributor and Injectors (Aircraft 5001 to 5134)
The fuel-flow distributor precisely meters the fuel to the injectors.
The injectors inject fuel into the combustion chamber. At shutdown, the
distributor drains fuel to an ecological drain system.
J L. Fuel Manifold (Aircraft 5135 and subsequent)
Metered fuel leaves the FCU, thru fuel flow transmitter and enters fuel
I manifold. The manifold consists of 2 separate halves, which form one
J continuous ring which encircles the combustion chamber frames.
M. Fuel Injectors (Aircraft 5135 and subsequent)
Integral with the continuous ring are 18 fuel injector hoses which
connect to 18 fuel injectors. Each injector has 2 independent fuel flow
passages, a primary and a secondary. The primary introduces fuel into
I the combuster at lower power settings (start-up-idle) the secondary
introduces fuel at higher power settings, resulting in 2 cones of fuel.
I N. Ecological Drain System (Aircraft 5001 to 5134)
This system prevents the fuel collected by the shut-down fuel drain
system from being discharged to the atmosphere. The ecological drain
tank collects this fuel which is then routed to an aspirator in the
engine fuel feed line to be consumed during the subsequent engine
operation.
SECTION 17
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OPERATING MANUAL
PSP 601A-6
3. AUTOMATIC PERFORMANCE RESERVE (APR) SYSTEM (Figures 2 and 4)
The APR system monitors engine thrust levels at high power settings and
automatically commands an increase in thrust on both engines if a
predetermined thrust loss is detected on one of them. To arm the system,
both engine speed control switches must be selected to ON and the APR switch
selected to ARM.
With the system armed and the engines operating in the Ni speed control mode,
an Ni drop to 5000 rpm (approximately 67.5% Ni) on either engine causes the
APR controller to command an N2 speed increase of 167 rpm (approximately 2%
Ni) on both engines. The engine still operating at the normal take-off Nx
has its Ni increased by approximately 2% while the engine affected by the Ni
drop reverts to N2 speed control mode, hence not responding to the Ni speed
increase command.
NOTE: The APR system does not affect or override the throttle lever inputs
to the FCU. Therefore, it is possible to advance the throttles and
obtain power settings higher than the normal (non-APR) take-off
thrust. Should this condition be followed by a power loss on one of
the engines, the other engine would respond to the APR command and
further increase Ni above its previously set higher power setting,
with the likely result of its inter-turbine temperature (ITT) limits
being exceeded.
Two system tests, a static and a dynamic system test, ensure system
serviceability before take-off- The static test is performed with the
engines running at idle, using a selector switch in the flight compartment.
The dynamic test is conducted automatically by the APR controller, with the
system armed and when the engines are accelerated through 83.5% Ni for
take-off. The dynamic system test cannot be repeated unless the
weight-on-wheels status changes or the system is selected off and re-armed.
4. ENGINE OIL SYSTEM (Figures 5 and 6)
Each engine is lubricated and cooled by its own self-contained oil system.
In addition to an engine-mounted oil storage tank, an oil replenishment
system located in the rear equipment bay is also provided.
Sensors are installed at suitable locations in the system to provide inputs
to the oil temperature and pressure indicators in the flight compartment.
Impending oil filter blockage is also monitored.
The principal components that make up the engine oil system are described in
a flow sequence, starting at the oil replenishment system through to the oil
return to the oil storage tank.
A. Oil Replenishment System
This system is used to add oil to the oil storage tanks on both engines.
It consists of a replenishment tank, an electric pump, an OIL LEVEL
CONTROL panel and a selector valve.
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Apr 10/95
OPERATING MANUAL
PS? 601A-6
The OIL LEVEL CONTROL panel indicates if the storage tanks are full and
also tests the replenishment system. The selector valve is manually
selected to the desired left or right storage tank. The electric pump
transfers oil from the replenishment tank to the selected storage tank.
The replenishment tank is gravity-filled. Its tank-mounted sight gauge
indicates oil level.
B. Oil Storage Tank
Each storage tank oil level can be determined by a dips tic mounted in the
filler cap. The tank can be directly gravity filled or remotely filled
using the replenishment system.
