Bombardier Challenger 605 - Power Plant

GENERAL

The Challenger 605 is equipped with two General Electric CF34-3B high-bypass ratio turbofan
engines.
The dual-assembly engine consists of a fan rotor (N1) and a compressor rotor (N2). The N1 rotor is
comprised of a single-stage fan connected through a shaft to a four-stage low-pressure turbine. The
N2 rotor is a 14-stage axial flow compressor connected through a shaft to a two-stage high-pressure
turbine. The accessory gearbox is mechanically driven by the N2 compressor.
Normal takeoff thrust rating is 8,729 pounds per engine. During engine-out operation, the automatic
performance reserve (APR) system increases thrust on the operable engine to 9,220 pounds.

FLAT-RATED THRUST

Outside air temperature and pressure altitude are determining factors in achieving takeoff and APR
power. Increases in ambient temperature or pressure altitude adversely affect the engine’s ability to
produce rated thrust. The CF34-3B is flat-rated to ISA + 15°C at sea level.

ENGINE CONSTRUCTION

Description

The CF34 power plant has two independently rotating major assemblies. The N1 section consists
of a fan rotor that is driven through a shaft by a four-stage low-pressure turbine. The N2 section is
comprised of a 14-stage axial flow compressor, a combustor, an accessory gearbox and a
two-stage high-pressure turbine. The compressor is driven by the high-pressure turbine.

Flow Distribution

Engine airflow passes through the single-stage fan, and is divided into two airflow paths:
• Bypass air – Air is accelerated by the single-stage N1 fan only, and is ducted around the
engine nacelle. Bypass airflow produces approximately 80% of the thrust at takeoff.
Thrust reversers are used to divert the bypass air forward to assist in airplane braking on
the ground.
• Core air – Air that is accelerated by the N1 fan enters the N2 core where it is
compressed, mixed with fuel, and ignited. The resulting combustion gases are
exhausted through the high-pressure two-stage N2 turbine, which drives the N2
assembly. The exhaust gases are then discharged through the low-pressure four-stage
N1 turbine to drive the N1 fan. Jet pipe thrust produces approximately 20% of the takeoff
thrust.

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ENGINE CONSTRUCTION (CONT'D)

Flow Distribution
Figure 19−10−1

Major Power Plant Components

N1 Fan

The N1 fan is a single-stage fan that consists of 28 titanium blades. A Kevlar blanket is
wrapped around the inlet housing to contain damage from a failed fan blade. The N1 fan rpm is
displayed on the EICAS page.

Variable Geometry (VG) Compressor

The VG system regulates airflow across the compressor, by changing the position of the
compressor inlet guide vanes and the first five stages of the stator vanes. Fuel metered by the
fuel control unit (FCU) is used to hydraulically change the vane angle. The VG system
optimizes the airflow’s angle of attack at the compressor blades, and provides compressor stall
and surge protection.

Accessory Gearbox

The engine-mounted accessory gearbox is driven by the N2 compressor. The gearbox drives
the:
• Engine lubrication pumps;
• Alternator that powers the N1 control amplifier;
• Engine-driven hydraulic pump (hydraulic pump 1A or 2A);
• Engine-driven fuel pump; and
• Integral drive AC generator (IDG).
Mounted on the gearbox is the air turbine starter (ATS).
The N2 rpm is displayed on the EICAS page.

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ENGINE FUEL SYSTEM

Description

Fuel is delivered to the fuel injectors at pressures and flow rates required to maintain the desired
engine thrust. The engine fuel system, in addition to providing fuel for combustion, is used for:
• Controlling and actuating the VG compressor linkage;
• Cooling the engine oil (heat exchange);
• Actuating and lubricating servos within the fuel control unit (FCU); and
• Providing motive flow for the main ejector and scavenge ejector pumps.
Combustion fuel can be interrupted by moving the thrust lever to SHUT OFF, or by selecting the
engine FIRE PUSH switch/light. The SHUT OFF position shuts off the fuel at the FCU. The FIRE
PUSH switch/light closes the fuel shutoff valve.

