|
OPERATING MANUAL
PSP 606
SECTION 17
POWER PLANT
TABLE OF CONTENTS
Subject
GENERAL
ENGINE FUEL SYSTEM
Engine-Driven Boost Pump and Motive Flow Pump
Fuel/Oil Heat Exchanger
Main Fuel Pump and Unit Fuel Control
Fuel Flowmeter Sensor
In-line Fuel Filter and Overspeed Fuel Shutoff Val
Fuel Flow Divider and Combustor Nozzles
Ecology Drain
ENGINE OIL SYSTEM
Oil Tanks
Pressure Delivery and Scavenge Pump Assembly
Oil Circulation
System Venting
Oil Filtering
ENGINE CONTROLS
Throttle Levers
Reverse Thrust Levers
Throttle Friction
ENGINE INSTRUMENTS
Signal Data Converter
Engine Instruments
Automatic Dimming
| Nl Fan Speed Indicator Compensator
|ENGINE BLEED AIR
Bleed Air Manifold
| Operation
Bleed Air Leak Detection and Warning System
STARTING AND IGNITION
Ground Starting
| In-Flight Starts
N2 rpm above 45%
Mach/Altitude within Windmilling Start Envelope
Mach/Altitude within Starter Assist Envelope
Continuous Ignition
| Engine Motoring (Fuel and Ignition Off)
canadair
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OPERATING MANUAL
PSP 606
Subject Page
THRUST REVERSING 27
Operation 29
ENGINE VIBRATION MONITORING SYSTEM 32
OVERSPEED PROTECTION 34
ENGINE SYNCHRONIZER SYSTEM 36
Operation
| Fault Warning 38
ICOWLINGS 38
Exterior Cowlings
Nose Cowl
Access Cowl Doors
Thrust Reverser Translating Sleeve 40
Interior Cowlings
Fan Duct Panels
| Core Cowls
ENGINE ANTI-ICING 41
Fan Spinner Anti-Icing
Splitter Ring and Inlet Vane Anti-Icing
|POWER PLANT DRAIN, VENT AND ECOLOGY SYSTEMS 41
Drains and Vents
| Ecology System 43
LIST OF ILLUSTRATIONS
Fi gure
Number Title Page
1 Power Plant - Schematic 2
2 Engine Fuel System - Schematic 4
3 Fuel Control Panel - Engine Fuel System Monitoring 6
4 Engine Oil System - Schematic 8
5 Oil Temperature and Pressure Indicators 9
6 Throttle Quadrant 10
17-
Page
Mar 01
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OPERATING MANUAL
PSP 606
Figure
Number Pi tl e Page
7
8
9
10
n
12
13
14
15
16
17
18
19
20
Engine Instruments
Engine Instruments Control Panel
Engine Bleed Air - Schematic
Bleed Air Control Panel
Bleed Air Leak Warning and Testing
Engine Start and Ignition Controls
Thrust Reverser Stowed and Deployed Positions
Thrust Reverser System - Schematic
Thrust Reversing Arming and Indicating
Engine Vibration Monitor Panel
Overspeed Protection System
Engine Synchronizer System Control Panel
Cowli ngs
Engine Anti-Icing
12
14
17
19
22
24
28
30
31
33
35
37
39
42
17-CONTENTS
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OPERATING MANUAL
SECTION 17
POWER PLANT
1. GENERAL (Figure 1)
The aircraft is powered by two Avco Lycoming ALF 502L-2 engines secured to the
rear fuselage by yoke structures bolted to the engine mounting torque box in
the rear fuselage equipment bay. The engines are twin spool turbofans with a
5:1 bypass ratio to provide low fuel consumption in cruise and improved single
engine take-off performance. The two spools, designated as the low pressure
(LP) and high pressure (HP) spools, are not connected mechanically, but are
related in operation by the air and fuel flow through the engine.
The engine airflow passes through a single-stage fan assembly and is divided
into two flow systems. The main airflow, bypass a i r , is routed by a fan duct
around the engine core and exhausts through the thrust reverser assembly. The
remaining airflow passes through the LP compressor into the engine core, which
consists of the HP compressor, combustion section and HP and LP turbine
assemblies. The hot gas is then exhausted through a primary exhaust nozzle and
mixed with the bypass exhaust a i r.
An automatic interstage air bleed system, located on the HP compressor casing,
bleeds HP compressor a i r , during transient phases of engine operation, to
prevent compressor s t a l l.
An accessory gear box, driven by the high pressure compressor through a bevel
gear and drive shaft, provides mountings for the power plant accessories and the
engine starter. Bleed a i r ports, located just to the rear of the high pressure
compressor, provide high pressure air for the aircraft's pneumatic system,
thrust reverser actuation, nose cowl anti-icing and cross feed engine
starting. The following additional features of the power plants are described
in this section:
Independent fuel control and distribution systems installed on each engine
Integral lubricating oil systems
The throttle quadrant assembly containing mechanically interlocked throttle
and thrust reverse levers
The engine instrument system featuring vertical scale and digital readout
displays
Engine starting and ignition systems
Pneumatically actuated thrust reverser assemblies which reverse the
direction of bypass flow thrust to assist aircraft braking
The engine vibration monitoring and warning system
SECTION 17
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OPERATING MANUAL
Power Plant - Schematic SECTION 17
Figure 1 Page 2
May 28/82
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OPERATING MANUAL
The engine overspeed protection system
The engine synchronizing system which automatically synchronizes the fan
speeds of the two engines
The engine thrust ratings are 7500 pounds thrust at take-off power and
7100 pounds at maximum continuous power.
2. ENGINE FUEL SYSTEM (Figures 2 and 3)
Each engine has a completely self-contained fuel system which meters fuel from
the aircraft1 s fuel tanks to the nozzles in the engine combustor section at the
correct pressure and rate of flow. The primary components of the system
include an engine-driven fuel boost pump and motive flow fuel pump, a fuel/oil
heat exchanger, a main fuel pump and fuel control unit and a fuel flow divider
assembly supplying metered fuel to the fuel nozzles in the combustor section.
Fuel flow, temperature and pressure sensors and an overspeed shutoff valve are
installed at suitable locations on the fuel system lines. The operation of the
system is monitored by advisory lights and indicators mounted on the fuel
control panel in the flight compartment (refer to Figure 3).
A. Engine-Driven Boost Pump and Motive Flow Pump
The combined engine-driven fuel boost pump and motive flow fuel pump
assembly is located on the rear face of the accessory gearbox and contains
two pump elements mounted side by side: a positive displacement gear pump,
which generates motive flow for the primary and scavenge ejectors of the
aircraft fuel system, and a centrifugal pump, which supplies fuel to the
main fuel pump via the fuel heater side of the fuel/oil heat exchanger.
The gear pump element contains a bypass valve which regulates the motive
flow discharge pressure.
B. Fuel/Oil Heat Exchanger
The fuel/oil heat exchanger is mounted on the top right side of the fan
casing and is divided into two sections: a fuel heater, which receives fuel
from the engine boost pump, and an oil cooler supplied with fuel from the
fuel control outlet. A thermal valve inside the fuel heater regulates the
amount of heat transferred to the fuel and a pressure-operated bypass valve
opens automatically i f the heater becomes obstructed.
C. Main Fuel Pump and Unit Fuel Control
The main fuel pump and fuel control unit is mounted on the front face of
the accessory gearbox and consists of a fuel control unit (FCU) with an
integral pump. The fuel pump, driven directly by the gearbox, boosts the
system fuel pressure to a value suitable for metering to the combustion
nozzles, and provides fuel servo pressure to the control devices within the
FCU.
SECTION 17
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OPERATING MANUAL
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LOW
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Engine Fuel System -
Figure 2
Schematic SECTION 17
Page 4
May 28/82
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OPERATING MANUAL
The FCU regulates high pressure compressor speed as a function of throttle
lever position, fan inlet temperature (T12) and high pressure compressor
discharge pressure (P3). When engine deceleration is commanded by the
throttle lever, the FCU provides a constant ratio of fuel flow to high
pressure compressor discharge pressure down to a preset minimum fuel flow,
An additional function of the FCU is to provide a control signal to the air
bleed actuator on the high pressure compressor. The actuator opens air
bleed ports at the sixth stage of the high pressure compressor during
starting, acceleration and low speed steady state operation of the engine
to prevent compressor surging and stalling.
The FCU is protected against particle contamination in the fuel by the main
fuel filter assembly, located in the fuel line to the FCU inlet. The
filter contains a disposable filtering element and a pressure-operated
bypass valve which opens automatically if there is excessive differential
pressure across the filter. At the same time, a differential pressure
switch causes the amber FILTER light on the fuel control panel to come on
to indicate filter bypass, (refer to Figure 3).
Fuel Flowmeter Sensor
An electrically operated fuel flowmeter sensor is located between the
outlet of the oil cooler and the overspeed shutoff valve. The sensor is
capable of measuring maximum fuel flows of 3500 pounds per hour. Fuel flow
indications below 185 pounds per hour are subject to inaccuracies generated
by the sensor (refer to Figure 7).