C. Oil Temperature Sensor
This sensor, located in the storage tank and connected to an indicator in
the f l i g h t compartment, allows monitoring of engine oil temperature
including the oil cooler operation.
D. Oil Circulation
Oil flows from the storage tank to the lube pump. The pressurized oil is
then directed through the f i l t e r , the cooler and then to the various engine
components requiring lubrication and cooling. Oil is returned to the
storage tank by scavenge pumps.
E. Oil Filter
The o i l f i l t e r consists of a f i l t e r element, a differential pressure switch
connected to an oil pressure impending bypass indicator, and a bypass
valve. The impending bypass indicator, located in the rear equipment bay,
warns of an impending blockage of the f i l t e r element. Should the f i l t er
become clogged, the bypass valve would open to allow unfiltered oil to
maintain engine lubrication.
F. Oil Cooler
The o i l cooler is an oil-to-fuel heat exchanger which uses fuel as a
cooling medium for the engine o i l .
G* Oil Pressure Sensor and Low Oil Pressure Switch
The pressure sensor and the low pressure switch, connected to their
indicator and warning light respectively in the flight compartment, provide
independent indications of engine oil pressure. Both circuits sense the
differential oil pressure between the lube pump discharge and the scavenge
pump suction.
SECTION 17
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Apr 02/87
OPERATING MANUAL
PS? 601A-6
ENGINE CONTROLS (Figure 7)
Each throttle lever with its hinged thrust reverse (TR) lever is connected to
i t s engine fuel control unit (FCU) through a single flexible cable extending
from the throttle quadrant in the flight compartment to the throttle control
gearbox on the engine.
This system transfers all throttle and thrust reverse lever movement to the
engine to command forward or reverse thrust as well as fuel shutoff. This
system also mechanically provides a t a c t i l e feedback to the throttle and thrust
reverse levers when the FCU is kept at, or should be returned to, idle by the
t h r o t t l e retarder system.
Individual throttle lever release latches prevent inadvertent selection of fuel
SHUT OFF or fuel-on (IDLE) and individual thrust reverse lever release latches
prevent inadvertent operation of thrust reverse levers. Mechanical interlocks
within the throttle quadrant also prevent a thrust reverse lever from being
operated unless its throttle lever is at IDLE, or prevent a throttle lever from
being advanced above IDLE when i ts thrust reverse lever is pulled up from the
forward idle position.
A t h r o t t l e lever friction adjustment control is also provided.
THRUST REVERSER (Figures 7 and 8)
Each engine is equipped with a thrust reverser to assist in aircraft braking
after landing. When the thrust reverser is deployed, a translating cowl moves
rearward on tracks driven by a pneumatic actuator and uncovers forward facing
cascade vanes. Blocker doors, operated by interconnected linkages, move inward
to block the fan air exhaust duct and redirect the fan exhaust air through the
cascade vanes. The system is powered by 14th stage bleed air.
A. Operation
Each thrust reverser is armed for operation by i ts associated REVERSE
THRUST switch/light. With the throttle lever at IDLE and either a
weight-on-wheels signal or a 16-knot wheel spin-up signal, raising the
thrust reverse lever to the deploy position initiates the following
deployment sequence:
- Wing and engine anti-icing shutoff valves, i f open, are closed to
conserve 14th stage bleed air for the reverser operation.
- 14th stage bleed air unlocks the reverser from i t s stowed position and
powers a pneumatic drive unit (PDU) which deploys the reverser.
• As the reverser is fully deployed, i t is locked in position by the PDU,
and the thrust reverse lever is mechanically released by a throttle
retarder system and is electrically released by a throttle solenoid, so
that thrust settings up to maximum reverse thrust can be selected.
SECTION 17
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OPERATING MANUAL
PSP 601A-6
Moving the thrust reverse lever back to the deploy position t h r o t t l e s the
engine down to reverse i d l e - Continued movement of the lever to the f u l ly
down position stows and locks the reverser, in the reverse sequence to the
deployment described above.
As the reverser is f u l l y stowed and locked in position by the PDU, the
t h r o t t l e retarder system mechanically releases the t h r o t t l e lever so that
forward thrust setting above IDLE can be selected again. Wing and engine
a n t i - i c i ng is also restored to normal controls.
