Gulfstream G550 User manual

FLIGHT CONTROLS
2A-27-10: General
1. General Description:
The aircraft flight controls allow the flight crew to guide the aircraft in the
longitudinal, vertical and horizontal axes (see Figure 2). The primary flight controls
are:•Elevator to control aircraft pitch
•Rudder to control aircraft yaw
•Aileron to control aircraft roll
The primary flight controls are positioned by moving the pilot and copilot control
yokes and rudder pedals. Both yokes are mechanically linked together so that
either crew position has full control authority and control inputs are transparent to
both crew members since movement of one set of controls will move the
corresponding set of controls. Each yoke has a dedicated cable connection to the
elevator and aileron control on that respective side, but since the yokes are
mechanically linked, moving one side elevator or aileron will move the flight
controls on both sides. For instance, the copilot yoke is cable-linked to the right
elevator and the right aileron but any movement of the copilot yoke also moves
the pilot yoke that is in turn cable-linked to the left aileron and left elevator. This
system of split control authority and linked yoke movement provides a means to
maintain some flight control movement if one of the flight controls or associated
linkages becomes jammed. If a malfunction prevents movement of either the
elevators or ailerons, the mechanical links joining the two pilot yokes can be
severed, thereby allowing movement of the left or right elevator or aileron that
remains operational by commands using the yoke connected to the side of the
free control surface.
The two sets of rudder pedals are similarly mechanically linked together, but both
are connected to the rudder by a single cable. For this reason, there is no
provision for interrupting the linkage between the two sets of pedals, since each
set does not have an independent route to the rudder.
The cable connections from the yokes and rudder pedals are continuous loop
installations, providing feedback to the moveable controls. The control cables
engage bell cranks that translate cable movement into displacement commands
for hydraulic actuators that boost contol inputs to move the flight controls. Each
hydraulic actuator has a single shaft, but dual piston chambers in order that the
actuator may be driven by both (or either) left and right hydraulic system. (The
hydraulic system power sources for the flight controls are shown in Figure 1.) The
rudder actuator may also be powered by the Auxiliary (AUX) hydraulic system in
the event of dual hydraulic power failure. Each of the hydraulic actuators is
connected to the associated flight control by pushrods and bell cranks to impart
mechanical movement to the control surface. A bungee piston filled with hydraulic
fluid moderates the rate of actuation of the flight control and provides an artificial
feel input to the flight crew through the closed loop cable system. The failure of a
single hydraulic system does not degrade flight control operation - the remaining
system provides adequate power for flight control movement, and the actuator
chamber for the failed system bypasses fluid so there is no resistance to flow. If
both (or all, in the case of the rudder) hydraulic systems fail, both actuator
chambers bypass fluid, and pilot input through the cable connection moves the
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internal shaft of the actuator and the associated pushrods and bellcranks without
powered assistance. More pilot force is required, but full movement of the flight
controls is attainable.
The autopilot is interfaced with the primary flight controls through electric servos
to move parallel cable linkages to the hydraulic actuators. Each flight control
surface has a Linear Variable Displacement Transducer (LVDT) that provides an
electrical signal proportional to flight control surface displacement from neutral.
The LVDT signal provides feedback to the autopilot for positioning the flight
controls, and also communicates control surface position for display on systems
and synoptic windows through interface with the Modular Avionics Units (MAUs).
The description of the Digital Automatic Flight Control Systems (autopilot) makes
up the entire content of Section 2B of this manual, and for that reason is not
covered in this Section. However, for the convenience of the reader, a tabulation
of the Crew Alerting System (CAS) messages associated with the autopilot is
included below.
The elevators and ailerons have trim tabs to position the flight controls with
aerodynamic forces to moderate the amount of physical effort to maintain the
control surfaces in the desired steady-state condition. The rudder is not equipped
with a trim tab, but instead has a mechanical trim input to reset the neutral
position of the rudder. The mechanical trim uses the hydraulic actuator to project
the rudder slightly into the windstream in the desired direction to compensate for
induced yaw.