Components and Operation

Engine-Driven Fuel Pump

The accessory gearbox-mounted fuel pump is comprised of three separate pumps contained
within a single housing. The engine-driven fuel pump provides high-pressure fuel at a flow rate
that exceeds the requirements of the engine at any power setting. Fuel pump pressure is used
to generate motive flow for the scavenge and main ejectors of the aircraft fuel system.

Fuel/Oil Heat Exchanger and Fuel Filter

A fuel/oil heat exchanger is used to warm engine fuel and cool engine oil. The fuel temperature
at the fuel filter is indicated on the SUMMARY page. A fuel filter is used to remove solid
contaminants from the fuel. If the filter becomes clogged, the fuel bypasses the filter to ensure
continued operation. A clogged filter is indicated by the L (R) FUEL FILTER caution EICAS
message.

Fuel Control Unit (FCU)

The fuel control unit is a hydromechanical metering device that supplies fuel in response to
mechanical inputs from the thrust levers. In addition, the FCU controls and actuates the VG inlet
guide vanes and stator vanes of the engine compressor.
During start and at low power, the FCU hydromechanically schedules the fuel.

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ENGINE FUEL SYSTEM (CONT'D)

Engine Fuel System − Schematic

Figure 19−10−2

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ENGINE FUEL SYSTEM (CONT'D)

Electronic Control Unit (ECU)

When at high thrust settings, the engine is controlled by an electronic control unit (ECU), which
works in tandem with the FCU. The ECU is also referred to as the N1 control amplifier. The ECU
(or N1 control amplifier) trims the FCU fuel output to maintain a N1 speed schedule for a given
thrust lever setting.

Fuel System Operation

N2 Speed Control (N1 <79%)

At low power settings (N1 below 79%), the FCU hydromechanically controls the N2 speed. In
N2 mode, the FCU adjusts N2 speed, so that matched movement of the thrust levers produces
nearly matched N2 rpm for the engines. N1 speeds, fuel flows, or ITT indications may differ
between engines.

N1 Speed Control (N1 >79%)

At takeoff, climb and cruise power settings (N1 above 79%), the N1 control amplifier controls
the engine N1 rpm. The amplifier trims the FCU fuel output to achieve the desired N1 rpm.
With the ENG SPEED switches selected to ON, automatic switchover from N2 to N1 speed
governing occurs at 79% N1. Matched movement of the thrust levers produces nearly matched
N1 rpm and nearly matched thrust between the engines.

NOTE
If an ENGINE SPEED switch is moved from ON to OFF at high
power settings, the engine will revert to N2 speed control. A rapid
increase in engine acceleration will occur, and an overtemperature
limit may be exceeded.

ENGINE CONTROL Panel

Figure 19−10−3

AUTOMATIC PERFORMANCE RESERVE SYSTEM

Description

During takeoff, the APR system monitors the N1 rpm of both engines. If a significant loss of N1 rpm
is sensed on one engine, the APR system automatically increases the thrust of the remaining
engine to the APR thrust rating.

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AUTOMATIC PERFORMANCE RESERVE SYSTEM (CONT'D)

APR activation does not override thrust lever input to the FCU, nor does it restrict movement of the
thrust lever.

Operation

The APR system is armed during takeoff when the APR switch is selected to ARM, and both
engines’ N1 rpm is above 79% (N1 speed mode). This is indicated by the APR ARM advisory
EICAS message. During a normal takeoff, the advisory message is removed five minutes after
APR arming.
During takeoff, the APR system monitors the N1 rpm of both engines through the DCUs. If an N1
rpm drop below 67.6% rpm is sensed at either engine, the system automatically commands both
engines to increase N1 speed. Only the normally operating engine can respond, which it does by
increasing the N1 fan speed by a minimum of 2%. This increase in rpm equates to an increase of
approximately 500 pounds of thrust. No roll-back in N1 rpm occurs when APR thrust has been
commanded and the APR system times-out.
When the APR system is activated, a green APR icon appears in the center of the N1 gauge of the
operating engine.

APR Test

The APR TEST switches allow the system to be tested on the ground. The system can only be
tested successfully with the engines operating. Normal system test operation results in the APR
TEST 1 (2) OK advisory EICAS message. If the system fails, an APR INOP caution EICAS
message is displayed.

ENGINE CONTROL Panel

Figure 19−10−4

ENGINE OIL SYSTEM

Description

Oil from each engine nacelle tank is circulated under pressure to lubricate the engine and
accessory gearbox.