In-line Fuel Filter and Overspeed Fuel Shutoff Valve
An in-line fuel filter protects the fuel system components downstream from
the fuel flow meter. The filter assembly is integral with the overspeed
shutoff value and consists of a removeable housing and a replaceable filter
element. A bypass valve, located in the filter housing, allows fuel to
bypass the filter element, if required. An impending bypass indicator
button on the filter housing pops out when the filter differential pressure
is excessive.
The overspeed fuel shutoff valve consists of a solenoid-operated valve
which is energized by an electrical signal transmitted from an overspeed
control unit on the fan casing. When energized, the valve closes the fuel
line to the fuel flow divider to shut down the engine, and opens a bypass
line to direct any pressurized fuel back to the fuel heater inlet line.
SECTION 17
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OPERATING MANUAL
PSP 606
FUEL TEMPERATURE INDICATOR
Scale range: -20°C to 70°C.
Shows temperature at left and right fuel heater outlets.
FUEL CONTROL
— PUSH ON OFF 1
I r-l
X-FLOW
W)
1 VALVE-]
CLOSED
FILTER
LOW
PRESS l
L
fc
N
G
F
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L i
LOW PRESSURE WARNING LIGHTS
Amber warning light come son to indicate low pressure at
associated engine fuel inlet port.
VALVE CLOSED LIGHTS
White light comes on whenever associated firewall fuel
shutoff valve is closed.
FILTER BYPASS WARNING INDICATORS
Amber light comes on when fuel pressure drop is detected
across associated main fuel filter.
NOTE
Refer to FUEL for details of aircraft fuel
system control and monitoring.
Fuel Control Panel - Engine Fuel
System Monitoring
Figure 3
SECTION 17
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OPERATING MANUAL
F. Fuel Flow Divider and Combustor Nozzles
A fuel flow divider downstream of the overspeed shutoff valve meters fuel,
according to a predetermined schedule to the primary and secondary ducts of
the l e f t and right fuel manifolds. Each manifold houses 14 fuel nozzles
connected into the engine combustion section; each fuel nozzle houses
primary and secondary fuel atomizer parts. During engine start-up, fuel
flows through the flow divider to the fuel manifold primary ducts and to
the fuel nozzle primary atomizers. At approximately 13% N2 rpm, a valve
in the flow divider begins to open, allowing fuel to flow through the fuel
manifold secondary ducts, to the fuel nozzle secondary atomizers. As
secondary fuel flow increases, primary fuel flow decreases until, at full
secondary fuel flow, primary fuel flow is reduced by 60%.
G. Ecology Drain
The ecology drain consists of a drain tank and float valve assembly located
on the bottom of the engine to collect the fuel which pools in the
combustor section following engine shutdowns or aborted starts. An
engine-mounted ejector pump automatically removes the fuel collected in the
drain tank and returns i t to the inlet of the engine fuel system.
3. ENGINE OIL SYSTEM (Figures 4 and 5)
Each engine is lubricated and cooled by i t s own self-contained oil system,
which consists of an oil tank, a pressure delivery and scavenge pump assembly,
f i l t e r s and a fuel/oil heat exchanger. Externally mounted lines and internal
channels direct pressurized oil from the pressure delivery pump to the varioir
lubrication points within the engine. Flight compartment instruments (refer
Figure 5) monitor oil temperature and pressure; warning lights, mounted on tne
oil pressure indicator, provide low oil pressure warnings.
A. Oil Tanks
Each oil tank has a capacity of 3.75 gallons (US) (3.12 imperial gallons,
14.2 litres) and is installed on the l e f t side of the engine fan casing.
A f i l l e r neck, located on the outside of the tank, prevents over-filling.
When the oil system is fully replenished, the tank contains 3.0 gallons (US)
2.5 imperial gallons, 11.4 litres) of o i l , leaving the remaining tank
volume available for oil expansion and a i r removal. The oil level in the
tank is checked with a dipstick attached to the f i l l e r cap, or visually
through two sight gauges.
B. Pressure Delivery and Scavenge Pump Assembly
The pressure delivery and scavenge pump assembly is driven by the accessory
gearbox and contains three pump elements: the pressure pump which provides
flow of pressurized oil to the lubrication points in the engine, and two
scavenge elements, the main and bearing scavenge pump, which direct
scavenged oil along a common return line to the oil tank. A bypass l i n e,
which incorporates a pressure regulating valve, prevents overpressure at
the outlet of the pressure pump.
SECTION 17
Page 7
May 28/82
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OPERATING MANUAL
LP SPOOL
TURBINE
BEARINGS
LEGEND
LUBRICATING OILSUPPLY
PUMP OIL SUPPLY
SCAVENGE OIL
l l l l l l l l l l l l l TANK VENT LINE
LEVEL
SIGHT
GAUGES
Engine Oil System - Schematic
Figure 4
SECTION 17
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May 28/82
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OPERATING MANUAL
OIL PRESSURE INDICATOR
Vertical scale indicator which displays oil pressure as detected by pressure
transmitter located between main oil filter and engine lubrication points.
Coloured light segments of vertical scales come on to indicate the following
range:
Low pressure warning range (red) 0 to 30 psi
Cautionary pressure range (yellow) 30 to 40 psi
Normal operating range (green) 40 to 120 psi
High pressure warning range (red) 120 to 130 psi
OIL TEMPERATURE INDICATOR
Vertical scale indicator. Displays oil temperature as detected by
temperature sensor located between main oil filter and engine
lubrication points. Coloured light segments of vertical scales come
on to indicate the following ranges:
LOW OIL PRESSURE LIGHTS
Red warning lights come on when oil pressure of
associated engine drops below 20 psi.
Normal operating range (green)
Warning range (red)
-20°Cto140°C
140°Cto180°C
Oil Temperature and Pressure Indicators
Figure 5
SECTION 17
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OPERATING MANUAL
PSP 606
THRUST REVERSE (TR) LEVERS
Select and regulate reverse thrust- Throttle
interlock solenoids prevent TR lever
movement beyond deploy position until
reverser assemblies are fully deployed.
Maximum reverse thrust stop at 92-V2
degrees from stowed position of TR lever.
THRUST REVERSE LEVER RELEASE LATCHES
Extend fingers under TR lever grips and lift latches
to release TR levers from stow locks.
GO-AROUND SWITCHES
Momentary push button switches. Pressing
either switch disengages automatic flight
control system and places HSI in the goaround
mode.
THROTTLE SAFETY LOCK
Prevents Throttle lever from
advancing beyond HIGH IDLE
when aircraft is airborne and
the thrust reverser is not
fully stowed.
THROTTLE LEVERS
Control forward thrust. Remain locked at
LOW IDLE position during thrust reverser
operation.
THROTTLE LEVER RELEASE LATCHES
Lift to advance throttle levers from SHUT
OFF position or retard throttle levers from
LOW IDLE position.
*^i
THROTTLE LEVER FRICTION ADJUSTMENT
Adjusts friction on throttle levers only. Rotate
control clockwise to increase friction.
THROTTLE SETTINGS
SHUT OFF - Located at rear throttle stop. Acts as engine fuel shut-off position.
LOW IDLE - Lowest forward thrust setting. When moving throttle forward
from SHUT OFF or rearward from higher power settings, LOW IDLE is
encountered as positive stop which is released by lifting throttle lever thumb
latch.
HIGH IDLE - Felt as detent as throttle is retarded. Serves as reference for
pilot, indicating approach of idle power settings as throttle is retarded.
Detent overcome by rearward pull of 5 to 8 pounds at throttle grip.
MAX POWER - Highest forward thrust setting. Located at forward throttle
lever stop.
Throttle Quadrant SECTION 17
figure 6 Page 10
Mar 01/85
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OPERATING MANUAL
C. Oil Circulation
Oil flows from the bottom of the oil tank through an external line to the
inlet of the pressure delivery pump. The pressurized oil is then directed
to the main oil filter and through distribution lines to the engine
bearings, accessories and LP spool reduction gear assembly. Scavenge oil
from the forward part of the engine and the rear high pressure turbine
bearing, drains into a sump on the accessory gearbox and is scavenged by
the main scavenge pump. Oil from the low pressure turbine bearings, at the
rear of the engine, is recovered by the No. 4 & 5 bearing scavenge pumps.
The discharge from the two scavenge pumps passes to a common return line
and is routed across both sections of the fuel/oil heat exchanger before
returning to the oil tank. The heat exchanger maintains oil temperature
within operational limits at all engine speeds and under environmental
extremes.
D. System Venting
The combined pumping capacity of the main and number two scavenge pumps is
approximately four times as large as the maximum system requirement. The
extra scavenge pump capacity produces a return flow, which contains a
relatively large volume of air. The air is separated from the oil by a
swirl chamber inside the accessory gear box and vented to atmosphere. A
line connecting the oil tank and the accessory gearbox vents the air space
in the oil tanks.
E. Oil Filtering
The oil system contains three filters: a main oil f i l t e r on the lubricating
oil supply line, and two filters located on the HP spool rear bearing and
the LP turbine bearings respectively. The main oil f i l t e r incorporates a
bypass valve which opens automatically to maintain the supply of lubricating
oil i f the f i l t e r becomes blocked.
