Gulfstream V User manual

FLIGHT CONTROLS
2A-27-10: General:
The flight controls system for the Gulfstream V is hydraulically powered, providing boost
to mechanical linkages to overcome aerodynamic forces associated with high speed
flight. Tandem type hydraulic actuators, receiving hydraulic fluid under pressure from two
independent systems (Left and Right hydraulic systems), are used to move the flight
control surfaces. Both hydraulic systems maintain a system pressure of 3000 psi. Loss of
system pressure by one hydraulic system has no effect on operation of the flight controls,
as the remaining system is capable of maintaining actuator load capacity. In the event of
total loss of hydraulic pressure in both hydraulic systems, the primary flight controls
revert to manual operation.
The flight controls are divided by function as follows (see Figure 2):
•Primary Flight Controls:
The primary flight controls consist of the ailerons, elevators and rudder.
•Secondary Flight Controls:
Secondary flight controls consist of flaps and spoilers.
•Trim Controls:
Trim controls consist of an aileron trim tab, horizontal stabilizer, two elevator trim
tabs and a trimmable rudder. The aileron and elevator trim tabs are mass
balanced to prevent control flutter.
The flight controls system is divided into the following subsystems:
•2A-27-20: Pitch Flight Control System
•2A-27-30: Yaw Flight Control System
•2A-27-40: Roll Flight Control System
•2A-27-50: Horizontal Stabilizer System
•2A-27-60: Flaps System
•2A-27-70: Spoiler System
•2A-27-80: Gust Lock System
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Flight Controls System:
Simplified Fluid Power
Diagram
Figure 1
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2A-27-20: Pitch Flight Controls System:
1. General Description:
The elevators are manually and electrically controlled, mechanically actuated and
hydraulically boosted. Elevator travel ranges from 24 degrees trailing edge up to
13 degrees trailing edge down. Elevator movement is visually depicted on the
FLIGHT CONTROLS synoptic page.
An elevator trim tab is installed on each elevator. Pitch trim position is displayed
on the FLIGHT CONTROLS synoptic page.
Aircraft pitch is accomplished by movement of the control columns which transmit
motion through conventional mechanical linkage (cables, cranks and pushrods) to
displace the elevator. Hydraulic boost to assist inputs from the control columns is
provided by a dual hydraulic boost actuator. The dual hydraulic actuator contains
two pistons; one piston is powered by the Left Hydraulic System (L SYS), and the
other by the Right Hydraulic System (R SYS). During normal flight conditions, the
elevator actuator is powered simultaneously by L SYS and R SYS at 3000 psi
each. The actuators feature automatic hardover prevention. If a hardover occurs
for over 0.5 second, hydraulic pressure to the actuator will be shut off.
There are two complete sets of elevator controls. Normally, elevator controls are
connected to each other and this duplication is transparent to the flight crew. In
Flight Controls System Components
Figure 2
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the unlikely event of an immovable elevator, the elevator controls can be
separated by pulling an elevator disconnect handle, located on the cockpit center
pedestal. Once separated, the operable elevator system is identified and used to
fly the aircraft.
Two stall barrier systems are incorporated into the elevator control system to
prevent a stall by forcing the control columns forward when the crew fails to
respond either to visual indications or to stick shaker vibrations which precede an
impending stall. The stall barrier system is normally left on, but if it is
malfunctioning, the crew can turn the system off, using the STALL BARR switch,
located on the cockpit center pedestal. When a high angle of attack is attained, a
shaker trip point detector activates the control column shaker motors. When a
more severe angle of attack is attained, a pusher trip detector activates its
respective stall barrier pusher.
NOTE:
The control column force can be manually overcome
by pilot or copilot.
2. Description of Subsystems, Units and Components:
A. Automatic Hardover Prevention System:
The elevator actuator incorporates a hardover prevention system which
compares inputs and outputs. If inputs to, and outputs from, the actuator do
not agree, hydraulic pressure to the affected side of the actuator is shut off
and a message is prompted for display on the Crew Alerting System (CAS)
display. The elevators are still operative, but without benefit of the affected
side’s hydraulic boost. The hardover prevention system receives power
from the Left and Right Essential DC bus.
B. Elevator Control Separation System:
(See Figure 4.)
The elevator control systems are dual and separable. In the unlikely event
that a single elevator control system were to become jammed, an elevator
disconnect system provides the means to separate the left and right
elevators from each other. This is accomplished through the use of an
elevator disconnect system, located on the left side of the cockpit center
pedestal and labeled ELEV DISC on its protective cover.
