Source: http://code7700.com/g450_flight_controls.htm
Timestamp: 2019-04-26 14:17:21+00:00

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The G450 flight control system is a combination of what first came out in the GII with all the lessons learned with the GIV, and all the improvements through the GV and G550. But there are a few quirks because the G450 is based on the GIV which does not have as clean a wing as the GV.
The Gulfstream elevator system has matured over the years, the basic system in the GII gained a full-time 3,000 psi hydraulic boost with the advent of the GIV and hard over protection with the GV. As the system evolved, less and less has been written so much of the basic system knowledge depends on earlier Gulfstream manuals.
The aircraft has a dual elevator installation to control aircraft pitch attitude. The elevators are composed of a baked graphite-epoxy material and connected together with a "U" shaped fitting at a common hinge point. Each of the cockpit yokes is connected to both of the aircraft elevators. Each yoke is also connected to the other yoke by a mechanical torque tube beneath the cockpit floor. Since both yokes are interconnected, moving one yoke moves both elevators.
Beneath the cockpit floor, a levered attach point on the pilot side of the yoke interconnect is joined to a push rod that moves a bell crank. On the bell crank are the terminal ends of a braided steel cable that runs aft to a hydraulic assist actuator in the tail compartment of the aircraft. The cable is routed beneath the aircraft floor using pulley connections to clear other installed equipment. The cable mates with the hydraulic assist actuator via a bell crank that translates pulley rotational motion into forward and aft motion. The actuator has a single shaft powered by two piston chambers, one chamber for each (left and right) hydraulic system. Both hydraulic systems normally power the actuator, but one system is sufficient for full elevator movement. The actuator is connected to the elevator by linkages and bell cranks, moving the elevator up or down about the pivot point on the aft of the horizontal stabilizer. The deflection range of the elevator is twenty-four degrees (24°) up and thirteen degrees (13°) down.
Both sides of the hydraulic actuator 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 actuator is bypassed or lost (in the instance of dual hydraulic system failure) the elevator remains 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.
The pilot yokes are mechanically connected to the elevators with no provisions to disconnect the link. Moving the yoke will move the elevator, just as any movement of the elevator will be translated to the yoke. The only provision to stop this movement is through the control gust lock.
More about this: Gust Lock.
Pilot/copilot inputs cause rotation of the cable sector which is push rod connected to the actuator input crank. Both the actuator input crank and the output crank rotate on the same common pivot point and are linked together by a pin and elongated slot arrangement. The pin and slot arrangement permits differential motion between the two cranks until the pin bottoms at either end of the slot. The pin is attached to the actuator input crank and the elongated hole is on the output crank. The differential motion permits the input crank to move initially when an input is sensed, this motion is transmitted to the servo control valve input lever and causes the hydraulic actuator to stroke to the selected position. The actuator drives the output crank which transmits motion to the elevators through a series of connecting push rods, idlers and cranks which are routed to the top of the vertical fin. Each elevator is driven separately by individual push rods which link to a yoke-type crank at the top of the vertical fin.
Autopilot servo input to the longitudinal control system is through the cable sector. There is a separate set of cables leading from the longitudinal autopilot servo which are connected to separate sectors at this point. The autopilot servo inputs are introduced at the cable sector crank and displace the system to obtain the desired attitude about the pitch axis as called for by the autopilot.
The pin and slot arrangement is what allows hydraulic boost and manual reversion. When hydraulic pressure is available, levers connected to the pin port hydraulic pressure inside the actuator which then provides mechanical force to move the elevator. When hydraulic pressure is not available, the mechanical connection from the elevator cable, to the pulley, to the input link arm drives the pin to either end of the slot. Once the pin bottoms out at either end of the slot, mechanical movement is sent to the elevator. Early Gulfstream manuals provide drawings of this, the G450 and later do not. There is, however, a good drawing in the FlightSafety G450 Pilot Training Manual, next . . .
[G450 Aircraft Operating Manual, §2A-27-10, ¶1.] Hydraulically powered aircraft components, except the engine thrust reversers, are redundantly protected with either an alternate hydraulic power source, dual (left and right) hydraulic actuators, hydraulic accumulator pressure or compressed nitrogen (N2) bottle pressure. Control surfaces used throughout the flight regime are powered using actuators connected to both hydraulic systems, with either system capable of independently powering the controls.
