Gulfstream IV User manual

FLIGHT CONTROLS
2A-27-10: General
The Gulfstream IV primary flight controls system, shown in Figure 1, is a mechanically
actuated, hydraulically operated system that provides boosted surface control to
overcome the aerodynamic forces associated with high speed flight. This allows the
aircraft to be comfortably and reliably steered through the pitch, roll and yaw axes.
The primary flight control surfaces (elevators, ailerons and rudder) are positioned by
tandem type hydraulic actuators. The actuators receive hydraulic operating pressure
from both the Combined and Flight hydraulic systems, as shown in Figure 2. Both
hydraulic systems maintain a system pressure of 3000 psi. Loss of a single hydraulic
system has no effect on operation of the primary 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.
Mechanical pitch, roll and yaw trim systems allow the flight crew to trim the aircraft. The
pitch trim system can also be controlled electrically by pitch trim switches on the control
wheels.
A gust lock secures the elevators, ailerons and rudder to prevent wind gust damage to
the surfaces.
Secondary flight controls, shown in Figure 1, include flaps, ground spoilers and
speedbrakes. These flight controls are hydraulically powered and electrically or
mechanically controlled. The mechanically operated horizontal stabilizer moves in
conjunction with the flaps to maintain longitudinal trim.
AnAngle-of-Attack (AOA) system provides outputs to the control column shakers, control
column pusher, approach indexers, normalized AOA display and stall barrier system. The
control column shakers provides early warning of a stall scenario by vibrating the control
column before the stall while the control column pusher automatically initiates lowering
the nose if the stall is imminent.
The Gulfstream IV uses an aircraft configuration warning system to monitor landing gear,
flap, speed brake and power lever position. If an unsafe configuration is detected, the
system provides a visual and / or aural warning.
On CAA certified aircraft, a flight control automatic failure detection system compares
control inputs to actuator outputs. If a malfunction is detected, the system shuts off power
to the affected actuator.
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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GIV Flight Controls /
Aerodynamic Axes
Figure 1
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GIV Flight Controls Fluid
Power Diagram
Figure 2
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2A-27-20: Pitch Flight Control System
1. General Description:
Aircraft movement about the lateral axis (pitch) is controlled by the position of the
elevators. The elevators are manually controlled, mechanically actuated and
hydraulically boosted airfoils mounted on the trailing edge of the horizontal
stabilizer. Total elevator travel ranges from 24°trailing edge up to 13°trailing edge
down.
The elevators are positioned by a tandem type hydraulic actuator. The actuator
receives hydraulic operating pressure simultaneously from both the Combined
and Flight hydraulic systems during normal operations. Loss of a single hydraulic
system has no effect on operation of the elevators, as the remaining system is
capable of maintaining actuator load capacity. In the event of total loss of
hydraulic pressure in both systems, the elevators revert to manual operation.
Manual reversion is also possible through use of a flight power shutoff valve and
its pedestal-mounted control handle.
A pitch trim system is used to position a trim tab attached to the trailing edge of
each elevator. The tabs are positioned either manually by a pedestal-mounted
control wheel or electrically by pitch trim switches on the control wheels.
An Angle-of-Attack (AOA) system provides outputs to the control column shakers,
control column pusher, approach indexers, normalized AOA display and stall
barrier system. The control column shakers provides early warning of a stall
scenario by vibrating the control column before the stall while the control column
pusher automatically initiates lowering the nose if the stall is imminent.
On CAA certified aircraft,aflight control automatic failure detection system
compares control column inputs to elevator actuator outputs. If a malfunction is
detected, the system shuts off either or both hydraulic power sources to the
affected actuator.
The pitch flight control system consists of the following subsystems, units and
components:
•Control column system
•Mechanical actuation system
•Hydraulic boost system
•Manual reversion system
•Pitch trim system
•Angle-of-attack / stall barrier system
•Failure detection system (CAA aircraft only)
2. Description of Subsystems, Units and Components:
A. Control Column System:
(See Figure 7.)