The elevator disconnect system consists of the protective cover, the
elevator disconnect handle (labeled ELEV DISC) and a power assist trigger
(labeled LIFT). If a jammed elevator is detected, the protective cover is
raised and the ELEV DISC handle is pulled. With the handle fully extended,
the pilot’s and copilot’s control columns are separated in the cockpit. The
pilot controls the left elevator and the copilot controls the right elevator. The
immovable elevator side will remain immovable and the movable side is
now free to be used to control the aircraft. The Stall Barrier system remains
functional in this configuration.
If the red power assist trigger was NOT used to actuate the ELEV DISC
handle, the elevator disconnect system can be reset at the discretion of the
flight crew.
If the force required to pull the ELEV DISC handle is too great, the red
power assist trigger, located under the ELEV DISC handle, is pulled. This
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actuates a gas-spring cartridge, providing between 110 to 150 pounds of
controlled force to fully extend the handle and separate the elevator control
systems.
NOTE:
If the red power assist trigger was used to actuate the
ELEV DISC handle, the system cannot be reset
without the use of special tools.
C. Pitch Trim System:
(See Figure 4 through Figure 6.)
(1) Elevator Trim Tabs:
A trim tab is installed on the trailing edge of each elevator. They are
mechanically operated through trim actuators located in each
elevator. The trim actuators can be operated manually or electrically
as described below. Elevator trim tab travel ranges from 22 degrees
trailing edge down to 8 degrees trailing edge up.
(2) Elevator Trim Tab Actuator Heater:
The elevator trim tab actuator heater generates heat to prevent the
actuators from freezing. These heaters generate heat until the
temperature rises to a threshold temperature, predetermined by the
manufacturer. The heaters are controlled by the L/R ELEV TRIM
HEAT circuit breakers, located on the Right Electronic Equipment
Rack (REER), and receives power from the Right Main AC bus.
(3) Manual Trim Control:
Manual control of pitch trim is accomplished by an interconnected
manual trim control wheel set. A control wheel is provided on each
side of the cockpit center pedestal. With electric pitch trim
disengaged, moving either manual trim control wheel adjusts pitch
trim to the desired setting. With electric pitch trim engaged, the
manual trim control wheels move automatically corresponding to the
amount of electric pitch trim movement.
(4) Electric Pitch Trim:
Located on the pilot’s flight panel, the PITCH TRIM ENG/DISENG
switch engages or disengages the electric pitch trim. Pitch trim is
automatically engaged when the autopilot is engaged. However,
pitch trim is not automatically disengaged when the autopilot is
disengaged. To disengage the pitch trim, the ENG/DISENG switch
must be depressed, causing the DISENG legend to illuminate in the
switch. Pitch trim can be engaged even if the autopilot is off by
depressing the ENG/DISENG switch. Should a failure occur
rendering the electric pitch trim inoperative, or pitch trim become
disengaged, messages are prompted for display on CAS. The type
of messages depend upon aircraft airspeed.
With electric pitch trim engaged, pitch trim can be adjusted through
use of a split-half pitch trim switch installed on each control wheel,
labeled NOSE DOWN / NOSE UP. Inadvertent actuation of pitch
trim, including runaway, is minimized through the split-half switch
design. In order for the pitch trim to be actuated, both halves of the
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switch must be simultaneously moved in the same direction. When
electric pitch trim reaches the nose up or nose down limit of
allowable travel, messages are prompted for display on CAS. The
type of messages depend upon whether or not the autopilot is
engaged.
D. Stall Barrier System:
(See Figure 3, Figure 4 and Figure 6.)
The stall barrier system consists of stick shaker motors located on each
control column, a stick pusher actuator located in the tail compartment and
Angle of Attack (AOA) probes located on each side of the fuselage forward
section.
The stall barrier system can be deactivated by depressing the autopilot
disconnect switches (A/P DISC / BARR DISC) on either control wheel, or
by selecting the STALL BARR switch, located on the center console, to the
OFF position. When the stall barrier system is selected off, a message is
prompted for display on CAS.
NOTE:
The autopilot servo input to the elevator control
system is through a separate set of cables connected
to the elevator actuator. Servo inputs are introduced at
the actuator. The force generated by the stall barrier
system will overcome autopilot force.