Forward and aft movement of the control column moves the input crank (through the green pulley on the left) through a series of cables and pulleys. The input crank is has a pin that feeds into an elongated hole in the output crank (the pulley in the diagram above). Hydraulic servos react to the pin hitting either side of the elongated hole (see the brown lever in the diagram above) and attempt to center the pulley, providing hydraulic power to move the output crank, which is attached to the elevator through a series of push rods and cranks. If hydraulic power is lost, the pin in the input crank moves the output crank directly.
With normal hydraulic pressure, the servo is working quickly to keep that pin centered so as soon as you make an input, the servo moves and the pin never really reaches either extreme. The result is almost a seamless translation for your pitch input to elevator response. Without hydraulic pressure pilot pitch inputs must move the actuator fully to move the pin to one extreme or another before any input is translated to the elevator. Not only will control pressures be higher, they will be delayed.
More about this: G450 Manual Reversion.
Movement of the elevators contrary to the commanded position are limited by a Hard Over Prevention System (HOPS). The system incorporates eight switches 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 installed on each side of a bracket attached to the command input side of the elevator actuator. 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.
The pin and slot arrangement discussed earlier keeps the elevator linkage centered in relation to the output of the actuator. As long as it is centered the four switches shown here should be centered. If they are not centered or if they indicate movement is opposite to what the actuator is doing, a HOPS signal is generated.
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 of a second, hydraulic pressure from both left and right systems is shut off to the elevator actuator.
The four internal switches do nothing other than measure pressures which should be balanced. If pressure in one direction overwhelms the other, a HOPS signal could be generated. It is conceivable the imbalance could have been generated by rapid pilot inputs, or so they say. The abnormal procedure gives the pilot the option of resetting the HOPS by pull and resetting the ELEV HYD S/O circuit breaker.
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.
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 “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 hard over 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 ELEV HYD S/O circuit breaker.
[GV Aircraft Operating Manual, §2A-27-20 ¶2.A.] The elevator actuator incorporates a hard over 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 hard over prevention system receives power from the Left and Right Essential DC bus.
The elevator is normally mechanically controlled and hydraulically operated. The HOPS system is purely electrical and will change the equation to make the elevator mechanically controlled and either partially or totally mechanically operated. It takes electrons to make the HOPS work and if you lose those electrons through either a circuit breaker or loss of left essential DC power, the hydraulic power will return.
More about this: G450 Elevator HOPS Activation.
The elevator has two trim tabs 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 175°F 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 22° trailing edge down (aircraft nose up) to 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.
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 bell cranks to the elevator trim tabs.
The flight crew has the option of electrically moving the elevator trim tabs. A push button, 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.
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.
The only time you will do this is when you need to control the pitch trim manually. When? Read on . . .
Manual Electric Trim (MET) - This function gives direct pilot control of the trim tab surface through the manual trim switches located on each control wheel. MET is active if AP is not engaged, provided that no failures are detected within the MET system. MET overrides the Mach trim functions when the manual trim switches are pushed.
Mach Trim (MT) - This function is automatically activated in the Mach region when AP is not engaged.
Autopilot Pitch Trim (APT) -This function is activated when the autopilot is engaged to alleviate elevator servo loads by controlling the trim tab surface as a function of servo current.
The manual trim cable from the cockpit ends at a spool just aft of the electric trim servo.
This spool is on a common shaft, connected to a second spool which has another cable that goes up to the elevator.
A geared chain is connected to the spool and just forward of that is another shaft connected to two electric trim servos.
One electric trim servo is connected to the pilot electric trim switch and the other to the copilot electric trim switch. Each is powered by ESS DC and protected by a circuit breaker on either the POP or CPOP panel.
Pitch trim, how hard can it be? It really is pretty easy, you just need to know how these four things work . . .
The Nose Down/Nose Up switches on the control wheels are connected to the elevator trim tabs via an electric motor.
If you press the EMER STAB switch, the stabilizer is no longer connected to the computers and is now controlled by the NOSE Down/Nose Up switches you were using for the elevator trim.
So that's why you've got those mechanical things inboard of your legs. They are, after all, connected to the elevator trim tabs mechanically. So why aren't they working?
You need to disengage the electric trim motors before you can use the mechanical trim wheel.
The autopilots employ 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 active 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.
Mach trim must be ON during all flight operations except as provided for in Section 05-02-40, Mach Trim Failure.