The pilot’s and copilot’s control columns mount to a common transverse
torque tube supported by a bearing on each end. Moving the control
column fore and aft rotates the torque tube that, in turn, transmits the
inputs rearward through conventional mechanical linkage. Adjustable stops
on the torque tube limit control column movement to eight inches aft and
five inches forward of the neutral position.
An eddy current damper is connected to the bottom of the control column
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output crank. The damper is designed to sense control column movement
and provide a counteracting damping force or artificial feel. Damping force
is generated in proportion to control column velocity. Should the eddy
current damper jam or fail, its internal clutch releases to allow control
column movement.
B. Mechanical Actuation System:
(See Figure 3.)
Control column inputs are transmitted rearward through a system of push-
pull rods, bellcranks, cables and a sector assembly to an input cable sector
below the base of the vertical stabilizer in the tail compartment. Connected
to the input cable sector are the elevator actuator, autopilot servo and
stability springs. Final output of the sector is to the elevator actuator input
lever.
Movement of the actuator input lever displaces the elevator actuator’s
servo control valve, directing Combined and Flight hydraulic system
pressure to the actuator cylinders. When under pressure, the actuator body
moves while the its piston remains motionless. Movement of the actuator
body in turn moves its output crank. Output crank movement is transmitted
upward and aft to the left and right elevators through a system of push-pull
rods and cranks, resulting in the desired elevator deflection.
Stability springs, commonly referred to as down springs, provide
approximately 13 pounds of pull on the control columns. This pulling force
keeps pilot feel forces out of the friction band at low airspeeds.
C. Hydraulic Boost System:
The elevator actuator is a dual tandem actuator consisting of two pistons
secured to a common shaft. The pistons move inside a common cylinder
divided to create two separate cylinders. One cylinder receives Flight
hydraulic system pressure while the other cylinder receives Combined
hydraulic system pressure.
Mechanical movement of the actuator input lever moves the servo control
valve from its neutral position. The servo control valve then directs
hydraulic pressure to one of the actuator’s two cylinders and connects
each cylinder’s opposite side to return. When the elevator reaches the
desired deflection, the servo control valve shifts to its neutral position to
lock hydraulic pressure within the actuator, in effect preventing further
surface movement.
The elevator actuator also has an internal hydraulic damper that provides
damping force proportional to the square of its input velocity. This damping
action ensures operational stability for the elevators when hydraulically
boosted whereby a portion of the actuator’s output is fed back to the input
system.
D. Manual Reversion System:
During normal flight operations, the Combined and Flight hydraulic
systems each supply and maintain 3000 psi to the elevator actuator. Loss
of system pressure due to a single system failure has no effect on
operation of the pitch flight control system.
Loss of system pressure from both hydraulic systems will automatically
revert the pitch flight control system to manual control. As pressure at the
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actuator drops below 60 psi, bypass valves within the actuator open to
allow the actuator piston to idle. With the actuator piston idling, the system
is said to be in manual reversion, technically explained in the following
paragraph.
The actuator input crank and output crank rotate on a common pivot point.
They are linked by a pin and elongated slot arrangement. The input crank
is on the pin and the output crank is on the slot. With the actuator piston
idling, the pin moves within the slot until it reaches either end. At this point,
mechanical contact is made and the input crank now drives the output
crank. The output crank in turn drives the left and right elevators through its
push-pull rods and cranks.
Manual reversion of the pitch flight control system is also possible by
closing a normally open flight power shutoff valve. The flight power shutoff
valve is a mechanically operated shutoff valve located between the
Combined and Flight hydraulic system pressure sources and the elevator
actuator (as well as the aileron, rudder and flight / ground spoiler actuator)
pressure lines. The valve consists of two mechanically connected but
hydraulically isolated sections. A controlex cable connects the valve to a
FLIGHT POWER SHUT OFF handle located on the left aft side of the
cockpit center pedestal. See Figure 8.
Moving the FLIGHT POWER SHUT OFF handle up from its stowed
(horizontal) position to the vertical position mechanically closes the flight
power shutoff valve. With the valve closed, operating pressure is removed
from the actuator, allowing the piston to idle.