(1) Stick Shaker Motor:
Two stick shaker motors are connected to the pilot and copilot
control columns. Upon activation, the motor provides a warning by
shaking the control column. The pilot’s stick shaker motor is
connected to the Left Essential DC bus through the SHAKER #1
circuit breaker and a relay activated by Data Acquisition Unit (DAU)
#1. The copilot’s stick shaker motor is connected to the Right Main
DC bus through SHAKER #2 circuit breaker and a relay activated by
DAU #2. If a failure is detected in either stick shaker, a message is
prompted for display on CAS.
(2) Stick Pusher Actuator:
The stick pusher actuator is controlled through two independent
electro-hydraulic valves. The left valve is controlled by either
channel of DAU #1 and Fault Warning Computer (FWC) #1, and the
right valve is controlled by either channel of the DAU #2 and FWC
#2, in order to eliminate single point failures. The #1 stick pusher
valve is connected to the Left Essential DC bus through the STALL
BARR VALVE #1 circuit breaker and the PUSH #1 relay. The #2
stick pusher valve is connected to the Right Main DC bus through
the STALL BARR VALVE #2 circuit breaker and the PUSH #2 relay.
The actuator is normally loaded in the retracted position. When the
stick pusher is activated, a message is prompted for display on
CAS, the solenoids are energized and the actuator extends.
Extension of the actuator provides mechanical force to push the
control column forward.
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The stall barrier system is constantly monitored for faults or failures.
If a fault or failure in the control column pusher is detected,
messages are prompted for display on CAS.
(3) Angle of Attack Probe:
Two heated Angle of Attack (AOA) probes provide aircraft AOA
information via the Aeronautical Radio Incorporated (ARINC) 429
databus to the DAUs. The left AOA probe provides information to
DAU #1 and the right AOA probe provides information to DAU #2.
AOA values are displayed on the Primary Flight Display (PFD).
The AOA system is constantly monitored for faults or failures. If a
probe heater failure occurs, or data transmitted by the probes to
their respective DAUs is in gross disagreement, messages are
prompted for display on CAS.
3. Controls and Indications:
(See Figure 3 through Figure 6.)
NOTE:
A full description of the Primary Flight Display can be
found in section 2B-02-00: Electronic Display System
Description. A full description of the FLIGHT
CONTROLS synoptic page can be found in section
2B-03-00: Engine Instruments and Crew Alerting
System Description.
A. Circuit Breakers (CBs):
The pitch flight controls system is protected by the following CBs:
Circuit Breaker Name: CB Panel: Location: Power Source:
L ELEV TRIM HEAT REER E-16 R MAIN AC Bus
R ELEV TRIM HEAT REER F-16 R MAIN AC Bus
LEFT ELEV HYD S/O POP C-5 L ESS DC Bus
RIGHT ELEV HYD S/O CPOP C-5 R ESS DC Bus
STALL BARR VALVE #1 POP E-5 L ESS DC Bus
STALL BARR VALVE #2 CPOP E-5 R ESS DC Bus
SHAKER #1 POP E-6 L ESS DC Bus
SHAKER #2 CPOP E-6 R MAIN DC Bus
AOA PROBE #1 POWER LEER F-4 L ESS DC Bus
AOA PROBE #2 POWER REER F-13 R ESS DC Bus
B. Crew Alerting System (CAS) Messages:
CAS messages associated with the pitch flight controls system are as
follows:
Area Monitored: CAS Message: Message Color:
Stall Warning System AOA MISCOMPARE Amber
AOA Probe(s) AOA PROBE 1/2 FAIL Amber
Integrated Avionics Computer(s) EL MISTRIM NOSE DN Amber
Integrated Avionics Computer(s) EL MISTRIM NOSE UP Amber
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Area Monitored: CAS Message: Message Color:
L/R Elevator Deactivation
Valve(s) L/R ELEV HYD OFF Amber
Electric Pitch Trim ELEV TRIM INOP Amber
Integrated Avionics Computer(s) MACH TRIM INOP Amber
Stall Warning System STALL BARRIER 1/2 Amber
STALL BARR Switch STALL BARRIER OFF Amber
Stall Warning System STCK PSH 1/2 FAULT Amber
Stall Warning System STICK PUSH UNAVAIL Amber
Stall Warning System STICK PUSH 1/2 FL Amber
Electric Pitch Trim EL TRIM DN LIMIT Blue
Electric Pitch Trim EL TRIM UP LIMIT Blue
Stall Warning System STICK SHAKE 1/2 FL Blue
4. Limitations:
A. Stall Barrier / Stall Warning:
(1) Takeoff Requirements:
Both stall warning / stall barrier systems must be operative for
takeoff.