4. ENGINE CONTROLS (Figure 6)
The engine controls consist of a throttle quadrant located on the centre
pedestal and the mechanical linkages between the throttle quadrant and the
engine fuel control units. Two throttle levers control forward thrust and two
thrust reverse (TR) levers control the operation of the thrust reversers.
Controlex push-pull cables run from the throttle quadrant under the cabin floor
to the engine pylon firewalls. At the firewalls, the cables terminate in
disconnect fittings connected to teleflex throttle controls which complete the
cable run to control boxes mounted forward of the FCUs. Rod assemblies link
the control boxes with the FCU power levers. Pressure seals are installed at
the fuselage sides where the cables pass through the pressure shell.
SECTION 17
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OPERATING MANUAL
PSP 606
NOTE
Refer to Figure 5 for OIL TEMP and
OIL PRESS operating ranges.
ITT INDICATOR
Normal operating range (green)
Caution range (yellow)
Warning range (red)
N1 % RPM INDICATOR
Normal operating range (green)
Warning range (red)
E3
BUGS
Manually set to desired references on N1 and
ITT indicators.
N2 % RPM INDICATOR
Low speed caution range (yellow) 0 to 53%
Normal operating range (green) 53 to 96%
High speed caution range (yellow) 96 to 98%
Warning range (red) 98 to 110%
FUEL FLOW INDICATOR
Normal operating range (green) 0 to 3500 pph
POT
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T.O./NORM SWITCH
Two position toggle switch.
T.O. - Fan speed indicator
compensation is on.
NORM - Fan speed indicator
compensation is off. Actual N1
speeds are shown on indicator.
DIGITAL DISPLAY ON/OFF DIGITAL DISPLAYS
SWITCH
Two position toggle switch
controls N1, ITT, N2and
FUEL FLOW digital displays
on and off.
Three figure readouts.
Displays are not included on
OIL TEMP and OIL PRESS
indicators. FUEL FLOW
display is given in multiples
of ten.
VERTICAL SCALES
Scales consist of light segments illuminated
by miniature lamps inside instrument. Light
segments are colour coded to indicate
normal operating, caution and warning
ranges.
POWER ON LIGHTS
Blue light segments located at bottom of
each vertical scale. Lights come on
whenever vertical scales and associated
digital readouts are receiving adequate
electrical power from SDC.
EFFECTIVITY
[ l ] Aircraft 1072, 1086 and subsequent and aircraft
incorporating Canadair Service Bulletin 600-0350.
Engine Instruments
Figure 7
SECTION 17
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OPERATING MANUAL
PSP 606
Throttle Levers
Forward thrust is controlled by moving the throttle levers between throttle
positions LOW IDLE, HIGH IDLE and MAX POWER. A fuel shutoff position, SHUT
OFF, is located at the rear throttle stop. The throttle levers are moved
forward from LOW IDLE without restriction until they encounter a positive
stop at the MAX POWER position. The intermediate position, HIGH IDLE, is
felt as a shallow detent as the throttles move forward through it. When
the throttle levers are retarded, a detent is encountered at HIGH IDLE and
a positive stop at LOW IDLE. The detent at HIGH IDLE is overcome by a
rearward force of 5 to 8 pounds applied at each throttle grip. The
positive stop at LOW IDLE is released by lifting the release latches under
the throttle grips. The throttle levers can then be retarded without
restriction to the SHUT OFF position.
Reverse Thrust Levers
The thrust reversers are deployed after the throttle levers have been
retarded to the LOW IDLE position. Reverse thrust is controlled by moving
the thrust reverse (TR) levers rearwards with the throttle levers at LOW
IDLE. The TR lever locks are released by lifting the TR lever release
latches allowing the levers to be pulled back to the deploy position. This
action locks the throttle levers at LOW IDLE but moves the throttle control
output linkages to the engine fuel control units (FCU) to a position
corresponding to high idle thrust. If the TR levers are operated with the
throttle levers set above LOW IDLE, feedback mechanisms, activated during
thrust reverser deployment, move the throttle levers rapidly back to LOW
IDLE. Interlock solenoids in the throttle quadrant prevent the TR levers
from being moved to full reverse thrust during the thrust reverser
deployment phase. When the thrust reversers are fully deployed, the
interlocks release the TR levers and the levers can be pulled rearwards to
give the desired amount of reverse thrust.
The thrust reverser stow sequence is initiated when the TR levers are
pushed fully forward to engage the TR lever locks. Freedom of movement is
returned to the throttle levers at the same time. A throttle safety lock
system prevents movement of the throttle levers above the LOW IDLE
position, if the thrust reversers are not fully stowed.
Throttle Friction
A throttle friction device built into the system includes an adjustment
control located at the base of the throttle quadrant. Turning the control
clockwise increases friction on the throttle levers only, eliminating
throttle creep due to control loads and vibration. Friction is reduced to
a minimum by rotating the control counterclockwise.
SECTION 17
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Mar 01/85
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OPERATING MANUAL
PSP 606
'- c_ R—1
PHOTOCELL
Part of automatic dimming control. Adjusts brightness
of the instrument displays with reference to
ambient fight in flight compartment.
POWER FAILURE WARNING LIGHT
Comes on whenever one of the dual
instrument power supplies fails.
TEST TOGGLE SWITCH
Each test position checks one of the engine instrument power sources.
On each instrument one digital readout comes on indicating 888 and
the opposite side vertical scale comes on indicating a full scale reading.
The remaining digital readouts and vertical scales are tested by setting
the switch to the second test position.
At both test positions the power failure warning light comes on, the
three left digits of each digit element on the fuel quantity indicator
show 8, and the remaining digits show 0.
DIMMING CONTROL KNOB
Permits adjustment of ambient to output
brightness ratio of displays to individual
preferences.
Engine Instruments Control Panel SECTION 17
Figure 8 Page 14
Mar 01/85
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OPERATING MANUAL
5. ENGINE INSTRUMENTS (Figures 7 and 8)
Six engine instruments monitor the following parameters: fan (Nl) rpm,
inter-turbine temperature (ITT), high pressure compressor (N2) rpm, fuel flow,
oil pressure and oil temperature. The indicator systems employ solid state
signal processing and electronic displays to eliminate the moving parts
associated with conventional dial or vertical tape instruments.
The principal components of the system include six indicating instruments and
an engine instrument control panel in the flight compartment, a signal data
converter (SDC) in the underfloor avionics bay, and various sensing devices
mounted on the engines. Dual power supplies are provided for the display and
signal processing circuits within the instruments.
A* Signal Data Converter
The SDC serves as the power supply for the engine instrument system. Two
28-volt dc inputs, from the battery bus and the 28-volt dc essential bus,
are divided within the SDC into dual lamp and signal processing power
supplies. The power supplies provide voltage-regulated dc power to the
display and signal processing circuits within the six instruments. The SDC
provides ambient temperature compensation for the ITT indicator.
The SDC also serves as the power supply for the fuel quantity indicator.
Instrument power supply fuses and components of the automatic dimming
control circuits are also located in the SDC.
B. Engine Instruments
Each instrument provides a vertical analog display of the relevant engine
parameter. Left and right engine displays on each instrument are separated
by a common central scale. The Ni, ITT, N2 and FUEL FLOW instruments
also contain digital readout displays below the vertical scales. The
vertical scales are colour-coded to indicate the normal operating,
cautionary and warning ranges of each system.
The vertical scales consist of coloured plastic light segments connected by
a fibre-optic system to an array of miniature incandescent lamps behind the
instrument display face. The electronic signal processing circuits inside
the instrument cause the lamps to come on in response to variations in the
signal received from the sensing device on the engine. The light generated
by the incandescent lamps is transmitted through the fibre-optic system to
the light segments on the display face of the instrument to produce the
vertical scale reading.
The signal processing circuits operate in a similar manner to produce the
digital displays below the vertical scales. The three-digit readouts
provide more accurate indications when compared with the readings on the
vertical scales.
SECTION 17
Page 15
May 28/82
OPERATING MANUAL
PSP 606
In the case of the Nl, ITT, N2 and FUEL FLOW instruments, the dual lamp
power sources supplied by the SDC create display redundancy in the
following manner: one of the lamp power sources supplies power to one of
the digital readouts and to the vertical scale on the opposite side of the
instrument. The other power source powers the second digital readout and
the remaining vertical scale. If one of the power sources fails, each
instrument loses one digital readout and the opposite side vertical scale
In the case of the OIL TEMP and OIL PRESS instruments, each power source
powers alternate light segments in the vertical scales. If one of the
power sources fails, alternate segments on the vertical scales remain on.
Additional features of the instruments include the central scale markings
illuminated by the instrument integral lighting system and blue power on
lights on the bottom of each vertical scale, which come on whenever the
instruments are receiving adequate power from the SDC.
Automatic Dimming
The engine instrument control panel, located beside the fuel control panel,
contains the system auto-dimming controls. A photocell on the control
panel monitors the ambient light level in the flight compartment and,
through a feedback circuit, automatically adjusts the brightness of the
displays to ensure optimum readability. A manual dimming control, also on
the control panel, allows the ambient to output brightness ratio of the
displays to be adjusted to individual preferences. The automatic and
manual dimming controls are operated through separate electronic circuits
so that one of the controls remains in operation if a failure of the other
occurs.