Components and Operation

The gearbox-driven main lubrication pump pressurizes the lubrication system. Oil flows from the
pump through an oil filter, a fuel/oil heat exchanger, and continues through the engine sumps to
the bearings and gearbox.

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ENGINE OIL SYSTEM (CONT'D)

The oil is returned to the oil tank by the gearbox-driven main lubrication and scavenge pumps. The
oil passes through scavenge screens for filtering prior to the oil pumps. The oil then flows through
a chip detector and a deaerator to the tank. Maximum oil consumption is 6.4 ounces or 0.05 US
gallons/hour.
Engine oil pressure and temperature indications are presented on the EICAS page. To provide
system redundancy, a pressure switch and separate pressure transmitter are used to monitor the
engine oil pressure. When low oil pressure is detected by the pressure switch, the L (R) ENG OIL
PRESS warning EICAS message is presented. If the pressure transmitter detects low oil pressure,
the EICAS digital oil pressure readout changes to red.
Chip detector and impending oil filter bypass indications are provided in the aft equipment bay on
junction box 5 (JB5), but are not presented on EICAS.
Pressurized refilling of the engine oil tanks is provided by a replenishment tank system, located in
the aft equipment bay. The system remotely gauges engine oil tank level, and is used to transfer
oil to the engine-mounted tanks.

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ENGINE OIL SYSTEM (CONT'D)

Oil System Schematic

Figure 19−10−5

Engine Oil Replenishment System

The Challenger 605 is equipped with an oil replenishment system, located in the aft equipment
bay. The system consists of a replenishment tank, holding approximately 6.0 U.S. quarts of oil,
a pump, selector valve, and control panel.

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ENGINE OIL SYSTEM (CONT'D)

If a low oil tank quantity is indicated during the test, the nacelle oil tanks can be filled from the
replenishment tank. Pump and selector valves, located next to the replenishment tank, are used
to transfer oil from the replenishment tank to the applicable engine-mounted tank.
Engine Oil Replenishment Procedure

Oil levels should be checked between 15 minutes to two hours after engine shutdown. The
engines must be motored if the replenishment period is exceeded. Maximum refill allowable is
2 U.S. quarts, then the engine must be dry-motored for at least 30 seconds prior to adding
more oil.
Oil replenishment is accomplished as follows:
1. Note the oil quantity on the oil replenishment tank gauge.
2. The system power switch is selected ON, illuminating the green ON light.
3. The ENG OIL PRESS TO TEST switch/light is activated to illuminate (test) the green
LH (RH)-FULL indications.
4. The selector valve is rotated to the L or R position, as required, to pump oil from the oil
replenishment tank to the associated engine oil tank.
5. The selector valve is released when the pump automatically shuts off, and the
LH (RH)-FULL legend illuminates. A full level is indicated by the illumination of the
respective side green light (LH FULL or RH FULL).
6. The system power switch is selected OFF.
7. Oil quantity used for each engine is noted.

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ENGINE OIL SYSTEM (CONT'D)

Oil Replenishment System

Figure 19−10−6

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ENGINE BLEED AIR SYSTEM

Description

Engine bleed air is extracted from the 7th, 10th and 14th stages of the engine compressor, and
used by the airplane systems below:
• 7th stage: Pressurization of oil seals, and the venting of engine sumps in the lubrication
system.
• 10th stage: Pneumatic supply to the air conditioning and pressurization systems, and
engine starting.
• 14th stage: Pneumatic supply to the engine cowl and wing anti-ice systems, or thrust
reverser systems.

Components and Operation

Ducting and check valves are used to direct the flow of bleed air from the engine to the 10th- and
14th-stage bleed air manifolds. The nacelle and pylon ducts are monitored for bleed air leakage by
the engine fire and jet pipe overheat detection systems.
For additional information, refer to Chapter 9, Fire Protection, and Chapter 18, Pneumatic System.

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ENGINE BLEED AIR SYSTEM (CONT'D)

Engine Bleed Air System − Schematic

Figure 19−10−7
ENGINE STARTING SYSTEM

Description

The starting system consists of the starter control valve and the air turbine starter. The air turbine
starter drives the accessory gearbox, which in turn drives the N2 core section. DC electrical power
and air from the 10th-stage bleed air manifold are required to open the starter control valve and
engage the air turbine starter.