After the above sequence to stow and lock the reverser is complete, the
reverser can be disarmed by i t s associated REVERSE THRUST switch/light.
B. Safety Features
Each reverser is protected by the safety features described below.
(1) Throttle Retarder System
The t h r o t t le retarder system is a mechanical system connected to the
thrust reverser and the engine FCU control linkage. If the engine is
t h r o t t l ed above i d l e and there is an inadvertent thrust reverser
deployment, the t h r o t t l e retarder system returns the FCU to i d l e , and,
through the i n t e r l i n k cable, pulls the t h r o t t l e lever to IDLE. This
system also operates during normal operation of the reverser, as
described in paragraph 6.A. above.
(2) Throttle Lockout System
The throttle lockout system prevents movement above IDLE of a
previously retarded throttle lever if the aircraft is airborne and the
thrust reverser moves from the fully stowed position. Under such
conditions, initial movement of the thrust reverser energizes a
throttle lockout solenoid which locks the throttle linkage at the FCU
and prevents throttle lever movement beyond IDLE. If the reverser is
returned to the stowed position, the throttle lockout solenoid is
de-energized and freedom of movement is returned to the throttle lever.
(3) Auto Stow System
In the case of an uncommanded movement of the reverser from the stowed
position, a microswitch commands the PDU to return the reverser to the
stowed position.
SECTION 17
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Apr 02/87
OPERATING MANUAL
PSP 601A-6
(4) Emergency Stow System
If the REVERSER UNLOCKED light is on and the auto stow system fails to
stow the reverser, the appropriate THRUST REVERSER EMERG STOW
switch/light can be pressed to ensure positive operation of the sys^--
to the stowed position.
(5) Safety Relay
Actuation of the thrust reverse lever with the aircraft not on the
ground causes the REVERSE THRUST UNSAFE TO ARM light to come on.
ENGINE INSTRUMENTS (Figure 9)
Engine instruments monitor Nl %rpm, inter-turbine temperature (ITT), N2 %rpm,
fuel flow, oil temperature and oil pressure.
A. Signal Data Converter (SDC)
The SDC controls the power supply and provides automatic dimming to the
engine indicator systems.. Two power supplies are divided within the SDC
into dual lamp-processing and signal-processing power supplies.
B. Engine Instruments
Each instrument provides a vertical analog display of the relevant engine
variable using a series of miniature incandescent lamps inside the
indicator, which provide the light source for the vertical row of coloured
light segments.
Digital displays on the Nl, ITT, N2 and FUEL FLOW indicators provide more
accurate indications when compared with the readings on the vertical scales.
ENGINE BLEED AIR
Engine bleed air consists of two systems, each with its own source. One source
is at the 10th stage and the other is at the 14th stage of the compressor of
each engine. Each system contains distribution ducting, shutoff valves,
isolator valves and check valves. Both systems are protected by a bleed air
leak detection system. The flight compartment BLEED AIR and ANTI-ICE control
panels provide the necessary controls and indicators.
SECTION 17
Page 8
Apr 02/87
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OPERATING KANUAL
PSP 601A-6
Tenth Stage Bleed Air (Figures 10 and 11)
The 10th stage bleed a i r system can be supplied from the l e f t and right
engines or from the APU, or from a ground a i r supply unit through an
external connection on the lower left side of the rear fuselage. The 10th
stage system supplies bleed air to the following systems:
Air conditioning/pressurization
Cabin pressurization control
Footwarmer/demister and emergency pressurization
Engine s t a r t i ng
A bleed air isolator valve is normally closed to separate the l e f t and
right d i s t r i b u t i o n ducting. This isolator valve is automatically opened by
the engine s t a r t system to ensure air supply to both engines regardless of
the a i r source. It can also be selected open when required, i . e . , to
supply both ACUs from a single engine bleed source.
A l e f t and r i g h t pressure indicator receives signals from two sensors, ont
on each side of the isolator valve, for continuous monitoring. For
operation of LH and RH footwarmer/demister valves and l e f t and r i g h t ACU
valves, refer to Section 2.