(2) Use of System:
Operative stall barrier systems must be ON during all flight
operations, unless required to be selected OFF for procedural
reasons. Refer to Section 05-13-50: Stall Barrier Malfunction, for
additional information.
B. Mach Trim / Electric Elevator Trim Inoperative Speed:
With both mach trim compensators inoperative, or electric elevator trim
inoperative, the maximum operating limit speed is 0.80 MT.
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Stall Barrier System Block
Diagram
Figure 3
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Pitch Flight Controls
System Controls and
Indications (Cockpit
Center Pedestal)
Figure 4
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Electric Pitch Trim Engage / Disengage Switch Controls and Indications
Figure 5
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Electric Pitch Trim / Stall Barrier System Controls and Indications
Figure 6
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2A-27-30: Yaw Flight Controls System:
1. General Description:
The purpose of the yaw flight controls system is to provide the crew with a means
of controlling aircraft movement about the vertical axis when aircraft speed allows
aerodynamic use of the rudder. The rudder is a manually and electrically
controlled, mechanically actuated, hydraulically boosted airfoil located on the
trailing edge of the vertical stabilizer. Total rudder travel is 22 degrees in either
direction. Rudder movement is depicted on the FLIGHT CONTROLS synoptic
page.
Movement of the aircraft around the yaw axis is accomplished by the movement
of the rudder pedals which transmit inputs through conventional mechanical
linkage (cables and bellcranks) to displace the rudder. A dual hydraulic actuator
boosts inputs to the rudder surface. The rudder can be operated without the
assistance of the hydraulic actuator. In the event of a loss of both normal hydraulic
systems, the Auxiliary Hydraulic (AUX) system can also supply power to the
rudder actuator in flight through the selection of the STBY RUD (Standby Rudder)
switch on the lower portion of the pilot’s flight panel.
2. Description of Subsystems, Units and Components:
A. Automatic Hardover Prevention System:
Automatic hardover prevention is incorporated into the rudder boost
actuator. Switches monitor inputs to, and outputs from, the rudder actuator.
If the inputs and outputs disagree for 0.5 second or longer, hydraulic
pressure to the affected side of the actuator is shut off and a message is
displayed on the Crew Alerting System (CAS) display. The rudder is still
operative, but without benefit of the affected side’s hydraulic boost. The
rudder hardover prevention system receives power from the Right
Essential DC bus.
B. Redundant Hydraulic Power Sources:
Hydraulic power to the rudder actuator is normally provided by the L SYS
and R SYS. In addition, the AUX system can also supply power to the
rudder actuator in flight through the selection of the STBY RUD switch.
C. Automatic Overload Limiting System:
Force modulating valves within the rudder actuator provide protection of
the aircraft rudder against overload. Rudder surface movement is limited
by these valves when airspeeds increase airloads against the rudder.
When the hinge movement limit is reached, force-modulating valves shift,
reducing pressure to the limit actuator load output. This action causes a
logic-computed (blue RUDDER LIMIT) advisory message to be displayed
on CAS. Any further input force at the pedals cannot further displace the
rudder.
In addition to rudder limiting, hydraulic pressure to the rudder actuator is
monitored. Should a pressure differential between the L SYS and Right
Hydraulic System (R SYS) exceed 700-1000 psi, or a single summed
output pressure of less than 500 psi be detected, a blue SINGLE RUDDER
advisory message is displayed on CAS.
D. Standby Rudder System:
(See Figure 7.)
In the event of dual hydraulic system failure in flight, a Standby Rudder
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system can be activated. Selection of the STBY RUD switch, located on
the lower portion of the pilot’s flight panel, to ON activates the AUX pump.
AUX pressure opens a pilot-pressure operated valve to provide pressure to
the rudder actuator and a message is displayed on CAS. System operation
can be viewed on the HYDRAULICS synoptic page. Rudder operation will
be normal from the flight crew’s perspective until nose landing gear weight-
on-wheels is achieved, at which point AUX pressure is removed from the
actuator. The Yaw Damper system continues to function normally while the
Standby Rudder system is in operation.
E. Yaw Damper System:
(See Figure 7.)
A yaw damping channel is integrated within each of the two Flight
Guidance Computers (FGCs). Inputs received by the active FGC is cross-
checked and valid commands are transmitted to the rudder dual trim servo
for output to the rudder actuator. Detected faults prompt messages for
display on CAS. The yaw damper system is fail-operational in that should
one yaw damping channel (or FGC) fail, the remaining FGC can support
full system operation.