[G450 Aircraft Operating Manual §2B-08-50, ¶1.A.] Mach trim is automatically engaged at power-up and only functions in the MACH region when autopilot is disengaged. Mach trim operation requires at least one micro air data computer to be valid, and no unflagged MACH misnomers between the multiple ADCs.
The Mach trim function is part of the autopilot and trims the aircraft nose-up for increasing Mach number and nose-down for decreasing Mach number during manual flight within the transonic flight region (0.85- 0.90 Mach for GV-SP, 0.80-0.93 Mach for GIV-X) by commanding elevator trim tab movement. The Mach trim function provides trim rate commands to the trim servo that drives the elevator trim tabs. The Mach trim control law with the addition of trim tab position feedback is used by the Mach trim function to control the aircraft. The rate command computed by the Mach trim function is based on air data Mach and elevator trim tab position information.
For GIV-X, the delta trim tab is applied as a function of delta Mach over the 0.80 to 0.93 transonic region is applied linearly to a maximum of 6.5°. For GV-SP, the delta trim tab is applied as a function of delta Mach over the 0.85 to 0.90 transonic region is applied linearly to a maximum of 1.7°.
All the Mach trim system does is anticipate the nose down pitch trim required by comparing air data with computed data for the wing and apply the necessary elevator trim. If the autopilot is engaged, that happens automatically through the autopilot elevator trim input to the elevator. If the autopilot is not engaged, it happens automatically from the autopilot to the elevator trim tab without any other autopilot functions.
You have to have at least one to dispatch.
There isn't a lot out there about G450 critical Mach numbers but based on the limitations, lets call it 0.75MT.
At less than critical Mach the entire wing is subsonic and the wing works just like you think it should.
As you increase the speed just above critical Mach, the forward section of the wing becomes supersonic. The faster you go, the more of the wing becomes supersonic. The boundary between supersonic and subsonic is a shock wave and and laminar flow over the wing separates and you lose lift, hence the nose down moment. A higher wing sweep improves things but a degree of Mach tuck will still exist at higher speeds.
The Mach trim system is applying nose up trim for you at higher speeds, apparently in increments small enough to be transparent to you. Gulfstream doesn't think you would be able to make these inputs in timely enough a fashion and have placed a limit of 0.75MT if the system should fail and will not let you dispatch without it.
More about this: G450 Mach Trim Failure.
More about this: High Speed Flight.
The rudder in the GII/III could kill you. It was improved somewhat in the GIV and fixed completely in the GV. The G450 rudder system comes from the GV and that is a good thing.
[G450 AOM, §2A-27-30, ¶1] The rudder is positioned by inputs from the pilot or copilot rudder pedals, or by autopilot electro-servos. The pilot and copilot rudder pedals are connected by a common torque tube so that either may control rudder movement. The common torque tube is connected by a bellcrank to a single stranded wire cable loop located on the right side of the aircraft beneath the cockpit and cabin floor. The cable loop incorporates pulleys and bellcranks to route the cable around other installations beneath the aircraft floor.
This mechanical set up is not as direct as it may seem. The input from your feet is transmitted via bellcranks and cables but end up in a bellcrank that is connected to the rudder via a pin and slot. Your inputs go to the pin and the hydraulic actuator attempts to center the slot around the pin. It is all transparent to you until you lose hydraulic pressure. In that case you need to shove that pin an extra distance to get any rudder deflection. It can be felt in your feet as a little extra "play" in the rudder. Of course the only play involved is within that elongated slot. You can see the pin and slot in the diagram which follows in the next section.
[G450 MM, §27-20-00 ¶2.D.] The right rudder cable connects to the upper part of the input sector while the left rudder cable connects to the lower side of the sector. Both the input sector and the output crank are joined at the center rear crank arm by a pin and slot arrangement. When the hydraulic actuator strokes, it drives the output crank which is a yoke-type arrangement rotating on the same axis as the sector assembly. There is a differential motion of ±4° between the sector and output crank due to the pin and slot arrangement (the slot is in the output crank and the pin is secured to the sector assembly). This differential motion is sufficient to create maximum valve displacement and to allow ±1-1/2° yaw damper authority from pilot selected position. In the event of complete hydraulic failure, the free motion of ±4° would exist between input and output until the pin bottoms in the slot to drive the output crank.