Fourteenth Stage Bleed Air (Figures 11 and 12)
The 14th stage bleed a i r system is supplied only by the l e f t and right
engines. The 14th stage system supplies bleed air to the following systems:
Wing a n t i - i c i ng
Engine a n t i - i c i ng
Thrust reverser
The operation of engine and wing a n t i - i c i ng valves (including the isolator
valve) is controlled by the ANTI-ICE control panel (refer to Section 14).
Bleed Air Leak Detection and Warning System (Figure 13)
Six temperature sensors ( f i r e - w i r e type) are attached to the bleed a ir
ducts and are connected to two bleed air leak detection control units.
Dual detection loops are provided for the l e f t and r i g h t sections of the
10th stage bleed air system, and single loops are provided for the 14th
stage bleed a i r system, the fuselage pylons and the a n t i - i c i n g ducts
running through the fuselage and wings. If a leak occurs, the hot bleed
a i r escaping i s detected by the temperature sensors and i n i t i a t e s a warning
signal.
SECTION 17
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Apr 02/87
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cftanenoer
OPERATING MANUAL
PSP 601A-6
The warning signal is picked up by its leak detection control unit and
transmitted to the flight compartment via a centrally located flashing DUCT
FAIL warning light as well as an appropriate individual DUCT FAIL warning
light on the control panel for the system affected by the leak.
The individual DUCT FAIL warning light identifies the defective duct whicn
can then be depressurized and isolated using the BLEED AIR control panel.
A bleed air leak annunciator panel, behind the copilot's seat, also
provides for fault isolation through eight latching magnetic indicators.
With the exception of the indicators on the bleed air leak annunciator
panel, all of the warning indicators go out when their associated
temperature sensors have cooled sufficiently. Warnings and testing of the
bleed air leak detection system are summarized in Figure 13.
9. ENGINE STARTING AND IGNITION SYSTEMS (Figure 14)
The engine starting and ignition systems consist of a pneumatically driven air
turbine starter, ignition-exciter boxes and igniter plugs for each engine. The
systems are controlled by individual switch/lights in the flight compartment.
The starter transmits starting torque to the N2 rotor through the accessory
gearbox. An automatic centrifugal shutoff switch opens at a preset rpm to
protect the starter against overspeed. The 10th stage bleed air system
supplies the starter through a starter valve.
Each engine has two ignition systems, A and B, each system connected to its
igniter plug in the combustion chamber. The systems are powered by 115-volt ac
power.
A. Ground Starting
The APU, an external ground air source or an operating engine can be used
to supply the 10th stage bleed air system for engine starting.
An IGN A and/or IGN B switch/light arm(s) the associated igniter plugs on
both engines. With the 10th stage bleed air system pressurized, pressing a
START switch/light initiates the starting sequence on the associated
engine. Refer to Figure 14 for description of the starting sequence.
B. In-Flight Starts
An IN FLIGHT START switch/light provides a separate power supply and fires
both igniter plugs (A and B) on the selected engine without any other
switch/light having to be operated- However, if the windmilling rpm is
less than 131 N2, starter assist using the START switch/light is required.
SECTION 17
Page 10
Apr 02/87
OPERATING MANUAL
PSP 601A-6
C. Continuous Ignition
A CONT IGN switch/light provides continuous i g n i t i on to both engines
through the pre-selected A and/or B i g n i t e r plugs. Firing of the igniter
plugs is continuous until the CONT IGN switch/light is pressed out.
10. ENGINE VIBRATION MONITORING SYSTEM (Figure 15)
The engine vibration monitoring (EVM) system provides a continuous indication
of the vibration level of each engine. The main components of the system
include a transducer mounted on the compressor casing of each engine, a signal
conditioner and an indicator panel in the f l i g h t compartment.
Each transducer generates an electrical signal proportional to the intensity of
engine vibration. The signal conditioner converts these signals into values
readable on the EVM indicator.
An alarm c i r c u i t causes an amber caution light on the EVM indicator panel to
come on i f the vibration level of either engine exceeds 1.7 MILS for a period
greater than 3 seconds. This 3-second delay prevents spurious warnings caused
by high transient engine vibrations.
SECTION 17
Page 11
Apr 02/87
OPERATING MANUAL
PS? 601A-6
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Figure 1
SECTION 17
Page 12
Apr 02/87
OPERATING MANUAL
PSP 601A-6
BLEED
AIR f
4-
ECOLOGICAL
DRAIN
TANK
FIREWALL!