The secondary flight controls and functions are:
•Wing flaps - enhance wing lift characteristics
•Flight and Ground Spoilers - reduce wing lift and add overall drag
•Movable horizontal stabilizer - aligns the elevator with aircraft angle of
attack
•Yaw damper - uses the autopilot rudder servo to moderate aircraft heading
oscillation
Some functions of the secondary flight controls are integrated with the operation
of the primary flight controls and other functions mutually complement the
operation of other secondary controls.
•As flaps extend, moving the wing center of lift aft, a downward pitch
moment is created - the moveable stabilizer automatically compensates for
the pitch moment by trimming downward. The opposite movement occurs
as flaps are retracted and the nose of the aircraft pitches up.
•When ailerons are used, flight spoilers on the downward wing activate to
increase roll rate and provide yaw into the turn. Ground spoilers decrease
ground roll distance during landings and aborted takeoffs.
•The yaw damper provides a degree of turn coordination for aileron roll
commands, provided wing flaps are not selected to more than thirty
degrees (30°).
Safety features incorporated into the flight controls system include:
•A stick shaker warning and a stick pusher stall prevention actuator.
•A gust lock that prevents damage to flight controls while the aircraft is
secured on the ground.
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2. Primary and Secondary Flight Controls Subsections:
The primary and secondary flight controls are discussed in the following
subsections:
•2A-27-20: Elevator Pitch Control
•2A-27-30: Rudder Yaw Control
•2A-27-40: Aileron Roll Control
•2A-27-50: Horizontal Stabilizer and Wing Flaps
•2A-27-60: Stall Warning and Prevention System
•2A-27-70: Speed Brake and Ground Spoilers
•2A-27-80: Flight Controls Gust Lock
3. Autopilot Crew Alerting System (CAS) Messages:
The following CAS messages are associated with the operation of the autopilot
and integrated subsystems:
Area Monitored: CAS Message: Message Color:
Flight Guidance Computer
Internal Monitor AP 1-2 Fail Amber
Flight Guidance Panel AP
Engage Switch AP Control Switch Stuck Blue
AP Engage Switches AP Engage Inhibit -Sw
Active Blue
Air Data System AP Inhibit - ADS Blue
Inertial Reference System AP Inhibit - IRS Blue
Control Column Force AP Inhibit - Left Column Blue
Control Column Force AP Inhibit - Right
Column Blue
Control Wheel Force AP Inhibit - Left Wheel Blue
Control Wheel Force AP Inhibit - Right Wheel Blue
Manual Trim Wheel AP Inhibit - Man Trim
Active Blue
Autopilot Quick Disconnect
Switch AP Inhibit - QD Blue
Stall Shaker AP Inhibit - Stall Blue
Autopilot Touch Control
Steering Switch AP Inhibit - TCS Blue
Control Wheel Electric Trim
Switch AP Inhibit -Trim Cmd Blue
Weight On Wheels (WOW)
System AP Inhibit - WOW Blue
Autopilot Power Source AP 1-2 Power Fail Blue
Autopilot Elevator Trim Servo AP / Trim Fail Blue
Flight Guidance Panel Speed
Window Check Speed Target Blue
Flight Guidance Computer /
WOW System FGC - WOW Fault Blue
Take Off and Go Around
(TOGA) Engage Switch / Flight
Guidance Panel Manual Speed
Go Around Pitch Blue
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Flight Controls System:
Simplified Fluid Power
Diagram
Figure 1
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Flight Controls System Components
Figure 2
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2A-27-20: Elevator Pitch Control
1. General Description:
The aircraft has a dual elevator installation to control aircraft pitch attitude. The
elevators are composed of a baked graphite-epoxy material. Each of the cockpit
yokes is connected to one of the aircraft elevators. The pilot yoke is connected to
the left elevator, the copilot yoke to the right elevator. Each yoke is also connected
to the other by a mechanical torque tube beneath the cockpit floor. Since both
yokes are interconnected, moving one yoke moves both elevators.