The resultant advantage of the flight power shutoff provision is the ability to
bypass a malfunctioning actuator (such as would be the need in the
unlikely event of an actuator jam) and manually fly the aircraft. Although
control column effort and response time to inputs are increased while in
manual reversion, the aircraft remains capable of positive and harmonious
control.
E. Pitch Trim System:
(See Figure 5 and Figure 9.)
(1) Elevator Trim Tabs:
A trim tab is installed on the trailing edge of each elevator. The tabs
are mechanically positioned through cable-driven drum actuators
located in each elevator. The trim actuators can be operated
manually or electrically as described in the following paragraphs.
Elevator trim tab travel ranges from 22 ±1°tab trailing edge down
(aircraft nose up) to 8 ±1°tab trailing edge up (aircraft nose down).
(2) Manual Trim Control:
Manual control of pitch trim is accomplished by an interconnected
manual trim control wheel set. A trim control wheel and elevator trim
scale are 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; the opposite wheel
moves in unison. With electric pitch trim engaged, both manual trim
control wheels move in unison corresponding to the amount of
electric pitch trim movement. Each elevator trim scale range is
incremented to a maximum of 22 units aircraft nose-up (22 ±1°tab
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trailing edge down) and 8 units aircraft nose-down (8 ±1°tab trailing
edge up).
Mechanical stops limit elevator trim wheel movement to 6.6 turns
from each stop. A shear rivet installed in the cockpit portion of the
system prevents application of excessive force by shearing at
approximately 31 pounds of force.
(3) Electric Pitch Trim:
Located on the pilot’sflight panel, the PITCH TRIM ENG / DISENG
switch engages or disengages the electric pitch trim. Pitch trim is
also engaged whenever the autopilot is engaged. With electric pitch
trim engaged (amber DISEN switch legend extinguished), pitch trim
can be adjusted through use of a split-half pitch trim switch
(sometimes referred to as a “beep”switch) installed on the outboard
grip of each control wheel. Switch positions are labeled NOSE
DOWN and 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 switch must be
simultaneously moved in the same direction.
Movement of the electric pitch trim switch to NOSE DOWN or NOSE
UP actuates the autopilot elevator trim servo. The trim servo is
connected to the manual trim control wheel set by a chain. The
chain-driven movement of the manual trim control wheel set in turn
positions the elevator trim tabs.
The electric pitch trim is normally checked by the flight crew on the
first flight of the day, during the Before Starting Engines checklist. A
check usually consists of running the elevator trim fully up, then fully
down, using normal methods, i.e., using both halves of the switch
simultaneously. This is followed by attempting to run the pitch trim
using each half of the switch alone. Any movement resulting from
using either half of the switch alone indicates a malfunction that
should be corrected before flight. The check is concluded by setting
pitch trim for the takeoff Center of Gravity (CG) condition as
determined using the Airplane Flight Manual.
(4) Elevator Trim Tab Actuator Heat System:
Electrically-heated elevator trim tab actuators are incorporated on
airplanes SN 1380 and subsequent and SN 1000 through 1379
havingASC 342. These actuators are designed to alleviate frozen or
stiff trim tab actuators possible in extreme cold temperatures. The
system receives power from Phase C of the Left Main AC bus.
Operation of the system is automatic and transparent to the flight
crew.
F. Angle-of-Attack / Stall Barrier System:
(See Figure 4, Figure 10 and Figure 11.)
While in flight, the Angle-of-Attack (AOA) system monitors aircraft AOA to
provide warnings of an approaching stall. If AOA continues to increase
toward aerodynamic stall, the system applies a nose down control input
through the stall barrier system.
The AOA / stall barrier system consists of the following units and
components:
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•AOA probes
•AOA display and approach indexer
•Stall warning computers
•Pilot and copilot control column shaker motors
•Stall barrier system
AOA becomes fully functional as the aircraft becomes airborne, i.e., when
the nutcracker shifts to theAIR mode. The control column shaker, however,
is disabled for the first five seconds following rotation to eliminate nuisance
activity. Pulling the SHAKER #1 and / or SHAKER #2 circuit breakers, as
appropriate, is the only way to disable a control column shaker in flight;
however, such action completely disables the associated stall warning
computer(s). System design is such that either stall warning computer is
capable of operating the control column shaker and control column pusher
should the other computer become disabled.