[G450 AOM, &2A-27-30 ¶1] The rudder actuator has a single shaft with a piston chamber for each (left and right) hydraulic system. Both hydraulic systems provide up to 3,000 psi pressure to assist in moving the rudder surface. Internal regulator valves limit the pressure output of the two pistons within the hydraulic actuator to a maximum of 1,500 psi. The output end of the hydraulic actuator shaft is connected to linkages that move the rudder around the pivot point connections on the vertical stabilizer. If one hydraulic system fails, the regulator valve of the remaining system shifts to provide up to three thousand (3,000) psi to move the rudder.
The rudder pedals are connected to the rudder actuator at a pin inside an elongated slot. The slot allows the actuator to apply hydraulic pressure in an attempt to keep the pin centered. It also gives the yaw damper the ability to apply up to 1-1/2° of motion beyond rudder pedal position. Hydraulic pressure is normally halved but if a single system were to fail, the other system is used at 3,000 psi. If all hydraulic pressure is lost, rudder pedal movement is transmitted directly after the 4° of the slot is overcome.
Automatic rudder compensation for aircraft yaw is provided by the yaw damper function of the autopilot. The yaw damper is normally engaged even if the autopilot is not operating. If the autopilot is engaged, the yaw damper must be engaged, since autopilot rudder commands use the yaw damper circuits to displace the rudder.
The autopilot processor detects an uncommanded yaw displacement by monitoring data from the IRUs. The amount of rudder displacement necessary is a function of airspeed / Mach number.
The yaw damper function is a redundant dual-channel installation. Since only one yaw damper channel is necessary for rudder control, the active channel alternates on each flight segment (a function of weight-on-wheels) to prolong system life. If the active channel fails, the standby channel will automatically assume yaw damper control.
The yaw damper function also provides a rudder input for aircraft turn coordination provided the flaps are not set to thirty degrees (30°) or more. The yaw damper will add up to five degrees (5°) of rudder in the direction of turn without pilot rudder input.
Autopilot control of the rudder is the same functional process as the yaw damper, but the amount of rudder displacement available to the autopilot is greater (up to the 22° limit).
Swept wing aircraft tend to oscillate in yaw and roll, usually so slightly it is imperceptible. You can see these oscillations if you pick a spot on the windscreen against a spot in the distance. The spot on the windscreen should trace a circle around the spot in the distance.
Yaw oscillations are naturally dampened by a restoring force provided by the vertical fin, but roll moments are not. Roll motions become yaw motions and can be difficult to counteract. The effect is made worse with increasing altitude, since the vertical fin has less to "grab" and becomes less effective. The effect is made worse still as the mass of the wings increases since each oscillation carries more momentum. This, then, explains the limitation. If you haven't been trained to manually counteract dutch roll — even if you have — a minimum speed gives the vertical fin more effectiveness and will keep you safe.
For more about this, see: Stability and Control.
[G450 AOM, &2A-27-30 ¶1] Mechanical stops are incorporated into the rudder mounting structure to physically limit rudder displacement to a maximum of 22° either side of neutral, although full displacement is available only at low airspeeds. As airspeed increases, the air load on the rudder surface increases proportionally. When the air load on the rudder surface equals the available hydraulic pressure output of the rudder actuator, no further rudder displacement is possible. The Monitor and Warning System (MWS) software monitors aircraft speed, angle of attack and rudder displacement to formulate an advisory message informing the flight crew when maximum rudder displacement has been reached.
The system automatically limits rudder travel to the maximum usable based on air loads, up to mechanical stops at 22° available only at low airspeeds.
[G450 AOM, &2A-27-30 ¶2.C.] The G450 does not have a separate load limiter unit, but relies instead upon MWS software to compute maximum rudder deflection for a given airspeed, matching air loads on the rudder surface with the hydraulic function of the rudder actuator. Although at low airspeeds full rudder travel of 22° is available, at higher airspeeds less rudder travel is necessary to achieve the desired amount of aircraft heading control. To avoid excessive loads, the rudder hydraulic actuator uses internal pressure switches to signal the MWS when full hydraulic pressure output of the actuator has been reached. The MWS formulates a CAS blue advisory message text of “Rudder Limit” for display on the CAS window indicating the maximum rudder hydraulic power assist condition.