Braided steel cables run from each yoke to hydraulic assist actuators in the tail of
the aircraft. The cables are routed beneath the aircraft floor using pulley
connections to clear other installed equipment. The cables mate with the hydraulic
assist actuators via bellcranks that translate pulley rotational motion into forward
and aft motion. The actuators each have a single shaft powered by two piston
chambers, one chamber for each (left and right) hydraulic system. Both hydraulic
systems normally power the actuators, but one system is sufficient for full elevator
movement. The actuators are connected to the respective elevator by linkages
and bellcranks, moving the elevator up or down about the pivot points on the aft of
the horizontal stabilizer. The deflection range of the elevators is twenty-four
degrees (24°) up and thirteen degrees (13°) down.
Each connection of yoke to elevator is a continuous loop. Incorporated into the
loops adjacent to the actuators is a bungee cylinder filled with viscous fluid to
resist yoke / elevator movement in order to provide artificial feel to each yoke.
Each elevator also has a stability spring incorporated into the cable linkage to
provide a forward pull to the control yoke and to contribute additional feel input.
Both sides of the hydraulic actuators are monitored to assure correct operation.
The cockpit cable input motion must result in a corresponding actuator output
motion, and similarly the output side of the actuator should not move without
cockpit input. If input and output do not correspond, actuator hydraulic pressure is
bypassed to prevent movement of the elevator.
Anytime hydraulic pressure to the actuators is bypassed or lost (in the instance of
dual hydraulic system failure) the elevators remain operable with manual yoke
movement that positions the actuator shaft and connecting linkages to the
elevator. Control forces will be higher, since normal hydraulic assist provides a six
(6) to one (1) boost advantage to move the elevator surfaces.
Each elevator is equipped with a trim tab that uses aerodynamic pressure to aid in
positioning the control surface. The trim tabs are controlled manually by rotating a
wheel on the cockpit pedestal or electrically using switches on the control yokes.
Manual trim uses a dedicated braided wire connection from the cockpit to a
mechanical linkage in the tail. Electrical switch trim movement commands an
electric servo to move the same linkage.
Both the elevators and elevator trim incorporate Rotary Variable Differential
Transducers (RVDTs) to feed back position information to the autopilot for
elevator control and trim and to the ModularAvionics Units (MAUs) for formulation
of control surface position display on the Flight Controls 2/3 synoptic page. RVDTs
measure the angle of the elevators and trim tabs and transmit an electrical signal
proportional to displacement from a neutral position.
When the autopilot is engaged, the elevators are positioned by electric servos that
move parallel cable connections to the hydraulic actuators. The autopilot also
uses the electric trim servo to move the trim tabs, minimizing hydraulic actuator
force.
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If a malfunction or failure in any portion of the loop between a cockpit yoke and the
corresponding elevator prevents control surface movement, the mechanical
torque tube connection between the two control yokes can be separated to allow
control of the aircraft with the free (unjammed) elevator. The autopilot may be
used in single elevator operation.
2. Description of Subsystems, Units and Components:
A. Elevator Hard Over Prevention System (HOPS):
Movements of the elevators contrary to the commanded position are
limited by a Hard Over Prevention System (HOPS), illustrated in Figure 3.
The system incorporates eight switches for each elevator to monitor
mechanical and hydraulic elevator operation. Four external mechanical
switches are integrated into the elevator control linkage to provide a
comparison reference for four switches mounted internally within the
hydraulic actuator. Of the four external switches, two are for left hydraulic
system reference and two are for right hydraulic system reference. Of the
two switches for each hydraulic system, one provides a up elevator
command reference and the other provides a down elevator command
reference. The switches are plunger-type contact switches, and are
installed on each side of a bracket attached to the command input side of
the elevator actuator. On one side of the bracket are the up elevator
command input switches for the left and right hydraulic systems. On the
other side of the bracket are the down elevator command input switches for
the left and right hydraulic systems. Inserted between the switches in the
bracket is a cam-type arm mated to the elevator hydraulic actuator output
linkage. The cam arm is positioned with a defined amount of clearance
between the plunger-type switches. Under normal conditions, the bracket
holding the switches moves with elevator command input and the cam arm
moves with the elevator hydraulic actuator output, so the clearance
between the switches and the arm is maintained.