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ENGINE STARTING SYSTEM (CONT'D)

The 10th-stage bleed air manifold can pressurized by the:

• APU;
• External air cart; or
• 10th-stage bleed air from the operating engine (crossbleed).
The following minimum bleed air pressures should be available for an engine start:
• APU – 40 psi;
• External air start cart – 45 psi; or
• Engine crossbleed – 60 psi (approximately 85% N2).
L (R) 10th-stage bleed air manifold pressure is displayed on the SUMMARY page.

CAUTION
During an engine crossbleed air start, ensure that intake and exhaust
areas of the operating engine are secure. When starting engines in
close quarters, consideration should be given to the effects of jet
blast.

Starting System − Schematic

Figure 19−10−8

Components and Operation

Starter Control Valve

The starter control valve is mounted next to the air turbine starter, inside the engine nacelle. It
controls the initial rate of engine acceleration upon engagement of the starter, by regulating the
amount of air supplied to the starter. The starter control valve is controlled by the START and
STOP switches on the START/IGNITION panel.

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ENGINE STARTING SYSTEM (CONT'D)

Air Turbine Starter

The air turbine starter is mounted on the accessory gearbox. The air turbine starter converts
pneumatic energy into mechanical motion. The starter mechanically engages the accessory
gearbox through a clutch, and accelerates the N2 section of the power plant. The starter is
capable of dry-motoring the engine up to approximately 28% N2 rpm. During a normal start, the
starter remains engaged until 55% N2, to assist the engine in accelerating to idle speed. At 55%
N2 rpm, the starter control valve closes and the air turbine starter disengages.
For subsequent starts or relights, the starter clutch does not require that engine rotation be
completely stopped before engaging the starter. The air turbine starter may be engaged at any
rpm up to 55% N2 rpm (starter cutout speed).
Starter disengagement may be commanded at any time by pressing the engine STOP
switch/light.

Engine Start/Ignition Panel

Figure 19−10−9

Start Sequence

Each engine has a set of START and STOP switch/lights on the ENGINE START/IGNITION
panel. When the engine START switch/light is pressed, the following occurs:
• Left, right and isolation 10th-stage bleed air SOVs open;
• Starter control valve on the associated engine opens to allow pressure from the
10th-stage manifold to engage the air turbine starter;
• When the starter control valve opens, a white light illuminates in the associated START
switch/light; and
• At 55% N2 rpm, the start control valve is de-energized and the air turbine starter
disengages.
The air turbine starter is subject to the following starter engagement limits:
STARTER ENGAGEMENT LIMITATIONS

START # ENGAGEMENT TIME COOLING PERIOD

1 30 seconds none
2 30 seconds none
3 & subsequent 30 seconds 5 minutes

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ENGINE STARTING SYSTEM (CONT'D)

DRY MOTORING # ENGAGEMENT TIME COOLING PERIOD

1 90 seconds 5 minutes
2 & subsequent 30 seconds 5 minutes
Dry-Motoring

Dry-motoring is performed with ignition off and thrust levers at SHUT OFF.
Dry-motoring may be used for engine ground starts and engine airstarts.

IGNITION SYSTEM

Description

Two independent alternating current (AC) ignition systems are provided for each engine. Each
ignition system consists of one ignition exciter and one igniter plug.
Ignition system A is powered by the AC essential bus. Ignition system B is powered by the battery
bus through a static inverter.

Operation

There are three ignition system modes:

Normal – Prior to engine start, the ignition is armed by selecting either A or B ignition switch/light
at the START/IGNITION panel. When an ignition system is armed, the corresponding white ARM
light illuminates. Ignition is energized upon starter engagement, and is de-energized at starter
cutout (55% N2). While an ignition system is energized, the corresponding green ON light
illuminates, and the IGNITION A (B) advisory EICAS message appears.
Continuous ignition – Continuous ignition may be selected by pressing the CONT ignition
switch/light on the START/IGNITION panel. Both ignitions A and B are energized. When
continuous ignition is selected, both green ON lights illuminate, and the IGNITION A/B advisory
EICAS message appears.
Continuous ignition must be used during the following flight conditions:
• Takeoff and landings on contaminated runways;
• Takeoff with high crosswind components (greater than 10 knots);
• Flight through moderate or heavier intensity rain;
• Flight through moderate or heavier intensity turbulence; or
• Flight in the vicinity of thunderstorms.
Autoignition (aerodynamic stall protection) – Automatic activation of continuous ignition is
based upon angle-of-attack (AOA) data. Both systems A and B are energized by the stall warning
computer, and remain on until the airplane flight attitude is corrected.