[G450 MM, §27-20-00 ¶3.A.(1)] The rudder trim control wheel located on the aft end of the center console incorporates integral stops. Rotating the trim control wheel to the left or right transmits motion through a torque rod to turn the forward cable drum located beneath the cockpit floor. The trim cables wrap around the trim actuator cable drum a total of nine times, and are fastened with ball locks at both ends of the drum. Rotation of the drum about the Acme screw produces a linear shaft travel of 1.08 inches, with an ultimate axial load of 1500 pounds either under tension or compression. Since no stops are incorporated in the trim actuator shaft, travel is determined by the integral stops in the rudder trim control wheel. The rudder trim actuator is irreversible up to and including 250 cycles per second, and contains provisions that prevent the Acme screw torque from being transmitted to the rod ends. Movement of the Acme screw is applied to linkage that operates the control valve on the rudder actuator to deflect the rudder to affect trim. The rudder trim control wheel can be rotated 6-5/8 turns stop to stop. This rotation will be reflected on the trim indicator which is part of the rudder trim control wheel. The indicator is marked to indicate 10 units left and 10 units right of zero (rudder faired).
[G450 AOM, ¶2A-27-30 ¶2.B.] There is no trim tab on the rudder, rather the whole rudder panel moves in response to trim input. The rudder may be displaced up seven and a half degrees left or right with trim commands.
Rudder trim is purely mechanical: movement of the trim wheel translates directly to a physical movement of the cables and the actuator. If the rudder pedals or cables are jammed, the rudder trim cables can move the rudder independently.
Movements of the rudder contrary to the commanded position are limited by a Hard Over Prevention System (HOPS) that incorporates eight switches to monitor mechanical and hydraulic rudder operation. If a malfunction occurs and the rudder moves opposite to or further from the commanded direction, the cam arm that is attached to the output linkage of the rudder 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.
In normal conditions, all four switches sense stabilized pressures since hydraulic outputs positioning the rudder are balanced by air load pressures on the rudder surface acting against actuator pressures. When a malfunction occurs and the rudder 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 rudder movement persists for longer than one half second, hydraulic pressure from both left and right systems is shut off from the rudder actuator.
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. The rudder hydraulic actuator will continue to function using the remaining hydraulic system.
If both hydraulic systems are shut off, an amber caution message of "Rudder Hydraulics Off" is displayed. Manual rudder control may remain possible, depending upon the cause of the hard over condition. If the cause of the condition is thought to be momentary, and the use of the rudder is deemed necessary for continued safe flight and landing, the hydraulic shut off valves may be reset by cycling the RUDDER HYD S/O circuit breaker. If the cause has not been rectified, the shut off valves will close and hydraulic boost for the rudder will be unavailable. Loss of rudder hydraulic pressure will also prevent yaw damper (and autopilot rudder) operation.
If only one hydraulic system is shut off to the rudder, the remaining system will provide full boost to the rudder and yaw damper / autopilot rudder operation. The amber caution CAS message of "Rudder Hydraulics Off" will be accompanied by a blue advisory "Single Rudder" message.
Automatic hard over 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.
There are rudder HOPS switches and shut off valves for both hydraulic system inputs to the rudder. A Rudder HOPS will generate a Rudder Hydraulics Off CAS message. A HOPS can be activated for a single system in which case a Single Rudder CAS message will also be displayed. The HOPS is triggered and actuated electrically. If electrical power is lost or the rudder hydraulic shut off circuit breaker is pulled, rudder hydraulic pressure will be restored.
More about this: G450 Rudder HOPS.
[G450 MM, §27-10-00 ¶2.] Both the pilot and copilot control wheels are mechanically connected. By displacing either of the control wheels, push-pull rods, bell cranks and 3/16 inch diameter cables are utilized to operate a control valve on the aileron servo actuator. The aileron actuator mechanical input system is tied in with the flight spoiler system. The aileron actuator is cradled in the power boost linkage between input crank and output crank with the lever ratios designed for a 6.5:1 boost ratio. For every unit of work done by the pilot, the actuator does four.
I'm not sure where they are coming up with the "for every unit of work done by the pilot, the actuator does four." If anybody can figure this out, please "contact Eddie" below.
Control valve displacement is caused by a relative rotation of input and output cranks about a common pivot. The valve is ported so that the output follows input with the motion of the actuator housing providing follow-up (feedback). Relative motion between input and output levers is limited by a pin in slot arrangement to the amount necessary to provide sufficient valve displacement to meet maximum piston velocity. The free play appears on manual reversion since valve error goes to 100% when there is no hydraulic power available. The actuator assembly contains an input damper which provides a force proportional approximately to the square of input velocity. The damper ensures stability in this type of power-boosted system which feeds a portion of the output load into the input system inertia. The aileron servo actuator is a tandem arrangement of two double-acting balanced cylinders supplied by separate hydraulic systems. If all hydraulics should fail or be shut off, the primary flight controls will revert to manual operation. The actuator contains cylinder bypass valves which open when system pressure drops below 60 psi. If both hydraulic systems fail or are shut off manually, the aileron actuator piston is free to idle as the pilot controls the ailerons manually.