If a malfunction occurs and the elevator moves opposite to or further from
the commanded direction, the cam arm that is attached to the output
linkage of the elevator actuator will move to close the clearance gap
between the cam arm and the plunger-type switches, making contact with
the switches on the side of the bracket. When the plungers of the switches
are depressed, a relay closes and an electrical signal is sent to a
corresponding set of switches mounted internally within the hydraulic
actuator.
Four pressure switches monitor left and right hydraulic system pressures
within the pistons of the elevator actuator. In normal conditions, all four
switches sense stabilized pressures since hydraulic outputs positioning the
elevator are balanced by air load pressures on the elevator surface acting
against actuator pressures. When a malfunction occurs and the elevator
moves contrary to the commanded direction, the hydraulic actuator shaft
moves in the contrary direction, causing an increase in hydraulic pressure
on the opposite sides of the pistons within the actuator. The left and right
system opposite side pressure switches close, completing the circuit
initiated by closure of the bracket plunger switches, and an electrical signal
is sent to a timing relay. If the contrary elevator movement persists for
longer than one tenth (1/10) of a second, hydraulic pressure from both left
an right systems is shut off to the elevator actuator.
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Left Hydraulic System Right Hydraulic System
Up
Elevator
←
Down
Elevator
→
Up
Elevator
←
Down
Elevator
→
Hydraulic Actuator Shaft
Pressure
Switch Pressure Switch Pressure Switch Pressure Switch
In a similar manner, if a hydraulic malfunction causes the shaft of the
actuator to move in a direction opposite the commanded elevator direction,
the increased pressures in the wrong direction sides of the pistons would
close the monitoring switches, and movement of the actuator shaft would
cause the cam arm to contact the plunger switches on the external bracket,
and hydraulic pressure would be shut off to the elevator actuator after a
one tenth (1/10) of a second delay.
If a hydraulic malfunction in a single system (left or right) side of the
actuator moves the actuator shaft in a wrong direction, only the hydraulic
pressure of the malfunctioning system is shut off. For instance in a
stabilized condition, if the left hydraulic system piston attempts to move the
actuator shaft in the up direction, the increased pressure in the down
elevator side of the left piston will close the monitor switch and actuator
shaft displacement will close both the left and right hydraulic system up
elevator bracket plunger switches, completing the shut off circuit for the left
hydraulic system after a one tenth (1/10) second delay. (In this case
hydraulic pressure in the up elevator side of the right hydraulic system
piston will decrease due to the increase in area caused by actuator shaft
movement.) The elevator hydraulic actuator will continue to function using
the remaining hydraulic system.
The operation of the hydraulic shut off valves by the HOPS is signaled to
the MAUs (left elevator to MAU #1, right elevator to MAU #2) over ARINC-
429 connections. The shut off condition is monitored by the MWS, and a
CAS message corresponding to the condition is displayed on the CAS
window. If either or both hydraulic systems are shut off, an amber caution
message of “L (or) R Elevator Hydraulics Off” is displayed. If only a single
hydraulic system has been shut off, the remaining hydraulic system will
provide full elevator operation. If both hydraulic systems have been shut
off, manual elevator control may remain possible, depending upon the
cause of the hardover condition. If the cause of the condition is thought to
be momentary, and the use of the elevator is deemed necessary for
continued safe flight and landing, the hydraulic shut off valve(s) may be
reset by cycling the RIGHT ELEV HYD S/O and/or LEFT ELEV HYD S/O
circuit breaker. If the cause has not been rectified, the shut off valve(s) will
close and hydraulic boost for the elevator will be unavailable if both
hydraulic systems have been shut off. (HOPS is powered by the left
essential DC bus for the left elevator and the right essential DC bus for the
right elevator.)
To prevent the HOPS from shutting off hydraulic system pressure to the
elevator during normal flight maneuvers that may involve rapid changes in
elevator direction, HOPS activation is buffered by three elements:
•The clearance between bracket plunger switches and the actuator
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cam arm
•Inertia in the build up of hydraulic pressures in the elevator actuator
•The one tenth (1/10) second delay in electrically latching the shut off
of hydraulic pressure
B. Elevator Disconnect Handle:
(See Figure 4.)