Both yaw damper channels are controlled using the YAW DAMP ENG/
DISENG switch, located on the lower portion of the pilot’s flight panel.
Power to the Yaw Damper system is provided by the Left Essential DC Bus
(YD 1) and Right Essential DC Bus (YD 2). When the ENG/DISENG switch
is selected to DISENG, the system is disengaged and a message is
displayed on CAS.
F. Rudder Trim System:
(See Figure 8.)
The Rudder Trim system allows fine position adjustment of the rudder to a
desired position. This is accomplished through a trim adjustment knob,
located on the cockpit center pedestal, to a maximum indicated 7.5
degrees left or right. There is no rudder trim tab on the rudder control
surface; mechanical trim inputs to the rudder actuator offset the entire
surface.
3. Controls and Indications:
(See Figure 7 and Figure 8.)
NOTE:
A full description of the FLIGHT CONTROLS and
HYDRAULICS synoptic pages can be found in section
2B-03-00: Engine Instruments and Crew Alerting
System Description.
A. Circuit Breakers (CBs):
The yaw flight controls system is protected by the following CBs:
Circuit Breaker Name: CB Panel: Location: Power Source:
RUDDER HYD S/O CPOP C-3 R ESS DC Bus
IAC #1 POP D-8 L ESS DC Bus
IAC #2 CPOP D-8 R ESS DC Bus
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B. Crew Alerting System (CAS) Messages:
CAS messages associated with the yaw flight controls system are:
Area Monitored: CAS Message: Message Color:
Rudder System Deactivation
Valve RUDDER HYD OFF Amber
Yaw Dampers 1 and 2 YD 1-2 FAIL Amber
YAW DAMP Switch YAW DAMPER OFF Amber
Rudder System Logic RUDDER LIMIT Blue
Rudder Load Limiter Sensor SINGLE RUDDER Blue
4. Limitations:
A. Yaw Damper Inoperative Speed:
(1) Above 10000 Feet:
The maximum speed is 260 KTS / 0.80 MT.
(2) Below 10000 Feet:
The maximum speed is 250 KCAS.
(3) Above 20000 Feet:
The minimum speed is 210 KTS.
(4) Below 20000 Feet:
The minimum speed is in accordance with the following schedule
until ready to configure for approach and landing. VREF as presented
on the airspeed tape is the approach speed for landing in the current
flap setting.
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Sea Level to 5000 ft 5000 to 20000 ft
Fuel - lb Flaps 0, 10, 20 Flaps 39 Flaps 0, 10, 20 Flaps 39
23000 VREF VREF VREF 135
24000 VREF VREF VREF 141
25000 VREF VREF VREF 147
26000 VREF VREF VREF 153
27000 VREF VREF VREF 159
28000 VREF VREF 147 160
29000 VREF VREF 153 160
30000 VREF VREF 158 160
31000 VREF VREF 163 160
32000 VREF VREF 168 160
33000 VREF VREF 174 160
34000 VREF VREF 179 160
35000 VREF VREF 184 160
36000 VREF VREF 189 160
37000 VREF VREF 195 160
38000 VREF VREF 200 160
39000 VREF VREF 205 160
40000 VREF 154 211 160
41000 VREF 158 216 160
41300 VREF 159 217 160
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Yaw Flight Controls System Controls and Indications
Figure 7
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2A-27-40: Roll Flight Controls System:
1. General Description:
The ailerons are manually and electrically controlled, mechanically actuated and
hydraulically boosted airfoils located on the outboard trailing edge of the left and
right wings. Aileron travel is eleven (11) degrees up and down. Aileron movement
is visually depicted on the FLIGHT CONTROLS synoptic page.
Aircraft roll is accomplished by movement of the control wheels which transmits
motion through conventional mechanical linkage (cables, cranks and pushrods) to
displace the ailerons. Hydraulic boost is provided by a hydraulic boost actuator to
assist inputs from the control wheels. During normal flight conditions, aileron
actuators are powered simultaneously by Left and Right Hydraulic Systems (L
Rudder Trim Controls
Figure 8
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SYS and R SYS) at 3000 psi each. In the event of a single hydraulic system
failure, the actuators will remain powered by the operative hydraulic system.
2. Description of Subsystems, Units and Components:
A. Automatic Hardover Prevention:
Each aileron actuator incorporates a hardover prevention system which
compares inputs and outputs. If inputs to, and outputs from, either actuator
do not agree for over 0.5 second, hydraulic pressure from both hydraulic
systems to both actuators is shut off simultaneously and a message will be
displayed on the Crew Alerting System (CAS) display. The ailerons are still
operative, but without benefit of hydraulic boost. The hardover prevention
system receives power from the Left and Right Essential DC Bus.