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IGNITION SYSTEM (CONT'D)

CAUTION
Activation of the Stall Protection System Test will activate continuous
ignition.
Both ignition systems A and B are disabled on the associated engine when the L (R) ENGINE
FIRE PUSH switch/light is selected. If continuous ignition was in use, the white ON legend of the
CONT switch/light will extinguish, but the green ON lights in the switch/lights remain illuminated,
and the IGNITION A/B advisory EICAS message remains displayed (to advise the crew that
continuous ignition is still active on the unaffected engine).

Ignition System
Figure 19−10−10

VIBRATION MONITORING SYSTEM

Description

The power plant consists of two major rotating assemblies, the N1 fan and N2 core sections. Each
assembly is continuously monitored for vibration. Indications are displayed on the EICAS page.

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VIBRATION MONITORING SYSTEM (CONT'D)

Operation

N1 Fan

N1 fan vibration is displayed as a numeric readout on the EICAS page as FAN VIB. When the
N1 vibration level is 2.7 mils or greater, the color of the readout changes to amber. There is no
associated caution EICAS message.

N2 Core

N2 core vibration levels are continuously monitored, but are presented only when vibration
levels exceed a target value. An amber VIB icon appears in the middle of the N2 gauge when
the vibration target value is exceeded. There is no associated caution EICAS message.

Vibration Monitoring System Test

The system is tested by selecting the VIB switch on the ENGINE CONTROL panel to the TEST
position. In the TEST position, high vibration levels are simulated in the electrical circuitry. The
following indications appear on selection of the ENGINE VIB test switch:
• FAN VIB readouts increase to 3.6 mils, changing from green to amber, passing through
2.7 mils; and
• Amber VIB icons appear on the N2 dials.

ENGINE CONTROL Panel

Figure 19−10−11

THRUST LEVERS

Description

The thrust lever quadrant contains the thrust levers, thrust reverse levers, microswitches, and
internal locks and stops necessary to control the engines in forward and reverse thrust.

Operation

Thrust Levers

Most functions of the thrust levers are conventional in operation. Thrust lever quadrant settings
are SHUT OFF, IDLE, and MAX POWER.

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THRUST LEVERS (CONT'D)

Mechanical Stop

A mechanical stop prevents the thrust levers from inadvertent movement. When the thrust
lever is at SHUT OFF, the thrust lever is mechanically locked in that position. When the thrust
lever is at idle, the mechanical stop prevents the thrust lever from being accidentally moved to
SHUT OFF. The mechanical stop is released by lifting the idle/shutoff release latch on the
thrust lever.

Friction Knobs

A friction knob is set to a preset friction value for ATS operation, and is not adjustable by the
pilot.

Thrust Lever Position Measurement

Actual thrust lever positions are electrically measured by rotary variable differential
transformers (RVDTs), or sensed by microswitches that are housed within the thrust lever
quadrant. The information is provided to the N1 control amplifier, the flight control computers
(FCCs), and the data concentrator units (DCUs). Other aircraft systems receive thrust lever
position information, including:
• Landing gear warning system;
• Takeoff configuration warning system;
• Cabin pressurization; and
• Ground spoilers.

Takeoff/Go-Around (TOGA) Switches

A takeoff/go-around switch is mounted on each thrust lever. When pressed, the TOGA switch
signals the flight control computers to activate flight director modes accordingly.

Autothrottle System (ATS) Disconnect Switches

These switches, on the forward face of each thrust lever (ATS DISC), are associated with the
autothrottle system (ATS). In addition to control through the ATS control panel on the
glareshield, ATS disengage is also provided by these disconnect pushbutton switches.