The mechanical torque tube beneath the cockpit connecting the pilot and
copilot yokes may be disconnected if a malfunction in one of the cable
connections, hydraulic actuators or elevators renders the respective
elevator inoperative. Disconnecting the torque tube prevents both elevators
from being disabled by a malfunction in one elevator linkage.
An elevator disconnect handle is located on the pilot side of the center
pedestal beneath a protective cover. The handle is connected by a cable to
a pin securing the two halves of torque tube together. Pulling out on the
disconnect handle removes the pin and allows each yoke to move
independently. If the elevator linkage malfunction has resulted in opposite
movement between the two yokes, making retraction of the mating pin
difficult, a power assist gas-spring cartridge may be activated to provide
additional force to remove the pin. The power assisted disconnect is
activated by pulling a trigger beneath the disconnect handle, and provides
an upward lifting moment of thirty-three feet per second (33 ft./sec).
After the yokes have been separated, the malfunctioning elevator is
isolated and the operable elevator may be used to control the aircraft. with
manual or autopilot inputs. If it is discovered that separating the yokes
freed the previously malfunctioning elevator linkage, the yokes may be
reconnected by pushing in on the disconnect handle when the yokes are
aligned, provided that the power assist disconnect was not used. A reset of
the yoke torque tube coupling is not possible without special maintenance
tools after activation of the power assist disconnect.
C. Pitch Trim System:
(See Figure 4 through Figure 6.)
Each elevator has a trim tab installed on the trailing edge. The trim tabs are
manufactured from the same graphite-epoxy material as the elevators, but
incorporate a ceramic heating element that is continuously electrically
powered to maintain a temperature of one hundred seventy-five degrees
Fahrenheit, plus or minus twenty degrees (175°F±20) around the tab
actuator linkage. Elevator trim heat is powered by 115V AC from the right
main bus. The trim tabs have a range of movement of twenty-two degrees
(22°) trailing edge down (aircraft nose up) to eight degrees (8°) trailing
edge up (aircraft nose down). Limit switches are installed at the travel limits
that will prompt the display of Crew Alerting System (CAS) messages
notifying the crew that the elevator trim tabs are at maximum displacement.
If a trim limit message is displayed while the autopilot is engaged, extreme
care should be taken prior to disengaging the autopilot. An abrupt attitude
change will occur if trim displacement is not moderated prior to
disengaging the autopilot.
Operation of the trim tabs employs aerodynamic force to maintain the
elevator in the desired position. As the trim tab is moved from the neutral
position (faired with the elevator) into the airstream, the air impinging on
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the tab forces the hinged elevator into the opposite direction. As the
elevator is moved from the neutral position (faired with the horizontal
stabilizer), it also encounters pressure forces from the airstream, thus the
amount of elevator movement from neutral is determined by a balance of
airstream forces acting on both the trim tab and the elevator. Since the
surface area of the trim tab exposed to the airstream is less than the
surface area of the elevator, the elevator is deflected in much smaller
increments than trim tab displacement (excluding other factors, trim tab
effectiveness is a ratio of tab surface area to elevator surface area).
The flight crew manually controls the amount of trim tab deflection by
moving control wheels on either side of the center console. The wheels are
hubs connected to a common axial shaft, so that moving one wheel moves
the other. The shaft is connected to a continuous loop of wire cables that
connect through a series of pulleys and bellcranks to the elevator trim tabs.
Rotating a trim wheel forward positions the trim tab up forcing the elevator
down resulting in an aircraft nose down moment. Trim wheel rotation aft
results in an aircraft nose up moment.
The flight crew has the option of electrically moving the elevator trim tabs.
A pushbutton, labelled PITCH TRIM ENG / DISENG, located to the left of
the standby flight instruments on the lower instrument panel enables
electrical operation of trim switches mounted on the outboard side of the
control yokes. Electric pitch trim is normally engaged. The amber DISENG
legend in the pushbutton will illuminate if the button is not pushed in to
engage electric trim. The yoke trim switches are composed of split halves.
Both halves must be moved in the same direction to move the elevator
trim. The split switch design helps to prevent accidental trim input. The
switches are wired to an electric servomotor that is located in the tail of the
aircraft and incorporated into the cable linkage to the trim tabs. An electric
signal from a cockpit trim switch results in the servomotor rotating an
attached pulley, moving the elevator trim cable loop in the desired
direction. The manual trim wheels on the pedestal will rotate with electric
trim inputs, since the control cabling is a continuous loop.