[G450 AOM, §2A-27-40 ¶1] The maximum aileron deflection is 11° up or down. To increase the response time for roll commands, two of the three spoiler panels on each wing are mechanically coupled to the aileron control cable linkage. The two outboard spoilers on the falling wing (with the aileron deflected up) deploy, creating additional disruption of airflow over the wing to increase roll rates. The spoiler panels on the upward moving wing (with aileron deflected downward) do not move, remaining faired with the upper wing surface.
[G450 AOM, §2A-27-40 ¶2.A.] As both yokes rotate, bell cranks translate rotational movement into linear displacement of the control cables leading to the hydraulic actuators that move the ailerons. The hydraulic actuators are located on the aft wing beam between the flaps and the ailerons. The actuators boost manual or autopilot control inputs and position the ailerons in the commanded direction. Each hydraulic actuator is powered by both hydraulic systems. The actuators have a central shaft surrounded by a dual piston chamber. Each chamber is powered by a dedicated left or right hydraulic system. The pistons are moved by hydraulic pressure ported to the extend or retract sides of the chamber by control valves moved by the aileron cable linkage. The pistons extend or retract the central shaft of the actuator that in turn moves a bell crank and linkage to position the aileron. A mechanical linkage, termed a force link, between the control cable input to the actuator and the output of the actuator to the aileron parallels actuator operation and provides the sensing element for operation of the HOPS provision. Loss of a single hydraulic system will not degrade aileron activation since the remaining system is capable of providing sufficient power for a full range of aileron displacement.
The control wheels are connected to the ailerons with cables and can, if required, manually move the ailerons. The cables end in a pin located in a slot of each actuator. The actuator attempts to keep the pin centered using hydraulic pressure. Artificial feel is provided by hydraulic bungees that dampen motion. If the hydraulic pressure fails, the pin moves the entire actuator directly, albeit requiring more force and resulting in a little "slop" as the pin moves in the slot.
[G450 AOM, §2A-27-40 ¶1] A single trim tab on the left aileron is manually positioned to reduce control yoke forces in maintaining a stabilized condition around the longitudinal axis. An aileron trim wheel on the aft section of the cockpit center pedestal is linked by a cable directly to the trim tab. The trim wheel is rotated left and right to produce the corresponding trim input to position the left aileron. Since the ailerons are linked through the crossover cable loop and the mechanical connection between the control yokes, moving the left aileron will move the right aileron in the opposite direction for a equal amount of deflection.
[G450 AOM, §2A-27-40 ¶2.E.] The aileron trim tab is a manually operated by cable linkage from the aileron trim wheel on the aft section of the cockpit center pedestal to the trim tab bell crank actuator. Moving the trim wheel to the left positions the trim tab down into the airstream, deflecting the left aileron up to reduce lift on the left wing, moving the wing downward. At the same time the connecting linkage moves the right aileron down to increase lift on the right wing and move the wing up. The range of travel for the aileron trim tab is 15° up or down.
The cockpit aileron trim wheel is purely mechanical, connected to a single aileron trim tab on the left aileron via cables.
[G450 AOM, §2A-27-40 ¶1] The autopilot is integrated into the cable linkage to the ailerons through an electric servo. The aileron cables are wrapped around a drum driven by the servo. The autopilot servo rotates the drum, producing cable inputs to position the ailerons.
Flight spoilers deploy upward on the side of the upward displaced aileron to contribute to the decrease in lift on the downward rolling wing. By interrupting additional airflow over the wing on the inside of a turn, roll rate into the turn is increased. Spoilers are activated by the same cable linkage that controls aileron deployment. A bell crank in the aileron cables on the aft wing beam provides an input to a mixing and summing series of pushrods and hinges to provide a mechanical command to the hydraulic actuator for the two outboard spoiler panels. The amount of flight spoiler extension is proportional to aileron displacement. Maximum aileron displacement is 11°, nominally 10° from neutral and maximum flight spoiler extension is 26° up from the faired position.