Thrust Reverse Levers

The thrust reverse levers control the operation of the thrust reverser system. See the thrust
reverser description in this chapter for further details.

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THRUST LEVERS (CONT'D)

Thrust Levers
Figure 19−10−12

THRUST REVERSER SYSTEM

Description

The thrust reverser system is used to assist in stopping the aircraft on landing and during a
rejected takeoff (RTO). The system is operable on the ground only.

Operation

The thrust reversers are armed when the appropriate switches on the THRUST REVERSER panel
are selected to ARM, and the respective 14th-stage bleed air shutoff valves are opened. When
armed, a L (R) REV ARMED advisory EICAS message is displayed.

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THRUST REVERSER SYSTEM (CONT'D)

Reverse thrust is generated by blocking fan bypass airflow, and redirecting it forward through a
series of cascade vanes. Bleed air, from the 14th stage of the compressor, pneumatically activates
a power drive unit (PDU), which mechanically moves the engine translating cowls rearward by
means of a flexible driveshaft and ballscrew actuators. As the translating cowl moves rearward,
blocker doors rotate to redirect fan airflow forward through the cascade vanes.
Reverser deployment is accomplished by squeezing the thrust reverser triggers and applying
upward pressure on the thrust reverse levers. Thrust reverse lever movement is initially restricted
to approximately 20 degrees by a solenoid stop and a reverse thrust lever lock. This locks the
thrust lever in the IDLE position, and prevents thrust from being applied. When the translating cowl
reaches full aft travel, the solenoid stop is released, allowing the reverse lever to be operated
though its full range, and reverse thrust to be applied.
During normal thrust reverser deployment, an amber REV icon appears in the engine N1 gauge
while the reverser is in transit. When the reverser is fully deployed, the REV icon changes to
green. Reverser deployment is achieved in approximately 5 seconds.

Anti-Ice Disable

On touchdown, or during rejected takeoff with the wing and/or cowl anti-ice system on, the
anti-ice systems are automatically disabled while the thrust reversers are activated. This
disabling action redirects all 14th-stage bleed air to the thrust reverser PDU to ensure proper
operation.

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THRUST REVERSER SYSTEM (CONT'D)

Thrust Reverser Operation

Figure 19−10−13

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THRUST REVERSER SYSTEM (CONT'D)

GND SPOILERS/THRUST REVERSER Panel

Figure 19−10−14
The thrust reverse system incorporates the following safety features:
• In-flight protection – The reversers are locked out in flight. WOW (weight-on-wheels)
signals from the proximity sensing electronics unit (PSEU) or wheel spin-up (from the
anti-skid control unit) must be sensed in order to energize the thrust reverser deploy
relays.
• Flexshaft lockpin and lock cam – The reverse thrust PDU flexshaft is equipped with a
lock pin and lock cam, as well as a brake, which act to prevent uncommanded travel of
the reverser by more than 1/4 of an inch from the fully stowed position.
• Auto stow – In the event of flexshaft lock cam failure, resulting in uncommanded
movement of the thrust reverser away from the fully stowed position, the stow solenoid
will energize, and provided 14th-stage bleed air is available, the thrust reverser will be
returned to the fully stowed position.
• Automatic thrust lever retard system – An auto-retarding thrust lever mechanism
ensures that the thrust lever is at idle whenever the thrust reverser is in transit. In flight,
should a thrust reverser inadvertently deploy, the affected thrust lever is automatically
retarded to IDLE, and locked to minimize asymmetric thrust.
• Emergency stow – A thrust reverser emergency stow system is installed to stow an
unlocked thrust reverser. Selecting the respective UNLK switch/light on the THRUST
REVERSER panel directs 14th-stage bleed air to the PDU to stow the unlocked
reverser.

NOTE
While the UNLK switch/light is selected, 14th-stage bleed air is
continuously applied to the PDU, even if the reverser returns to the
stowed position.

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CONTROLS AND INDICATORS

The CF 34–3B engine controls consist of the following:

• Thrust levers
• Thrust reversers
• Engine start/ignition panel
• Engine control panel
The EICAS page and SUMMARY page provide analog and digital engine information, and system
warning/caution and advisory messages.