The test and warning functions of the engine instrument control panel are
shown in Figure 8.
Nl Fan Speed Indicator Compensator
Aircraft 1072, 1086 and subsequent and aircraft incorporating Canadair
Service Bulletin 600-0350 are fitted with an Nl fan speed indicator
compensator system. The system consists of a resistor, calibrated by the
manufacturer, on each engine overspeed controller and a two position
T.O./NORM toggle switch on the Nl %RPM indicator. For engines that produce
more than rated thrust at a given Nl rpm, the system will bias the affected
Nl %RPM indicators to read up to 2% high if the T.O./NORM switch is in the
T.O. position. When the T.O./NORM switch is in the NORM position, the
Nl %RPM indicator shows actual Nl rpm.
The system is normally used to set take-off thrust. With the T.O./NORM
switch in the T.O. position, take-off thrust is obtained from both engines,
without the possibility of an overthrust, if both engines are set at the Nl
rpm shown on the appropriate take-off thrust curve in the Airplane Flight
Manual.
SECTION 17
Page 16
Mar 01/85
canadair
chaHentjer
OPERATING MANUAL
PSP 606
LEGEND
RIGHT ENGINE BLEED
BLEED AIR
PRESSURE SENSING
APU AND GROUND AIR
TO CABIN
PRESSURIZATION
CONTROL
TO EMERGENCY
PRESSURIZATION
FOOTWARMER
AND WINDSHIELD
DEMIST
PRESSURE REGULATOR
AND SHUTOFF VALVE
GROUND AIR SUPPLY LEFT ENGINE BLEED
Engine Bleed Air - Schematic
Figure 9
SECTION 17
Page 17
Mar 01/85
ctiaitemjer
OPERATING MANUAL
PSP 606
6. ENGINE BLEED AIR
The engine bleed air system consists of a bleed air manifold which connects
supply and distribution ducting, electrically controlled and pneumatically
| operated valves and switch/light controls on the BLEED AIR and ANTI-ICE panels
in the flight compartment. The system is protected by a bleed air leak
detection system which provides a warning to the flight compartment i f a leak
occurs in the bleed air ducting.
A. Bleed Air Manifold (Figure 9)
The bleed air manifold consists of a series of ducts clamped together and
secured to the fuselage structure by clamps and t ie rods behind the rear
pressure bulkhead. Compressed air can be supplied to the manifold from the
left and right engines, from the APU or from an external source connected
to the APU fault panel and ground air connection under the left engine
(refer to Section 1, AIRCRAFT GENERAL, Figure 7).
The manifold is divided into four sections by the shutoff control valves
shown on Figure 9 which can be opened or closed to supply or isolate the
airflow to the various aircraft systems. The following aircraft services
make use of pressurized air tapped from the manifold:
Air conditioning and pressurization, supplied by lines from the left
and right sections
Wing anti-icing, supplied by lines from the lower section
Footwarmers, windshield demisting and emergency pressurization,
supplied by a line from the crossover section
Bleed air required for engine starting, thrust reverser actuation and nose
cowl anti-icing is tapped from lines between the engine supply duct and the
left or right bleed air shutoff valves. Small-diameter lines direct bleed
air from the manifold and the left engine to a j e t pump in the cabin
pressurization control system.
Two pressure transducers, one for each engine bleed air supply, are located
in the bleed air manifold downstream from the respective bleed air shutoff
valves. The transducers transmit pressure signals to the corresponding L
and R sections of the dual pointer pressure indicator on the BLEED AIR
panel.
All of the bleed air system valves are electrically controlled through
integral solenoids and pneumatically actuated ( i . e . the solenoids must be
energized and the valves pressurized before the valve ports open).
SECTION 17
Page 18
Mar 01/85
canactair
chanenQer
OPERATING MANUAL
PSP 606
DUCT MON SWITCH
Three-position DUCT MON toggle switch tests serviceability
of each of the detector loops A and B on the left and right
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.
BLEED AIR ISOL SWITCH/LIGHT
When pressed, upper bleed air isolation valve opens.
Valve closes when switch /light is pressed again.
Green OPEN light comes on whenever upper bleed air
isolation valve is open.
CKPT HEAT SWITCH
Three-position CKPT HEAT toggle switch controls
position of left and right footwarmer/demist pressure
regulating shutoff valves.
NORM - Right valve opens.
STBY - Left valve opens and right valve closes.
OFF - Left and right valves remain closed.
Selecting emergency pressurization on cabin
pressurization control panel opens left and closes right
valve, over-riding CKPT HEAT switch settings.
BLEED AIR PRESSURE GAUGE
Indicates pressure in bleed air manifold to the left and
right of upper bleed air isolator valve.
L ENG AND R ENG BLEED CLOSED SWITCH/LIGHTS
When pressed in, associated bleed air shutoff valve
opens and white BLEED CLOSED light goes out. When
pressed out, valve closes and light comes on.
Red L ENG of R ENG DUCT FAIL light comes on if the
bleed leak detection elements detect a failure in the
associated duct segment. Light goes out when the failed
duct is isolated and detection element cools.
Bleed Air Control Panel SECTION 17
Figure 10 Page 19
Mar 01/85
OPERATING MANUAL
PSP 606
B. Operation (Figure 10)
The engine bleed air controls consist of switch/lights on the BLEED AIR
| panel identified, from left to right, as L EN6, ISOL and R ENG.
The left and right bleed air shutoff valves are opened and closed by
pressing the L ENG and R ENG switch/lights. Pressing one of the
| switch/lights opens the associated shutoff valve and causes the white BLEED
CLOSED light to go out. The valve is closed by pressing the switch/light a
second time. Both bleed air shutoff valves open automatically when either
of the START switch/lights on the ENGINE START panel is pressed during the
engine start sequence.
I If the L ENG switch/light is pressed in with the upper bleed air isolation
valve open or closed, or if the R ENG switch/light is pressed in with the
valve open, an electrical interlock automatically operates to shut the APU
load control valve. When the APU is used as a pneumatic source during
engine starting, the L ENG and R ENG switch/lights must be off.
| The upper bleed air isolator valve is used to isolate the left and right
manifold ducts. The valve is opened by pressing in the ISOL switch/light
and closed by pressing the switch/light a second time. The green OPEN
light comes on whenever the valve is open.
I The lower isolator valve is normally left closed but can be opened, if an
I anti-ice valve fails to open or an engine fails, by pressing in the ISOL
I OPEN switch/light on the ANTI-ICE panel (refer to Section 14, ICE/RAIN
I PROTECTION).
Normally the crossover manifold is pressurized by setting the CKPT HEAT
| switch on the BLEED AIR panel to NORM to open the right footwarmer/demist
pressure regulator and shutoff valve. Alternatively, the left valve can be
opened by setting the switch to STBY to extract bleed air directly from a
line ahead of the left bleed air shutoff valve. When the switch is set to
STBY, the right valve will close if open. Selection of emergency
pressurization on the cabin pressurization control panel opens the left and
closes the right valve regardless of the position of the CKPT HEAT switch.
The left valve is located ahead of the left bleed air shutoff valve so that
emergency pressurization is available regardless of any failure in the
bleed air manifold or right engine.
The two anti-icing valves on the manifold rear section are opened by the I three-position WING switch on the ANTI-ICE panel (refer to Section 14,
ICE/RAIN PROTECTION).
SECTION 17
Page 20
Mar 01/85
canadair
chanenQer
OPERATING MANUAL
PSP 606
Bleed Air Leak Detection and Warning System (Figure 11)
Because of the high temperature of the air passed through the bleed air
manifold and the anti-icing ducts, a leak detection and warning system is
provided. The flight crew can depressurize and isolate a defective duct.
The system consists of heat-sensitive detection elements attached to the
bleed air ducts and electrically connected to two bleed air leak detection
control units in the underfloor avionics bay. The control units contain
dual detection loop circuits for the l e f t , right and crossover sections of
the bleed air manifold, and single loop circuits for the rear section and
the anti-icing ducts running through the fuselage and wings. On the l e ft
and right sections the detection elements are attached to the exterior of
the metallic insulating material surrounding the ducts. If a leak occurs,
the hot bleed air escapes through regularly spaced holes in the insulating
material and flows across the detection elements to i n i t i a te a warning
signal. In general, any of the detection elements initiates a warning
signal i f its impedance drops below a preset value.
The bleed air leak detection control units receive warning signals from the
detection elements and activate the following flight compartment warning
indicators:
The BLEED AIR LEAK DETECT switch/light on the centre instrument
panel. The red DUCT FAIL light of the switch/light flashes whenever a
bleed air leak is detected by any of the detection elements. The
switch/light also includes a system test function (refer to Figure 11).
The red DUCT FAIL lights on the BLEED AIR control panel. The DUCT
FAIL lights come on i f bleed air leakage is detected on the l e f t and
right sections of the bleed air manifold (refer to Figure 10).
The red DUCT FAIL light on the anti-ice control panel. The light comes
on i f a leak is detected in the wing anti-icing ducts.