SHUTOFF
(VALVE
AIRCRAFT
FUEL
SYSTEM W
11 Hftl I j
F™5H
TO RIGHT
ENGINE ^F WBT
e=FECTIVnY: A C 6001TO 613*
LEFT ENGINE ILLUSTRATED
LEGEND
• I • I • MOTIVE FUEL SUPPLY
i l l l l l l l f l t l METERED PRESSURIZED FUEL
* * * * * * ECOLOGICAL DRAIN LINE
ftT] DIFFERENTIAL PRESSURE SWITCH
Engine Fuel System - Schematic
Figure 2 (Sheet 1)
SECTION 17
Page 13
Apr 10/95
OPERATING MANUAL
PSP 601A-6
1 j MAI«=OU)~H
UlLUHVUV; A/C619S AND SUBS
LEFT ENGINE ILLUSTRATED
LEGEND
• ( • I B MOTIVE FUEL SUPPLY
millttllir METBIED PRESSURIZED FUEL
fp"] OfFFERBCTIAL PRESSURE SWITCH
Engine Fuel System - Schematic
Figure 2 (Sheet 2)
SECTION 17
Page 14
Apr 10/95
OPERATING MANUAL
PSP 601A-6
FUELTEMPERATURE INDICATOR
Shows temperature at left and right toef heater outlets.
A!C 6001 TO 5134:
Nofnw opefflDHQ rai
Cwbcmry range ftraDoer}
and
>VC 5136 AND 8UB8:
r e to acre
-2CrCto5*C
6CTCto70*C
No*nrfop*rata^ range (ferem) 4*Cto120*C
Ceitfomry range fj^iiow) -66~Cto4*C
FUEL CONTROL
PUSH ON/OFF 1
RTANK
PUMP EJCTRS
d
LOW PRESSURE WARNING UGHTS
Amber warning light comes on to indicate low pressure at
associated engine fuel inlet port.
VALVE CLOSED UGHTS
White light comes on whenever associated firewall fuel ahutoff
valve is closed.
FILTER BYPASS WARNING UGHTS
Amber light comes on when fuel pressure drop is detected
across associated main fuel fitter.
NOTES
1 Refer to FUEL for details of aircraft fuel
system control and rnonrtoring.
2 On A/C 5135 and SUBS, the fuel
LOW PRESS lights wiH be on
until the pumps are selected ON.
CENTRE INSTRUMENT PANEL
Fuel Control Panel -
Engine Fuel System Monitoring
Figure 3
SECTION 17
Page 14A
Apr 10/95
OPBUTMG MANUAL
PSP 601A-6
THIS PAGE INTENTIONALLY LEFT BUNK
SECTION 17
Page 14B
Apr 10/95
ctiat/encjer
OPERATING MANUAL
PSP 601A-6
READY LIGHT
Green light comes on to confirm system
readiness after system is initially armed.
When APR operates, light goes out when
L ON or R ON light comes on.
L ON/R ON LIGHTS
Green lights come on to
indicate left or right engine is
responding to APR command
following a power loss.
APR SELECTOR SWITCH
Three-position toggle switch:
ARM - Arms system if both ENG SPEED
CONTROLS switches are on, both
engines are in Nl speed control mode
and APR light is out.
OFF - De-activates system.
TEST/RESET - Initiates static test of
system (refer to NORMAL
PROCEDURES). Resets system after a
fault is cleared.
APR LIGHTS
Amber light comes on when:
- APR selector switch is off prior to take-off.
- APR self-monitoring circuits detect fault in APR
or engine fuel control system.
READY
li TEST I
] I
APR
ENG. SPEED
CONTROL
ON ON
Green light comes on during
static and dynamic tests of the
system (refer to NORMAL
PROCEDURES).
ENGINE SPEED CONTROL SWITCHES
Two-position toggle switches.
ON - Engine speed control i s m N l mode
when NT rpm exceeds nominal 79.1 %.
OFF - Engine speed control is in N2
mode regardless of Nl rpm.