B. Aileron Control Separation:
The aileron control systems are dual and separable. In the unlikely event
that a single aileron control system were to become jammed, an aileron
disconnect system provides the means to separate the left and right
ailerons from each other. This is accomplished through the use of an
aileron disconnect system, located on the right side of the cockpit center
pedestal and labeled AIL DISC on its protective cover.
The aileron disconnect system consists of the protective cover, the aileron
disconnect handle (labeled AIL DISC) and a power assist trigger (labeled
LIFT). If a jammed aileron is detected, the protective cover is raised and
the AIL DISC handle is pulled. With the handle fully extended, the pilot’s
and copilot’s control wheels are separated and the aileron systems are fully
isolated and independent – the pilot controls the left aileron and the copilot
controls the right aileron. The immovable aileron side will remain
immovable and the movable side is now free to be used to control the
aircraft. If the power assist trigger was NOT used to actuate the AIL DISC
handle, the aileron disconnect system can be reset at the discretion of the
flight crew.
If the force required to pull the AIL DISC handle is too great, the power
assist trigger, located under the AIL DISC handle, is pulled. This actuates a
gas-spring cartridge, providing between 110 to 150 pounds of controlled
force to fully extend the handle and separate the aileron control systems.
NOTE:
If the power assist trigger was used to actuate the AIL
DISC handle, the system cannot be reset without the
use of special tools.
C. Flight Spoiler Supplementation:
The outboard two spoilers on each wing act as flight spoilers in conjunction
with the ailerons to improve roll response of the aircraft. Their function as
supplements to the ailerons is fully automatic and transparent to the crew.
Spoiler travel will vary in accordance with the control wheel inputs up to a
maximum of 47(±3) degrees upward deflection. The spoilers are
hydraulically powered, normally by both L SYS and R SYS, but will function
normally with only one source of hydraulic pressure. The operation of flight
spoilers (as well as ground spoilers and speed brakes) can be inhibited
using the SPOILER CONTROL switch, located on the left side of the
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cockpit center pedestal. Selection of the switch to off causes a message to
be displayed on CAS.
D. Aileron Trim System:
(See Figure 9.)
(1) Aileron Trim Tab:
The aileron trim tab is a manually operated, mechanically actuated
airfoil located on the inboard trailing edge of the left aileron. The
aileron trim tab moves opposite to the left aileron (tab trailing edge
down, left aileron trailing edge up; tab trailing edge up; left aileron
trailing edge down) to control aircraft roll trim. It is mechanically
actuated by an aileron trim wheel which is located on the aft end of
the cockpit center pedestal. The range of travel for the aileron trim
tab is 15 degrees up or down.
(2) Aileron Trim Tab Actuator Heater:
The aileron trim tab actuator heater provides heat to prevent the trim
tab actuator from freezing. The heater provides heat until the
temperature rises to a threshold temperature, predetermined by the
manufacturer. The heater is controlled by the AILTRIM HEAT circuit
breaker on the Right Electronic Equipment Rack (REER) and
receives power from the Right Main AC bus.
3. Controls and Indications:
(See Figure 9.)
NOTE:
A full description of the FLIGHT CONTROLS synoptic
page can be found in section 2B-03-00: Engine
Instruments and Crew Alerting System Description.
A. Circuit Breakers (CBs):
The Roll Flight Control system is protected by the following CBs:
Circuit Breaker Name: CB Panel: Location: Power Source:
LEFT AIL HYD S/O POP C-4 L ESS DC Bus
RIGHT AIL HYD S/O CPOP C-4 R ESS DC Bus
SPLR FLT PWR S/O CPOP E-3 R ESS DC Bus
AIL TRIM HEAT REER D-16 R MAIN AC Bus
B. Crew Alerting System (CAS) Messages:
CAS messages associated with the Roll Flight Control system are:
Area Monitored: CAS Message: Message Color:
L/R Aileron System Deactivation
Valve L/R AIL HYD OFF Amber
Spoiler System Deactivation
Valve SPOILERS HYD OFF Amber
4. Limitations:
There are no limitations established for the roll flight controls system at the time of
this revision.
OPERATING MANUAL
PRODUCTION AIRCRAFT SYSTEMS2A-27-00
Page 24
May 22/01
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