The GIV Aircraft Operating Manual calls this a "Simplified Drawing" of the roll control system. It is so simple, I guess, it has never appeared in newer aircraft. I like it because it explains the relationship of the ailerons to the spoilers better than the text. If an aileron goes up, the pulleys and levers to the two inboard spoilers cause those panels to go up proportionally.
[G450 AOM, §2A-27-40 ¶1] The ailerons, like the other primary flight controls, are monitored for correct operation by a Hard Over Prevention System (HOPS). The HOPS will shut off hydraulic pressure to both aileron actuators and the spoiler actuators if aileron movement does not correspond to flight crew or autopilot commands. Additionally, if a spoiler panel malfunctions and extends or remains extended contrary to commanded aileron position, a switch labelled LATERAL CONTROL can be used to shut off hydraulic power to both the spoilers and the ailerons. The switch is located adjacent to the speed brake handle on the left side of the center console. When the switch is activated, the amber “OFF” legend within the switch illuminates. If no hydraulic pressure is available to the aileron actuators, the hydraulic control valves of the actuators will bypass any residual pressure and full aileron deflection will be available using manual control inputs from the cockpit yokes. A slower response time to aileron manual operation can be expected due to the lack of boost authority and the absence of assist from the flight spoilers.
[G450 AOM, §2A-27-40 ¶2.B.] The conformity of the motion of aileron hydraulic actuators to aileron control inputs from cockpit yokes or the autopilot is monitored by force link mechanisms. The force links are the sensing elements that initiate aileron HOPS activation. The mechanisms are essentially telescoping tubes with two sections - one section that travels within the other. The end of one section of the tube is connected to the cable linkage control inputs to the hydraulic actuator and the end of the other tube section of the is connected to the actuator shaft output to the aileron. The dual tube mechanism thus parallels the action of the aileron actuator. Each tube section is spring loaded to resist the push and pull motion transmitted through the actuator to move the aileron. If a control cable input to the actuator is in the push direction and the actuator shaft does not transmit a push assist or the aileron does not move, the spring at the push end of the force link tube is compressed, and an electrical contact within the spring initiates a signal to the lateral control shutoff valves on the left and right hydraulic systems to bypass system pressure, thus deactivating both the ailerons and the spoilers. The electrical signal incorporates a one half second delay before activating the hydraulic pressure bypass. Once activated, the lateral control shutoff valves are latched to the bypass position. NOTE: When the aileron HOPS activates, hydraulic power is no longer available to operate the speedbrakes or the ground spoilers.
The aileron HOPS uses force links connected to each actuator to sense a difference in the push or pull on an actuator versus what the actuator actually does. If the difference is too great for more than a half second, it electrically fires the lateral control shutoff switch to remove hydraulic pressure from the ailerons and spoiler system. This is done with a solenoid which requires electrical power to remain closed. Pulling the valve's circuit breaker opens it and resets the logic in the system. Closing the circuit breaker will allow the system to reclose if the condition causing it close in the first place still exists.
The cockpit Lateral Control Switch is the pilot's way of triggering a HOPS, shutting off hydraulic pressure to the ailerons and spoilers via an electric solenoid. Once activated, it can only be deactivated by pulling the circuit breaker. There is no AFM procedure to use this switch.
[G450 AOM , §2A-27-50 ¶1.] The horizontal stabilizer will automatically compensate for pitch changes as the flaps extend by moving leading edge down, moderating elevator trim change and conserving elevator deflection for landing. As flaps are retracted, the stabilizer moves leading edge up to reduce elevator forces as the wing center of lift transits forward.
In normal operating mode, the horizontal stabilizer position is determined by flap position, however the stabilizer has an emergency operating mode that provides a means to position the stabilizer independent of flap position. This emergency stabilizer mode is enabled with a EMER STAB pushbutton switch. When the EMER STAB button is selected, the amber ARM legend within the switch is illuminated and stabilizer position is then controlled by the electric elevator trim switches on the cockpit yokes. Stabilizer position is monitored on the scale display shown on the left side of the HSI on the PFD and/or on the Flight Controls synoptic page.
An alternate control path is also provided for operation of the wing flaps if a malfunction occurs in the normal primary control channel. A pushbutton switch adjacent to the EMER STAB switch, labelled ALT FLAP may be selected to employ a command path from a set of position switches in the flap handle to a corresponding set of follow-up switches on the flap drive gear box.
GII, GIII, and GIV — The G450 stabilizer is electric, there is no mechanical linkage between the flaps and the stabilizer. On top of that, a computer controls everything and there is a different stabilizer position for every flap position. You must have AC power to do move the stabilizer under normal circumstances, despite what the preflight checklist tells you.