CENTRE INSTRUMENT PANEL
APR and Engine Speed Control Panel SECTION 17
Figure 4 Page 15
Apr 02/87
OPERATING MANUAL
PSP 601A-6
t
LD
BYPASS VALVE
DIFFERENTIAL PRESSURE
SENSOR OR SWITCH
Engine Oil System - Schematic
Figure 5
SECTION 17
Page 16
Apr 02/87
OPERATING MANUAL
PSP 601A-6
OIL PRESSURE INDICATOR
Vertical scale indicator displays ofl pressure of each engine.
Coloured fight segments of vertical scales come on
to indicate the following range.
NZ 5001 TO 5134:
Low pressure warning Ene (red) 25 psi
Normal operating range (green) 25 to 95 psi
Cautionary pressure range (yellow) 95 to 100 psi
High pressure warning line (red) 100 psi
OIL
TEMP
—»1t0 —
1—140 —
|«» «» IOMI mmm
\mmm ^m
— 120 —
«^» «a»
! « • • SMS
«•» mm
— 100«-»
!•» ^» »••• GO sea
• • • OBi
aa» 0 aaa
aaaB -70 -IB*
0
jd j
j 1 q 1
IJ 1
•i I
J 1
"1 1
R
A/C 5135 AND SUBS:
Low pressure warning fine (red) 25 psi
Normal operating range (green) 25 to 115 psi
Cautionary pressure range (yellow) 115 to 130 psi
High pressure warning line (red) 130 psi
OIL TEMPERATURE aMDICATOft
Vertical scale M o t o r dspisyt oil
Coloured fight segments of vertical come on to indicate the
LOW OIL PRESSURE LIGHTS
Red warning lights coma on whan oil pressure of
associated angina drops below 2B i . 3 pax.
Normal operaUiy lenye tarean)
^M6onajYranga (yaaow)
Warning fine (red)
-20°Cto140°C
140°Cto180°C
ieo°c
ENGINE INSTRUMENT PANEL
Oil Temperature and Pressure Indicators SECTION 17
Figure 6 pa g e 17
Apr 10/95
OPERATING MANUAL
PS? 601A-6
THRUST REVERSER EMERG STOW
SWITCH/LIGHTS
When pressed, power is applied directly to
arming and stow solenoid valves to initiate
stowage of reverser.
Amber REVERSE UNLOCKED light comes on
whenever reverser moves from fully stowed
position and remains on until reverser is returned
to fully stowed position.
Green REVERSE THRUST light comes on when
reverser reaches fully deployed position and goes
out immediately when reverser moves from
deployed position.
THRUST REVERSER
EMERG STOW
REVERSER
UNLOCKED!
REVERSE
THRUST
REVERSER
UNLOCKED
REVERSE
THRUST
THRUST REVSISE (TR) LEVERS
With throttle levels at IDLE, putting
onTR levels deploy levsisers if
following conditions met:
- REVERSE THRUST switch/lights
have armed reversers.
• Aircraft on ground or wheel spinup
exceeds 16 knots.
Throttle solenoids prevent TR lever
movernem beyond deploy (or reverse
idle) position until reverser
assemblies fully deployed.
Once reversers fully deployed, TR
levers regulate reverse thrust from
reverse idle to maximum reverse
power.
Reverser operation shuts off 14th
stage bleed air to engine and wing
GO-AROUND SWITCHES
Momentarily push button switches.
These switches are associated with
go-around mode of flight director
system.
Returning TR levers to forward IDLE
(fully down) stow reversers. Once
reversers stowed, throttle levers can
be moved forward to increase thrust.
NOTE Reverser deployment do not
prevent throttle levers from being
selected to SHUTOFF.
THROTTLE SETTINGS
SHUTOFF - Shuts off fuel to engine
at the FCU. Located at rear throttle
IDLE - Lowest forward thrust
setting. Located at idle throttle lever
stop.
MAX POWER - Highest forward
thrust setting. Located at forward
throttle lever stop.
PUSH LEFT PUSH RIGHT
GLARESHIELD
THROTTLE LEVERS
Control forward thrust and acts a
fuel shutoffs. Remain locked at IDLE
position during thrust reverser
operation.
THROTTLE LEVER RELEASE LATCHES
Lift to advance throttle levers from SHUTOFF to
IDLE positions or retard throttle levers from IDLE
to SHUTOFF positions.