The bleed air leak annunciator panel behind the copilot's seat. The
panel is used primarily for fault isolation and contains seven latching
magnetic indicators. Each indicator has two positions, a set black
position for the no-fault condition and a white reset position which
appears after a bleed leak has been detected. The reset position
remains showing on the indicator after the associated detection element
has cooled and electrical power has been removed from the aircraft.
The indicators are returned to the set position by pressing the
IND RESET button on the panel.
With the exception of the indicators on the bleed air leak annunciator
panel, all of the warning indicators go out when their associated detection
elements have cooled sufficiently. Testing of the bleed air leak detection
system is summarized in Figure 11.
SECTION 17
Page 21
Mar 01/85
canaaair
chanenQer
OPERATING MANUAL
PSP 606
EFFECTIVITY
H I Panel on A/C incorporating SB 600-0495.
For panel on other A/C, refer to Section 14.
ANTI-ICE CONTROL PANEL [7]
WING ANTI-ICE DUCT FAIL LIGHT
Red DUCT FAIL light comes on if bleed
air leak is detected in airfoil anti-icing
ducts running along fuselage.
©
ANTI-ICE
LEFT REAR
FUS
O
I
RIGHT
FUS o
RIGHT
WING o
.MANIFOLD
LEFT RIGHT1
©
FUS o
LEFT
WING
O
o o
IND RESET SYSTEM TEST
®
® ©
BLEED AIR LEAK
©
BLEED AIR
LEAK DETECT
DUCT
FAIL
PUSH TO TEST
©
BLEED AIR LEAK 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 elements to simulate bleed air leak.
Flashing DUCT FAIL light on switch/light and steady
DUCT FAIL lights on bleed air and anti-ice panels come
on if leak detection system is serviceable.
BLEED AIR LEAK ANNUNCIATOR PANEL
Panel indicators 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 manifold sections.
Reset positions are magnetically latched to remain on
after associated detection element has cooled or
electrical power is removed from aircraft. Pressing
IND RESET button returns positions to set.
Pressing SYSTEM TEST switch tests system by
grounding all detection elements to simulate bleed air
leak. All the DUCT FAIL lights come on and all seven
indicators on panel show white if leak detection
system is serviceable.
Bleed Air Leak Warning and Testing
Figure 11
SECTION 17
Page 22
Feb 12/88
canadair
chauenQer
OPERATING MANUAL
PSP 606
STARTING AND IGNITION (Figure 12)
The engine starting and ignition systems consist of a pneumatically driven air
turbine starter (ATS), ignition exciter boxes and igniter plugs. The systems
are controlled electrically by switch/lights in the flight compartment.
The ATS, bolted to a drive pad on the accessory gearbox, transmits starting
torque to the HP spool through the accessory gearbox drive shaft. The unit
contains a single-stage axial turbine connected through reduction gearing and a
clutch assembly to an output shaft. An electrical cut-out switch inside the
ATS, actuated by a flyweight governor, and the clutch assembly operate to shut
down the ATS at a preset output shaft rpm and protect it against turbine
overspeed. Pressurized air, supplied through the bleed air manifold, enters
the ATS through the start control valve (SCV).
The dual ignition system on each engine includes two igniter plugs installed at
the 5-o'clock and 7-o'clock positions in the engine combustor section. The
plugs are fired by pulsed high energy dc electrical power provided by two
exciter boxes, designated A and B, attached to the fan casing. Shielded high
tension cables connect exciter box A to the left igniter plug and exciter box B
to the right igniter plug.
Electrical power for both the starting and ignition components of the system is
provided by the battery bus.
A. Ground Starting
Electrical power for engine starting is available from three sources on the
aircraft; the battery, the APU generator or the integrated drive generator
(IDG) of an operating engine. External electrical power can be connected,
if necessary, at the dc external power receptacle below the right engine.
The bleed air manifold is capable of supplying pressurized air for engine
starting from the APU, from an operating engine or from an external source
connected to APU fault and air start panel under the left engine.
Normally, the APU generates the required electrical and pneumatic services
for starting.
The IGN A and IGN B switch/lights arm the exciter boxes and igniters that
are used during engine start or continuous ignition. Pressing the IGN A
switch/light causes the green IGN A light to come on immediately,
indicating that exciter box A and the left igniter plug of both engines are
armed. Pressing the IGN B switch/light similarly amis exciter box B and
the right igniter plug of both engines and causes the green IGN B light to
come on.
As ground starts can be accomplished using only one of the igniter plugs,
IGN A and IGN B should be used alternately during successive engine starts
to extend the service life of the ignition components.
SECTION 17
Page 23
Mar 01/85
canactair
ctiaiiencjer
OPERATING MANUAL
PSP 606
CONT IGN SWITCH/LIGHT
When pressed, green CONT IGN light comes on and
continuous ignition is supplied to both engines through
IGN A and/or IGN B switch/lights.
IGNITION SWITCH/LIGHTS
Pressing IGN A switch/light arms exciter box A and left
igniter plug of both engines for start and continuous ignition
operation.
Pressing IGN B switch/light arms exciter box B and right
igniter plug of both engines for start and continuous ignition
operation.
Green IGN A and IGN B lights come on immediately when
associated switch/lights are pressed.
Blue ON lights come on when associated igniter plugs on one
or both of the engines are in operation.
START SWITCH/LIGHTS
Pressing switch/light causes green START light to
come on and initiates engine start sequence by
energizing start and ignition relays.
STOP SWITCH/LIGHT
Pressing switch/light stops engine start sequence.
Amber STOP light comes on 30 seconds after START switch
is pressed if engine has failed to start.
IN FLT START SWITCH/LIGHT
When pressed in, fires both igniter plugs on associated
engine and green IN FLT START light comes on.
RELIGHT SWITCH
Setting switch to ON fires both igniter plugs on both engines.
Plugs continue to fire until switch is returned to OFF.
REVERSE THRUST
LEFT RIGHT
1 UNSAFE 1
TO ARM
I! ARMED
UNSAFE 1
TO ARM
ARMED 1
'«- PUSH TO ARM -
©
Engine Start and Ignition Controls SECTION 17
Figure 12 Page 24
Mar 01/85
canadair
chaiiencjer
OPERATING MANUAL
PSP 606
Ignition selections can be cancelled by pressing the IGN A or IGN B
switch/light a second time. The IGN A or IGN B light goes out and the
associated ignition components do not operate when the START switch/light
is pressed.
With the bleed air manifold pressurized, and assuming that the IGN A/ON
switch/light has been pressed, pressing the left START switch/light
initiates the left engine starting sequence as follows (The starting
sequence for the left engine is described. The starting sequence for the
right engine is similar except where noted):
The green START light comes on and 28-volt dc power from the battery
bus is supplied to a 30-second time delay relay between the left START
and STOP switch/lights.
The left engine start, the bleed air and the ignition A on relays are
energized through the closed contacts of the left STOP switch/light.
Power is supplied through the contacts of the energized relays to
operate exciter box A and the left igniter plug of the left engine,
and open the following valves: the left and right bleed air shutoff
valves, the bleed air isolator valve and the left SCV.
The left SCV opens and the SCV position indicator switch closes
latching the start and ignition relays through the electrical cut-out
switch in the ATS.
The left ATS begins to rotate bringing the left engine up to starting
speed. At 15% N2 rpm the left throttle lever is moved from SHUT OFF
to LOW IDLE. At 45 to 47% N2 rpm the electrical cut-out switch in the
ATS opens, de-energizing the left engine start, the bleed air and
ignition A on relays.
The left START light goes out and the left and right bleed air shutoff
valves, the bleed air isolator valve and the left SCV close.
As engine speed overtakes the speed of the ATS output shaft, the ATS
clutch assembly opens and the ATS runs down.
The start sequence is completed when the ATS is disengaged from the
engine. Stabilized LOW IDLE speed on the ground is 42.5 to 53.5% N2 rpm on
a standard day at sea level.
I f the engine fails to start within 30 seconds after the START switch/light
has been pressed, the 30-second time-delay relay closes, causing the left
amber STOP light to come on. Pressing the STOP switch/light de-energizes
the left engine start, the bleed air and the ignition A on relays to stop
the engine start sequence. When the left SCV closes, the contacts of the
SCV position indicator switch open and the left START and STOP lights go
out. As the lights go out the time-delay relay re-opens and the system is
ready for another start attempt.
SECTION 17
Page 25
Mar 01/85
canadair
chauenQer
OPERATING MANUAL
PSP 606
The STOP light comes on to indicate a fault in the starting system or the
use of improper start procedures. Allowing the start attempt to continue
for any length of time after the light comes on could damage system
components.
In-Right Starts
Depending on the N2 rpm, Mach number and altitude, the start and ignition
controls can be used in one of the following ways to start an engine in
flight.
(1) N2 rpm above 45%
If the N2 rpm has not decreased below 45%, an immediate in-flight
start may be attempted by setting the throttle lever to HIGH IDLE and
using either the RELIGHT switch on the centre pedestal or the
appropriate IN FLT START switch/light to energize the igniter plugs.