When the autopilot is engaged, autopilot servomotors move the elevators
and also control elevator trim with the same servomotor employed by the
yoke electric trim switches. The PITCH TRIM ENG / DISENG switch is
automatically engaged whenever the autopilot is engaged (to enable
autopilot trim). However the reverse is not true - electric pitch trim will not
disengage when the autopilot is disengaged, but must be selected off with
the switch.
Elevator trim must be within a defined range for takeoff, with the specific
setting within the range determined by aircraft Center of Gravity (COG) and
takeoff gross weight. The limits of the acceptable pitch trim range are eight
degrees nose up to nineteen degrees nose up (8° - 19° up). The takeoff
elevator trim range is marked in green on the trim setting indices of the
manual trim wheels at each side of the center pedestal. The same limits
are shown in a green band on the trim scale of the Flight Controls 2/3
synoptic page. Failure to set pitch trim within the defined range will result in
a warning annunciation and CAS message as the power levers are
advanced for takeoff.
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D. Mach Trim:
The autopilot employs elevator control to provide the aircraft with Mach
trim. Mach trim is necessary because at high speed flight the center of lift
on the wing transits aft with increases in speed, producing a nose down
pitch moment termed Mach tuck. The autopilot electrically repositions the
elevator trim to neutralize the nose down force. Mach trim is an automatic
function of the Flight Guidance Computers.
NOTE:
The autopilot does not have to be engaged to provide
Mach trim. Automatic Mach trim is available whenever
the PITCH TRIM ENG/DISENG switch is engaged.
3. Controls and Indications:
(See Figure 4 through Figure 6.)
NOTE:
A full description of the Flight Controls 2/3 synoptic
page appears in section 2B-07-00.
A. Circuit Breakers (CBs):
The following CBs protect elevator pitch control:
Circuit Breaker Name: CB Panel: Location: Power Source:
ELEV SERVO #1 POP D-4 L ESS DC Bus
ELEV SERVO #2 CPOP D-4 R ESS DC Bus
PITCH TRIM SERVO #1 POP E-2 L ESS DC Bus
PITCH TRIM SERVO #2 CPOP E-1 R ESS DC Bus
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
B. Crew Alerting System (CAS) Messages:
The following CAS messages are associated with the elevator pitch
controls:
Area Monitored: CAS Message: Message Color:
Elevator Trim Tab RVDTs vrs
takeoff range Aircraft Configuration Red
Elevator and Trim Tab RVDTs and
MAUs Elevator Mistrim Nose Down Amber
Elevator and Trim Tab RVDTs and
MAUs Elevator Mistrim Nose Up Amber
L/R Elevator HOPS and MAUs L/R Elevator Hydraulics
Off Amber
Electric Pitch Trim Servos Elevator Trim 1-2 Fail Amber
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Area Monitored: CAS Message: Message Color:
Electric Pitch Trim Servos / Trim
ENG/DISENG Switch Mach Trim Off NOTE
(inhibited below 0.82
Mach)
Amber
Electric Pitch Trim Servos and
MAUs Pitch Trim 1-2 Power Fail Blue
Electric Pitch Trim Servos and
MAUs Elevator Trim Down Limit Blue
Electric Pitch Trim Servos and
MAUs Elevator Trim Up Limit Blue
4. Limitations:
A. Mach Trim / Electric Elevator Trim Functions:
(1) Use of Mach Trim / Electric Elevator Trim Functions:
Mach trim / electric elevator trim must be ON during all flight
operations except as provided for in Section 05-02-40: Mach Trim
Failure.
(2) With both Mach Trim / Electric Elevator Trim Inoperative:
M
mo
is reduced to 0.80M
T
.
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Hard Over Prevention
System (HOPS)
Figure 3 (Sheet 1 of 2)
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Hard Over Prevention
System (HOPS)
Figure 3 (Sheet 2 of 2)
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Pitch Flight Controls
System Controls and
Indications (Cockpit
Center Pedestal)
Figure 4
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