THRUST REVERSE LEVER
RELEASE LATCHES
Lift to release TR levers from
forward IDLE stops. ^ 3
THROTTLE LEVER FRICTION ADJUSTMENT
Adjusts friction on throttle levers only. Rotate
control clockwise to increase friction.
CENTRE PEDESTAL
REVERSE THRUST
SWITCH/UGHTS
When pressed in, arms thrust
reverser system and puts on amber
ARMED light.
When pressed out, providing thrust
reverser stowed* disarms thrust
reverser system and puts out amber
ARMED tight.
Amber UNSAFE TO ARM light
comes on if:
- electrical fault exists in reverser
F tEVERSE THRUST 3\
LEFT RIGHT ( j
| UNSAFE 1
j TO ARM j
ARMED
1 PUSH TO
1 UNSAFE 1
j TO ARM j
j ARMED
ARM
deploy is selected or deploy
switch fault occurs during flight.
CENTRE PEDESTAL
Throttle Quadrant and Thrust Reverser Controls
and Indicators
Figure 7
SECTION 17
Page 18
Apr 02/87
canatiair
ctiauenejer
OPERATING MANUAL
PSP 601A-6
STOWED POSITION
CASCADE
VANES
FAN AIR
BLOCKER DOORS
TRANSLATING COWL
TORQUE BOX
DEPLOYED POSITION
Thrust Reverser Stowed and Deployed Positions SECTION 17
Figure 8 Page 19
Apr 02/87
OPERATING MANUAL
PSP 601A-6
Engine Instruments and Control Panel
Figure 9
SECTION 17
Page 20
Apr 10/95
OPERATING MANUAL
PSP 601A-6
GROUND AIR
SUPPLY
A TO CABIN PRESSURIZATiON
? CONTROL SYSTEM
| JET PUMP
LEFT ENGINE
10TH STAGE
BLEEDS
RIGHT ENGINE
10TH STAGE
BLEEDS
ENGINE
START
SYSTai
CLOSE
EFFECTIVITY
H A/C 5001 TO 6134
LEGEND
BLEED AIR
ELECTRICAL SIGNAL
SHUTOFF VALVE
REGULAT1NG/SHUTOFF VALVE
CHECK VALVE
Tenth Stage Engine Bleed Air - Schematic
Figure 10
SECTION 17
Page 21
Apr 10/95
OPERATING MANUAL
PSP 601A-6
10TH AND 14TH STAGE SWITCH/UGHTS
I in, associated blood air shtitoff valve opens and
vriata BLEED CLOSED tight goes out Whan pressed out valve
tand light comas on.
BLEED AIR ISOL SWITCH/UGHT
Whan prassad in, bleed air isolator vafrve
opens. When prassad out valve closes.
Green OPEN light comes on whenever
bleed air aoiator valve is open.
BLEED AIR PRESSURE GAUGE
Indicates pressure in left and right
sections of 10TH stage bleed »r system.
EFFECTMTY: A/C 5135* SUBS
OVERHEAD PANEL
Bleed Air Control Panel SECTION 17
Figure 11 Page 22
Apr 10/95
OPERATING MANUAL
PSP 601A-6
TOWING
ANTMCING
TO ENGINE
ANTI-ICING
TO THRUST
REVERSER
LEFM4TH
STAGE
BJGINE
BLSD
TOWING
ANTMCING
TO ENGINE
ANTI-ICING
TO THRUST
REVERSER
CLOSE
-**
[ CLOSE THRUST
REVERSERS
IN
OPERATION
CLOSE
RIGHT 14TH
STAGE
ENGINE
BLEED
O
CLOSE U-+ RH ENG
RRE
PUSH
LEGEND
BLEED AIR
ELECTRICAL SIGNAL
SHUTOFF VALVE
REGULATING/SHUTOFF VALVE
CHECK VALVE
Fourteenth Stage Engine Bleed Air - Schematic
Figure 12
SECTION 17
Page 23
Apr 10/95
OPERATING MANUAL
PSP 601A-6
DUCT MON SWTTCH
Three-portion DUCT MON toggJe switch tests serviceability of each of the
detector loops A and B on the left and right 10th stage manifold sections.
LOOP A - Duct fail warning occurs if loop A of either section is damaged.