(2) Mach/Altitude within Windmilling Start Envelope
If the aircraft Mach/altitude is within the windmilling start envelope
and the engine is windmilling at 9 to 17% N2 rpm, the engine is
started by pressing the appropriate IN FLT START switch/light and
advancing the throttle lever to HIGH IDLE. Both of the igniter plugs
on the engine operate and the IGN A/ON and IGN B/ON lights come on.
The igniter plugs continue to operate until the IN FLT START
switch/light is pressed a second time.
(3) Mach/Altitude within Starter Assist Envelope
When an assist from the ATS is required, the start is accomplished by
pressurizing the bleed air manifold from the operating engine or from
the APU and pressing the IGN A, IGN B and START switch/lights. The
start sequence continues as in a normal ground start (refer to Ground
Starting).
Continuous Ignition (Figure 12)
Continuous ignition is obtained by pressing the CONT IGN switch/light and
one or both of the ignition switch/lights. Dual continuous ignition is
applied on both engines by pressing the IGN A, the IGN and CONT IGN
switch/lights. The green IGN A, IGN B and CONT IGN lights come on and
igniter boxes A and B fire their respective igniter plugs continuously in
both engines. The igniter plugs fire until the CONT IGN switch is pressed
a second time.
SECTION 17
Page 26
Mar 01/85
cacnhaadnaeinr tjer
OPERATING MANUAL
PSP 606
D. Engine Motoring (Fuel and Ignition Off) (Figure 12)
The engine can be dry-motored with fuel and ignition off by pressurizing
the bleed air manifold and pressing the START switch/light. The green
START light comes on and the ATS rotates the engine until the STOP
switch/light is pressed. The engine turns at a steady N2 rpm of 18 to
22%. Motoring may be continued for 5 minutes, followed by a 20-minute ATS
cooling period.
THRUST REVERSING (Figures 13, 14 and 15)
Each engine is equipped with a thrust reverser to assist in aircraft braking
after landing. Thrust reversing is accomplished by directing the fan exhaust
air, which constitutes the greater part of the total engine thrust, forward
through cascade vanes on the reverser fixed support structure. The principal
mechanical components of each reverser are a translating sleeve assembly,
flipper doors, blocker doors and three mechanical actuators connected by a
flexshaft system to a pneumatic drive unit (PDU). The PDU contains an air
motor, driven by high pressure air from the engine bleed air system.
Thrust reversing controls consist of the thrust reverse levers and their
associated microswitches on the throttle quadrant and arming switches on the
centre pedestal. Switch/lights on the glareshield advise the flight crew of
system status and act as emergency stow switches.
Each reverser is protected by the following safety features:
- A THRUST REVERSER EMERG STOW switch/light on the glareshield which, when
pressed, bypasses the normal reverser control circuits to initiate stowage
of the reverser.
An automatic stow electrical circuit which initiates stowage of the
reverser after any uncommanded movement of the reverser from the fully
stowed position.
A lock on the flexshaft system which, when engaged, limits the reverser to
0.25 inch of travel from the stowed position in the event of a system
malfunction. During normal operation of the reverser, the lock pin of the
lock is pneumatically withdrawn from the lock assembly to permit reverser
deployment.
A mechanical throttle feedback system which drives the throttle lever to
just below the HIGH IDLE position whenever the reverser is deploying. This
feature prevents the engine from producing more than idle thrust in the
event of an inadvertent thrust reverser deployment.
A mechanical interlock in the throttle quadrant which prevents full reverse
thrust from being selected on the thrust reverse lever unless the throttle
lever is in the LOW IDLE position. Conversely, the interlock prevents
operation of the throttle lever until the reverse lever is returned fully
forward to the stow position.
SECTION 17
Page 27
Mar 01/85
cacnhaadllaeinr ger
OPERATING MANUAL
PSP 606
STOWED POSITION
CASCADES
FAN AIR
FLIPPER
DOOR
TRANSLATING
SLEEVE
JET
EXHAUST
BLOCKER
DOOR
BLOCKER DOOR
LINKAGE
DEPLOYED POSITION
Thrust Reverser Stowed and
Deployed Positions
Figure 13
SECTION 17
Page 28
Feb 12/88
cacnftaaduaeirn qer
OPERATING MANUAL
PSP 606
A lock solenoid which prevents movement of the thrust reverse lever from
the deploy position toward higher reverse thrust settings until the
reverser is fully deployed.
A throttle safety lock consisting of a solenoid and a locking lever on the
centre pedestal just ahead of the throttle quadrant (refer to Figure 6).
The locking lever prevents the associated throttle lever from being moved
beyond the HIGH IDLE position whenever the thrust reverser moves from the
stowed position and a weight-off-wheels condition exists.
A. Operation
Except for test purposes, the thrust reversers are operated together. For
ease of description, the operation of one thrust reverser is described; the
operation of the other is similar.
The left reverser controls are armed prior to deployment by pressing the
left reverse thrust ARMED switch/light on the centre pedestal. When
pressed, the switch/light supplies 28-volt dc power to the contacts of a
thrust reverser weight-on-wheels relay. After landing, reverse thrust is
selected by retarding the throttle lever to LOW IDLE and pulling the thrust
reverse lever to the rear to operate the deploy switch in the throttle
quadrant. Closure of the deploy switch contacts, together with a signal
from the landing gear control unit, energizes the left weight-on-wheels
relay. On aircraft 1072, 1086 and subsequent and aircraft incorporating
Canadair Service Bulletin 600-0334, the weight-on-wheels relay is also
energized if wheel spin-up equivalent to a speed of 65 knots or more is
detected by the anti-skid control unit. Power from the closed contacts of
the weight-on-wheels relay is supplied to the arming and deploy solenoid
valves to initiate thrust reverser deployment as follows:
Bleed air enters the secondary lock actuator and drives the secondary
lock pin to the unlocked position. Movement of the locking pin
directs arming air to the directional valve actuator and moves the
feedback mechanism to the deploy position. Simultaneously, arming air
enters and opens the PDU inlet valve.
As the PDU inlet valve opens, the PDU brake actuator is pressurized,
releasing the PDU air motor brake. The unlocked air motor starts to
rotate in the deploy direction.
The output from the PDU air motor is transmitted through the SPUR
gearbox and the flexshaft system to the three mechanical actuators on
the translating sleeve. The actuators drive the translating sleeve
rearward to the fully deployed position.
SECTION 17
Page 29
Mar 01/85
cmnaaair
chanentjer
OPERATING MANUAL
PSP 606
Q
Z
Ui
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Thrust Reverser System
Figure 14
Schematic SECTION 17
Page 30
Mar 01/85
canatiair
chanenQer
OPERATING MANUAL
PSP 606
REVERSE THRUST SWITCH/LIGHTS
When pressed, amber ARMED light comes on and 28-volt dc
power is supplied to associated arming solenoid.
Amber UNSAFE TO ARM light is energized by associated
WOW and safety relays. Light comes on in the following
—If electrical fault occurs to operate the associated WOW
relay.
—If deploy is selected or deploy switch fault occurs during
flight.
REVERSE THRUST
LEFT RIGHT
UNSAFE
TO ARM
UNSAFE
TO ARM
ARMED
/ L - PUSH TO ARM •
©
THRUST REVERSER EMERG STOW SWITCH/LIGHT
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 stow position. Light is energized
through three sources:
—Stow switch on reverser assembly.
—Air motor brake position switch.
—Unlock switch on secondary lock assembly.
Green REVERSE THRUST light comes on when reverser
reaches fully deployed position and goes out immediately
when reverser moves from deployed position. Light is
energized through deploy switch on reverser assembly.
c== THRUST REVERSER
EMERG STOW
REVERSE
UNLOCKED
REVERSE
THRUST
PUSH
LEFT
REVERSE
UNLOCKED
o
REVERSETHRUST
PUSH
RIGHT
Thrust Reversing Arming and Indicating
Figure 15
SECTION 17
Page 31
Mar 01/85
OPERATING MANUAL
PSP 606
Near the fully deployed position, the deployed microswitch on the
reverser closes. When closed the switch causes the green REVERSE
THRUST light on the glareshield to come on and energizes the throttle
interlock solenoid to permit movement of the thrust reverse lever
beyond the deploy position. Simultaneously, the PDU feedback
mechanism acts to move the PDU directional valve to the closed
position, progressively slowing the air motor. When the reverser is
fully deployed, the brake deploy dump valve opens, resetting the air
motor brake and locking the reverser.
The blocker doors, pivoted into position by the rearward motion of the
translating sleeve, close off the fan exhaust and redirect the fan
thrust forward through the exposed cascade vanes.
Normally 2.0 seconds are required for the reverser to reach the fully
deployed position after the thrust reverse lever is set to the deploy
position.