(1) AOA Probes:
AnAOA probe / transducer assembly is installed on the left and right
forward fuselage. The cone-shaped probes freely rotate in the
airstream to provide AOA reference data for the stall warning / stall
barrier systems and AOA display data for the flight crew. The left
AOA probe provides data the No. 1 stall warning computer while the
right AOA probe provides data the No. 2 stall warning computer.
Heating elements prevent ice accumulation on the probe and
condensation within the transducer case.
(2) AOA Display and Approach Indexer:
AOA display data supplied by the probes includes the normalized
AOA display and the approach indexer. Normalized AOA display is
shown on the lower left portion of the Primary Flight Display (PFD)
and consists of a vertical scale marked from 0.2 to 1.1 in 0.1
increments. At the bottom of the scale is a three-digit display
surrounded by a pointer that provides AOA indication within a 0.01
resolution. As AOA changes, the display / pointer moves up and
down to correspond with the indication on the scale.
An AOA approach indexer on either side of the windshield center
post indicates the optimum AOA for approach and landing. The No.
1 AOA system drives the pilot’s indexer while the No. 2 AOA system
drives the copilot’s indexer. During approach and landing, the AOA
system illuminates each indexer’s red chevron if AOA is too high, a
green circle if AOA is correct or an amber chevron if AOA is too low.
(3) Stall Warning Computers:
The No. 1 and No. 2 stall warning computers receive the following
inputs and then provide outputs to the control column shaker motors
and stall barrier system:
•AOA reference data from the associated probes
•Altitude data from the DADCs
•Nutcracker mode from the nutcracker relays
•Flaps position from the 39°flap relay
(4) Control Column Shaker Motors:
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A control column shaker motor is attached to the pilot and copilot
control columns. When activated by a stall warning computer, the
motor drives an off-center weight that vibrates the control column.
Activation of one motor affects both control columns due to their
mechanical interconnection.
(5) Stall Barrier System:
A stall barrier system (control column pusher) is incorporated in the
pitch flight control system to prevent a stall by forcing the control
columns forward when the flight crew fails to respond to either visual
indications or to the control column vibrations that warn of
impending stall. The system consists of two normally closed stall
barrier valves and an actuating cylinder that is mechanically linked
to the elevator actuator input sector. One valve receives signals
from the No. 1 stall warning computer while the other valve receives
signals from the No. 2 stall warning computer. If one system fails,
the remaining system is capable of operating the system.
When a high AOA is reached, the control column shaker motors are
activated. When a more severe AOA is reached, the control column
pusher trip detector activates its respective stall barrier valve. The
activation signal will originate from whichever system is operating -
No. 1, No. 2 or both. When a stall barrier valve is activated,
Combined (or Utility) hydraulic system pressure is ported to the
extend side of the stall barrier actuating cylinder. As the cylinder
extends, it applies an input to the elevator actuator input sector. This
input causes the elevator actuator to drive the elevator trailing edges
down; the control column drives forward accordingly, to
approximately one inch forward of neutral. When AOA has
decreased more than one degree, the stall barrier system
disengages.
The force generated by the stall barrier system is sufficient to
overcome any autopilot force, however, the system can be manually
overcome by the flight crew.
The stall barrier system can be deactivated by pressing the BARR
DISC button on either control wheel. The BARR DISC button also
serves as the autopilot disconnect button, thus is also labeled A/P
DISC accordingly. Deactivation of the stall barrier system is also
possible through selection of the STALL BARRIER switchlight to
OFF. The switchlight is located on the cockpit center pedestal just
below the left HP fuel cock. An amber OFF legend in the switchlight
will illuminate when the system is deactivated and will extinguish
when activated.
(6) Stall Warning / Stall Barrier System Test:
The stall warning / stall barrier system is normally tested by the flight
crew on the first flight of the day or every eight hours of flight time.
The test is performed only on the ground and cannot be tested in
flight. It consists of the following steps:
(a) Select the STALL BARRIER switch to on. Verify amber OFF
legend is extinguished.
(b) On both the pilot’s and copilot’s display controllers, depress
the TEST function key.
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