LOOP B - Duct fail warning occurs if loop B of either section is damaged.
BOTH - In-flight switch position. Both detection loops are in operation on left
and right sections.
DUCT FAIL LIGHTS (4)
Red light comes on if the bleed leak
temperature sensors detect a failure in the associated
duct segment.
Light goes out when the failed duct is isolated and
temperature sensor cools.
OVERHEAD PANEL
WING ANTI-ICE DUCT FAIL LIGHT
Red DUCT FAIL light comes on if
bleed air leak is detected in wing left
and right anti-icing ducts running
along fuselage.
OVERHEAD PANEL
, 14 STAGE ,
- , * RIGHT LEFT ' o o
• RIGHT F U S LEFT o o
(RIGHT
O
'RIGHT o
-WING*
-10 STAGE
IND RESET SYSTEM TEST O O
BLEED AIR LEAK
8LEED AIR
LEAK DETECT
DUCT
FAIL
PUSH TO TEST
BLSD AIR LEAX DETECT SWITCH/LIGHT
Red DUCT FAIL light flashes if a bleed air leak is
detected by any of the detection elements.
PUSH TO TEST-- When pressed, system is tested by
grounding detection circuit to simulate bleed air leak.
Flashing DUCT FAIL light on swttch/tight and steady
DUCT FAIL lights on bleed air and arm-ice panels
come on if leak detection system is serviceable.
BLEED AIR LEAK ANNUNCIATOR PANEL
Panel indicate* s have two positions: a black set
position when no fault exists and a white reset
position visible when there is a bleed leak in the
associated ducting.
Reset positions VB magneocaUy latched to remain on
after associated temperature sensor has cooled or
electrical power is removed from aircraft. Pressing
IND RESET button returns positions to set.
Pressing SYSTfcM TEST switch tests system by
grounding detection circuit to simulate bleed air leak.
Afi the DUCT FAIL lights come on and ail eight
indicatots on panel show white if leak detection
system is serviceable.
CENTRE INSTRUMENT PANEL
Bleed Air Leak Warning and Testing SECTION 17
Figure 13 Page 24
Apr 02/87
cttanenejer
OPERATING MANUAL
PSP 601A-6
IGNITION SWITCH/LIGHTS
When pressed in, arms associated igniter plug of both
engines for start and continuous igntion operation.
When pressed out, disarms associated igniter plug of both
engines for start and continuous ignition operation.
Green IGN A (or SGN B) light comes on immediate
associated switch/light pressed.
White ON lightfs) comets) on when associated igniter
plugs on one or both of the engines are in operation.
START SWITCH/UGHTS
Momentarily pressing switch/light causes green START
light to come on and initiates engine start sequence:
- opens Of not previously opened) left and right bleed air
shutoff valves and isolator valve.
- opens associated starter valve.
- fires pre selected A and/or B igniter plugis) on
associated engine, ignition white ON tightts) cornels)
on.
When engine reaches 55° N2, the starter automatic shutoff
switch:
- closes (unless selected open by associated control
switch) left and right bleed air shutoff valves and
isolator valve.
doses associated starter valve.
- turns off pre selected A and/or B igniter plugis) of
1 engine, ignition white ON bghtls) goies) out.
STOP SWITCH/LIGHTS
Momentarily pressing switch/light stops engine start
CONT IGN SWITCH/LIGHT
When pressed in. green CONT IGN light comes on and
continuous ignition is supplied to both engines through
IGN A and/or IGN B switch/tight(s). When pressed out.
CONT IGN light goes out and continuous ignition is turned
off.
IN FUGHT START SWITCH/UGHTS
When pressed in, fires both igniter plugs on associated
engine and green IN FUGHT START light and white ON
light come on.
^A/hen pressed out, turns off both associated igniter plugs,
green IN FUGHT START light and white ON light.
Amber STOP light comes on GO seconds after START
switch is pressed if engine has failed to start.
OVERHEAD PANEL
Engine Start and Ignition Controls SECTION 17
Figure 14 Page 25
Apr 02/87
OPERATING MANUAL
PSP 601A-6
wxm
ENGINE
VIBRATION
OEEN
EFFECTTVTTY: A/C 5001 TO |
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