The reverser is returned to the stowed position by pushing the thrust
reverse lever fully forward to open the deploy switch contacts on the
throttle quadrant. Power remains applied to the arming solenoid valve
through the stowed switch on the reverser assembly but the deploy solenoid
is de-energized shutting off the supply of air to the directional valve
actuator. Spring force returns the directional valve actuator to the stow
position and the brake stow dump valve closes to release the PDU air motor
brake. The PDU air motor rotates in the stow direction driving the thrust
reverser assembly, via the flexshaft system and the translating sleeve
actuators, toward the stowed position. As the reverser approaches the stow
stops, the PDU feedback mechanism closes the PDU directional valve to stop
the air motor. At the same time, the stowed switch on the reverser
assembly opens removing power from the arming solenoid valve. Air pressure
trapped in the air motor and in the air motor brake actuator rapidly
decreases, allowing the brake to reset as the reverser assembly contacts
the stowed stops. The stow cycle is normally completed 4.0 seconds after
the thrust reverse levers are returned to the stow position.
ENGINE VIBRATION MONITORING SYSTEM (Figure 16)
The engine vibration monitoring (EVM) system provides the flight crew with a
continuous indication of the vibration level of each engine. The main
components of the system include single transducers mounted on the rear of the
high pressure compressor casing of each engine, a signal conditioner and an
indicator panel located on the pilot!s side console. The indicator panel
contains a dual-quadrant indicator, which gives a readout of vibration levels,
in inches per second, for each engine and a caution switch/light, which comes
on when a predetermined vibration level on one or both engines is exceeded.
The switch/light is pressed in to test the operation of the dual indicator and
the caution legend.
SECTION 17
Page 32
Mar 01/85
canadair
chauencjer
OPERATING MANUAL
PSP 606
ENGINE VIBRATION INDICATOR
Provides continuous indication of left and right engine v*xaton
levels as sensed by transducers mounted on engine compre
section.
ENGINE
VIBRATION
HIGH
VIB ENGINE VIBRATION CAUT»ON LIGHT
Amber HIGH VIB i^*t
the engines ha* •*/
that vibration level of one or both of
1 2 IN/SEC for more than 3 seconds.
PRESS TO TEST
Pressing HIGH vifi locator twitch provides functional test of
indicator and ft*gna< con&tionm indicator will show L and R reading
of 2.0 IN/SEC and * »*mc*> • prassed for 3 seconds, HIGH VIB light
comes on.
Press-to-test function don nor verify operation of engine mounted
transducers or cab* asaembbes Operation of these components
must be checked by noting indicator reading after engine start.
Engine Vibration Monitor Panel
Figure 16
SECTION 17
Page 33
Mar 01/85
cacnhaaanaeinr qer
OPERATING MANUAL
PSP 606
Each transducer contains a crystalline ceramic element which generates an
electrical signal proportional to the intensity of engine vibration. The
signals generated by the transducers are transmitted through shielded cables to
a signal conditioner, which converts them into analog dc voltages, suitable for
transmission to the EVM indicator.
The signal conditioner contains an alarm circuit which causes the amber HIGH
VIB caution light on the EVM indicator panel to come on if the vibration level
of either engine exceeds 1.2 inches per second for a period greater than
3 seconds. The 3-second delay in the alarm signal to the EVM indicator is
sufficient to prevent spurious warnings caused by high transient engine
vibrations.
The operation of the signal conditioner and the EVM indicator is tested by
pressing the HIGH VIB indicator switch/light on the EVM indicator panel. The
indicator pointers move to show a left and right vibration level of 2.0 inches
per second and, if the indicator switch is depressed for at least 3 seconds,
the amber HIGH VIB caution light comes on. This test function of the system
does not test the operation of the engine-mounted transducers or check the
electrical continuity of the signal transmission cables. The proper operation
of these components is verified by noting the vibration readings after engine
start.
. OVERSPEED PROTECTION (Figure 17)
The overspeed protection system automatically initiates engine shutdown if the
low pressure turbine overspeeds. Two magnetic speed sensors, located between
the low pressure turbine wheels, generate pulsed electrical signals, whose
frequencies are proportional to turbine speed, and transmit them to an
overspeed control unit on the fan casing. If the signal frequency from both of
the sensors exceeds a preset limit, the control unit energizes the solenoid of
an overspeed fuel shutoff valve, to shut down the engine (refer to ENGINE FUEL
SYSTEM). In addition, the control unit causes an overspeed warning light in
the flight compartment to come on, and activates an override relay to disarm
the overspeed protection system of the operating engine.
If desired, the overspeed protection system can be reset after an overspeed
shutdown by pressing the overspeed warning light of the inoperative engine.
Pressing the light causes it to go out and opens the overspeed fuel shutoff
valve. Overspeed protection for both engines is restored after the throttle of
the previously shutdown engine is moved forward from the SHUT OFF position.
Overspeed system electrical circuits are wired through switches on the throttle
quadrant and a weight-on-wheels relay. In general, the overspeed protection
system on one engine becomes inoperative in the air:
If an overspeed shutdown occurs on the opposite engine
If the throttle on the opposite engine is moved to SHUT OFF
SECTION 17
Page 34
Mar 01/85
cacnhaadnaeinr tjer
OPERATING MANUAL
PSP 606
o OVERSPEED
o
©
S\KF
©
^ 1
j
| ENGINE
k
i
RH
I
OVERSPEED
^ > ^v
_^\j
©
©
L—a—^^^
o
OVERSPEED WARNING LIGHTS (RED)
L ENG or R ENG light comes on to indicate overspeed shutdown of
left or right engine. Light remains on until system is reset.
While airborne, the overspeed protection system of one engine is
de-energized:
- Following overspeed shutdown of the opposite engine.
- If the throttle of the opposite engine is returned to the SHUT OFF
position.
During single engine operation on the ground, overspeed protection
system of operating engine remains energized regardless of throttle
setting of inoperative engine.
PRESS TO RESET
Following overspeed shutdown, system is reset by pressing
overspeed warning light.
- Light goes out.
- Overspeed fuel shutoff valve opens.
- Overspeed protection is re-established for both engines when the
throttle of the previously shut down engine is moved from the
SHUT OFF position.
OVERSPEED TEST
Three-position self-locking toggle switch. Setting switch to LH or RH
position allows left or right overspeed control unit to simulate low pressure
turbine overspeed 65% Nl rpm. Proper operation of system is indicated by
engine shutdown and appropriate warning light coming on.
Overspeed Protection System SECTION 17
figure 17 Page 35
Mar 01/85
canaetair
challenger
OPERATING MANUAL
PSP 606
During single engine operation on the ground, the action of the weight-on-wheels
relay ensures that overspeed protection remains available on the operating
engine regardless of the other engine's throttle position. The test switch
with LH and RH test positions tests the operation of the overspeed system.
Setting the switch to one of the test positions lowers the trip reference of
the overspeed control unit on the selected engine. If the Nl rpm of the test
engine is at or above 65%, the engine shuts down in the manner described above.
Electrical power for the overspeed system is provided by the battery bus.
1. ENGINE SYNCHRONIZER SYSTEM (If installed)
The engine synchronizer system allows the flight crew to synchronize the left
and right engine fan (Nl) speeds. The system is designed to compare the fan
speeds of both engines with one another and automatically reduce or increase
the right engine fan speed to match that of the left. The control panel,
ENGINE SYNC, is located on the centre instrument panel and is connected
electrically to a control unit in the underfloor avionics bay. The control
unit monitors Nl signals from each engine to control the operation of an
electrically powered actuator on the right engine. The actuator adjusts the
right engine fuel control unit (FCU) power lever through an operating range of
6 degrees to maintain the required fan speed.
A. Operation (Figure 18)
The control panel consists of two indicator lights, a SYNC INOP light and a
toggle switch with three positions: OFF, SET and ENGAGE. The indicator
lights come on, directing the pilot to advance or retard the right engine
throttle lever. The three toggle switch positions perform the following
functions:
OFF - Electrical power is removed from the system
SET - The system is energized and the right engine actuator is set to
the neutral (rig) position. If the fan speeds are not
synchronized to within 1.0£ of each other, one of the indicator
lights comes on to command an increase or decrease in right Nl
speed. If the indicator light showing an upward pointing arrow
comes on, the right throttle lever must be moved forward to gain
synchronization. If the light with the downward pointing arrow
is on, the right throttle lever must be retarded. Each light
goes out when the right engine Nl speed in within 1% of the left
engine Nl.
ENGAGE - The system is put into automatic operation and the left engine Nl
is used as a reference while the right engine FCU power lever
angle is increased or decreased to obtain engine synchronization.
Automatic synchronization is not possible when one or both of the
engines are operating at an Nl speed below 38£ or if the left
engine is operating at an Nl speed 12.5X above or below that of
the right engine. In either of these cases, the SYNC INOP light
comes on if ENGAGE is selected.
SECTION 17
Page 36
Mar 01/85
cacnhaadiiaeinr qer
OPERATING MANUAL
PSP 606
SYSTEM FAULT WARNING
With ENGAGE selected SYNC INOP light comes on to indicate
any of the following system faults:
- Actuator failure
- Loss of either input signal or if either engine speed drops below 38 % N1
- If engine speed difference is greater than 12.5 % N1
- If system is unable to synchronize fan speeds within 20 seconds
OENGINE SYNC
ENGAGE
|SYNC|
INOP SET
INDICATOR LIGHTS
When system is in SET mode, one of the two lights
comes on directing the pilot to advance or retard
the right engine throttle. Light goes out when N1
speeds are within 1 % of each other.