Engine Starting

Engine Start/Ignition Panel

Figure 19−10−15

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CONTROLS AND INDICATORS (CONT'D)

Engine Speed Controls

ENGINE CONTROL Panel

Figure 19−10−16

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CONTROLS AND INDICATORS (CONT'D)

Thrust Levers

Thrust Levers
Figure 19−10−17

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CONTROLS AND INDICATORS (CONT'D)

Reverse Thrust

GND SPOILERS/THRUST REVERSER Panel

Figure 19−10−18

Oil System

Impending Bypass and Chip Detector Annunciators (JB5)

Figure 19−10−19

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CONTROLS AND INDICATORS (CONT'D)

EICAS Page

N1 Indications
Figure 19−10−20

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CONTROLS AND INDICATORS (CONT'D)

ITT Indications
Figure 19−10−21

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CONTROLS AND INDICATORS (CONT'D)

N2 Indications
Figure 19−10−22

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CONTROLS AND INDICATORS (CONT'D)

Engine Secondary Indications

Figure 19−10−23

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EICAS MESSAGES
Message Meaning Aural Warning (If Any)
Left or right engine N1 or N2 or ITT above exceedance WARNING
ENGINE
level. (Triple Chime)
L ENG OIL PRESS
Oil pressure in the respective engine is less than 25 psi. “ENGINE OIL”
R ENG OIL PRESS
L ENG START SOV The respective start shutoff valve is not closed with that WARNING
R ENG START SOV engine running. (Triple Chime)

Either the APR has failed or it is not armed (APR to arm and N1 >79% and both eng
APR INOP
speed switches on).
APR CMD SET APR has been inadvertently activated on both engines.
L ENG MISCOMP The engine comparator has sensed a discrepancy in the respective engine
R ENG MISCOMP parameters (N1, ITT or N2).

L FUEL FILTER
Respective fuel filter is in impending bypass mode.
R FUEL FILTER
L FUEL LO PRESS
Insufficient pressure at the respective engine pump inlet.
R FUEL LO PRESS
L REV UNLOCKED
Reverser moved away from fully stowed position without a deploy command.
R REV UNLOCKED
L REV UNSAFE
Unarmed thrust reverser received a deploy command.
R REV UNSAFE
APR ARM Aircraft in takeoff configuration, dynamic test OK.
APR TEST 1 OK
The respective APR circuit has passed the self-test.
APR TEST 2 OK
L ENG SOV CLSD Fuel shutoff valve is closed after activation of the respective ENG FIRE PUSH
switch/light.
R ENG SOV CLSD
IGNITION A ’A’ igniters are activated, ’B’ igniters off.
IGNITION A/B Both ’A’ and B’ igniters are activated.
IGNITION B ’B’ igniters are activated, ’A’ igniters off.
L REV ARMED
The respective thrust reverser is armed.
R REV ARMED
DCU 1 APR FAIL
The respective DCU has failed its APR test.
DCU 2 APR FAIL
L ENG ECU FAIL
The respective engine ECU failed to respond to an APR test.
R ENG ECU FAIL

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 Bombardier Challenger 605 - Power Plant

POWER SUPPLY AND CIRCUIT BREAKER SUMMARY

CB
SYSTEM SUB−SYSTEM CB NAME BUS BAR CB LOCATION NOTES
PANEL

Power Plant Ignition System ENG IGN A AC ESS 3 C7

ENG IGN A DC BATT 1 L7
ENG IGN B DC BATT 1 L8

ENG IGN B BATT BUS 5 A1

Starter System ENG START L DC BATT 1 L10
ENG START R DC BATT 1 L9

Oil Pressure L OIL PRESS

DC BATT 1 L6
Indications TRANS

R OIL PRESS
DC ESS 4 B6
TRANS

OIL BYPASS IND DC APU

BATT 5 B3
DIRECT

ENG OIL POWER DC APU

BATT 5 B4
DIRECT
Vibration Monitor ENG VIB MON AC BUS 1 1 C7

Thrust Reversers THRUST REV

DC ESS 4 A4
AUTO STOW 1

THRUST REV
DC ESS 4 A5
AUTO STOW 2

THRUST REV 1 DC ESS 4 A6

THRUST REV 2 DC ESS 4 A7

Automatic APR
Performance DC BATT 5 A2
Reserve

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