RIGHT
THROTTLE
MODE SELECTOR SWITCH
Three position toggle switch.
OFF - Removes electrical power from system.
SET - Applies electrical power to system. Returns actuator linkage to
centered position and causes indicator lights to come on
directing pilot to advance or retard right engine throttle lever.
ENGAGE - System automatically synchronizes fan speeds (N1) by
adjusting right engine fuel control unit {FCU) power
lever angle.
Engine Synchronizer System Control Panel SECTION 17
Figure 18 page 37
Mar 01/85
canadair
chauencjer
OPERATING MANUAL
PSP 606
Before the toggle switch is set to OFF, the system should be operated in
the SET mode for at least 5 seconds to allow the system actuator to return
to its centered position.
B. Fault Warning
With ENGAGE selected, the SYNC INOP light on the control panel comes on in
response to the following system faults:
Loss of either Nl signal or if either engine speed drops below 38% N-j
I f engine speed difference is greater than 12.5% N-|
Actuator failure
I f the system is unable to synchronize fan speeds within 20 seconds.
Once the SYNC INOP light has come on, the system can only be re-engaged by
returning the toggle switch to the SET position for at least 5 seconds then
re-selecting ENGAGE.
12. COWLINGS (Figure 19)
The external cowling sections surround the engine, forming a completely
enclosed engine nacelle. In addition, internal cowlings which consist of fan
duct and core cowl panels, encase the engine core. The external cowlings
consist of a nose cowl, access cowl doors, fixed apron panels and the thrust
reverser translating sleeve. The interior cowlings consist of a four-segment
core cowl and a three-segment fan duct.
A. Exterior Cowlings
(1) Nose Cowl
The nose cowl provides the engine fan with a smooth and unrestricted
airflow and contains the cowling anti-icing ducting. The nose cowl is
bolted to the front flange of the engine fan section.
(2) Access Cowl Doors
Two access cowl doors enclose the engine sections between the trailing
edge of the nose cowl and the leading edge of the thrust reverser
translating sleeve. The doors are secured closed by three quickrelease
latches along the outboard split line. For servicing, the
upper and lower access cowl doors can be held in the open position by
support rod assemblies which are fixed to the doors at one end and
attached to engine brackets at the other end.
SECTION 17
Page 38
Mar 01/85
canadlair
chauencjer
OPERATING MANUAL
PSP 606
THRUST REVERSER TRANSLATING SLEEVE
UPPER ACCESS COWL DOOR
NOSE COWL
LOWER ACCESS COWL DOOR
ACCESS COWL DOORS CLOSED
FAN CASING
ACCESS COWL DOORS REMOVED
LOWER FAN DUCT PANEL
PRIMARY EXHAUST NOZZLE
ACCESS COWL DOORS AND FAN DUCT PANELS REMOVED
Cowlings SECTION 17
Figure 19 Page 39
Feb 12/88
OPERATING MANUAL
PSP 606
(3) Thrust Reverser Translating Sleeve
The thrust reverser translating sleeve surrounds the engine primary
exhaust nozzle and, during forward thrust operation, forms the outer
casing of the fan exhaust. During thrust reverser operation the
entire translating sleeve assembly is driven rearward by the thrust
reverser actuators (refer to THRUST REVERSING).
Interior Cowlings
(1) Fan Duct Panels
Three fan duct panels, two of which are removable, and a fixed support
beam form the exterior casing of the fan duct. The inboard fixed
panel inboard cut-outs which allow the transit of pneumatic ducting
from the service pylon to the engine. A fixed support beam houses the
integrated drive generator oil system cooler assembly which protrudes
into the fan duct cavity. The forward edges of the inboard fixed fan
duct panel and the fixed support beam assembly are bolted to an
attachment ring which is secured to the rear of the engine fan
casing. The rear edges are bolted to the forward portion of the
thrust reverser and provide full support for the reverser assembly.
The two removeable fan duct panels are secured to the fan duct casing
attachment ring and the thrust reverser forward flange by quick
release latches. All fan duct split lines are sealed against air
leakage.
(2) Core Cowls
Four core cowls, three of which are removable, encase the engine
between the fan section and the primary exhaust nozzle and form the
inner surface of the fan duct. The fixed core cowl, on the inboard
side of the engine, provides for the transit of lines from the service
pylon. Intake ducts located just to the rear of each cowl leading
edge admit cooling fan air into the engine core area.
SECTION 17
Page 40
Mar 01/85
cacntiaaduaeinr cjer
OPERATING MANUAL
PSP 606
13. ENGINE ANTI-ICING (Figure 20)
The engine incorporates the following anti-icing provisions:
The rotating fan spinner is heated continously by engine lubricating oil.
The splitter ring, which separates the fan and internal airflows, and the
hollow low pressure compressor inlet vanes are heated by bleed air tapped
from the sixth stage of the high pressure compressor.
A. Fan Spinner Anti-Icing
The hot oil anti-icing of the fan spinner forms an integral part of the
engine oil system and operates continuously, without control from the
flight compartment, whenever the engine is running. The flow of oil
through the spinner assembly prevents the build-up of ice under the most
extreme icing conditions.
B. Splitter Ring and Inlet Vane Anti-Icing
The splitter ring and inlet vane anti-icing components include an
electrically operated anti-icing valve and pressure switch, located on the
lower right side of the high pressure compressor casing, and pneumatic
tubing connecting the valve and pressure switch assembly to the engine fan
section. Internal ducts within the fan section route the anti-icing bleed
air to the splitter ring and the hollow inlet vanes. The left and right
systems are controlled by the same switch/lights on the anti-ice control
panel, designated ENGINES, LEFT and RIGHT, which control the engine air
intake anti-icing system (refer to SECTION 14, ICE/RAIN PROTECTION).
The engine anti-icing valve contains a solenoid which, when energized,
holds the valve closed. Pressing the appropriate ENGINES switch/light on
the anti-ice control panel de-energizes the solenoid, causing the valve to
open. This design feature ensures that, in the event of an electrical
power failure, the valve assumes the open position until electrical power
is restored.
14. POWER PLANT DRAIN, VENT AND ECOLOGY SYSTEMS
A. Drains and Vents
The power plant drain and vent system eliminates the accumulation of fluids
and vapours from the nacelle, the engine accessories and the gearbox.
Each drain and vent line is routed from an engine component to a common
drain manifold which vents into a drain mast in the lower access cowl
door. A vent is also provided between the ecology tank and the drain
manifold.
SECTION 17
Page 41
Mar 01/85
canaaatr
chauentjer
OPERATING MANUAL
PSP 606
WING
MAN
ANTI-ICE
OFF l
IR HEAT
ENGINES ,
PUSH ON/OFF
LEFT RIGHT
\\\ FAIL-]
III 0N 1
r?j
M/
[L/i
| FAIL ||
J ON 1
CTI/AUTO 1
IL /
>Gy/
VI.
FAULT
1 'SOL
j OPEN 1
| WSHLD
LOW LOW
HIGH HIGH
|| NOHT | NOHT |
1 TEST TEST
NOHT NOHT j
1 TEST | TEST 1
F
OFF/ K
RESET
FRONT
TEST
®
L-LEFT—BRIGHT -J
/cOSEDJ
ANTI-ICE CONTROL PANEL •
ENGINES SWITCH/LIGHTS
When pressed switch/lights open associated engine anti-ice
shutoff valves. Switch lights remain at held in position until
pressed a second time to close valves.
White ON light comes on whenever associated switch/light
is pressed and remains on until switch/light is pressed a
second time.
Amber FAIL light comes on if loss of bleed air pressure is
detected by pressure switch at associated engine anti-ice
shutoff valve. When switch/lights are pressed, FAIL lights
come on momentarily until pressure at anti-ice shutoff valves
exceeds 10 psi.
LP COMPRESSOR INLET GUIDE VANES
FAN SPINNER
SPLITTER RING AND INLET VANE ANTI-ICING
Bleed air from sixth stage of HP compressor flows through hollow inlet guide
vanes. Bleed air exits through perforations on trailing edge of vanes. Heat is
conducted from guide vanes to splitter ring assembly. Bleed air is controlled
from ANTI-ICE panel in flight compartment.
FAN SPINNER ANTI-ICING
Hot lubricating oil from LP spool reduction gearbox leaves duct at centre of
spool and sprays against nose of spinner assembly. Heat is conducted over
remainder of spinner surface to provide full anti-icing protection. System
operates whenever engine is running.
EFFECTIVITY
jjj Panel on A/C incorporating SB 600-0495.
For panel on other A/C, refer to Section 14.
SPLITTER RING
Engine Anti-Icing SECTION 17
Figure 20 Page 42
Feb 12/88
cacnhaadiiaeinr qer
OPERATING MANUAL
PSP 606
. Ecology System
The ecology system prevents the dumping of fuel overboard during engine
shutdown or aborted starts* Residual fuel from the combustion chamber
drain is collected in an engine-mounted tank. At the next engine run an
ecology ejector pump removes the collected fuel and returns it to the
engine fuel inlet. The ecology tank capacity allows one normal shutdown
and two wet start attempts without dumping fuel.
SECTION 17
Page 43
Mar 01/85
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