GV and G550 — The G450 has an emergency flap system that complicates the operation of the stabilizer without AC power. Do yourself a favor and ignore the AOM procedure for extending the flaps without AC power. Wait for the APU before extending flaps for the preflight.
The FSECU is located in the Baggage Compartment Electronic Equipment Rack (BEER). The FSECU is a dual processor/controller that receives external position commands for the flaps and horizontal stabilizer and also internally computes position commands for the horizontal stabilizer to coordinate stabilizer position with flap position. The FSECU responds to position commands for the flaps and stabilizer by providing control signals to the flap hydraulic motor to position the flaps and control commands to the stabilizer electronic motor control unit to move the stabilizer to a scheduled or commanded position.
Flap position commands are obtained from two RVDTs on the cockpit flap control lever. The FSECU compares the commanded position with the existing flap position derived from dual position resolvers installed on the outboard actuator of each wing flap. If there is a difference in the two positions, the FSECU signals operation of the Flap Hydraulic Control Module that controls the hydraulic motor that actuates the flaps to the commanded position.
If a malfunction such as a flap asymmetry prevents the flaps from moving, the FSECU cancels movement of the horizontal stabilizer. The horizontal stabilizer remains operative in the EMER STAB mode, and can be positioned with the cockpit yoke trim switches. If the EMER STAB mode is used to control the horizontal stabilizer, and subsequently control of the horizontal stabilizer is returned to normal operation (by deselecting the ARM state with the EMER STAB pushbutton), the FSECU will move the stabilizer to align with the existing flap setting according to the processor internal schedule. During this process, an advisory (blue) CAS message of “Stabilizer Syncing” will be displayed.
The flaps are hydraulically powered, electrically controlled. The horizontal stabilizer is electrically powered and electrically controlled. The electric control for both normally comes from a dedicated computer called the FSECU.
Your flap handle? That is nothing more than a series of electrical switches. The FSECU monitors the flap handle position versus the actual position of the flaps and horizontal stabilizer. The FSECU sends signals to the hydraulic flap motor and electrical stabilizer motor so everything agrees.
[G450 AOM , §2A-27-50 ¶1.A.] The horizontal stabilizer is an all metal structure mounted to the top of the vertical stabilizer at two points. The aft point consists of a shaft installed through an opening allowing the stabilizer to pivot. The front attachment is an actuator that moves the leading edge of the stabilizer up and down. The range of movement available for the horizontal stabilizer is from -1.00° (up limit) to -4.6° (down limit). The actuator is driven by the rotation of a torque tube within the vertical stabilizer that connects the actuator to an electric motor unit and gearbox located in the aft equipment bay.
The electric motor unit is operated by a single channel Flap / dual channel Stab Electronic Control Unit. The primary channel of the control unit is used in normal operation of the stabilizer when stabilizer position is coordinated with wing flap position. The normal control channel uses a set of motor windings powered by the Left main AC bus. The secondary channel of the control unit is used for operation of the stabilizer only independent of flap position in the emergency stabilizer mode. The secondary control channel uses a set of motor unit windings powered by the Right Standby AC bus, enabling emergency stabilizer operation during electrical malfunctions that require operation of the Hydraulic Motor Generator (HMG). Corresponding to the dual electric motor windings are two DC powered motor brakes that prevent movement of the motor drive shaft to inhibit uncommanded stabilizer movement. Both brakes are powered by the Essential DC bus. When a motor winding is employed by one of the control channels, the DC brake is first disengaged to allow the motor to turn in the desired direction. Additional protection against uncommanded stabilizer movement is provided by a locking ratchet installation, called a “no-back”, on the stabilizer actuator that prevents stabilizer motion from any source other than the stabilizer motor unit.
The horizontal stabilizer is purely electrical driven by a single motor with several windings which needs AC to run but includes DC windings to prevent it from running, an electrical brake. It is normally powered by left main AC and controlled by the FSECU, which determines stabilizer position based on flap position. It has an alternate mode designed for the day you are running on the HMG which uses right standby AC controlled by the up/down switch on each pilot's yoke when activated by the EMER STAB switch.
Two Linear Variable Displacement Transducers (LVDTs) mounted on the stabilizer provide feedback to control signals and position monitoring information.
Flap position based on feedback from the RVDTs during normal operation to move the stabilizer to compensate for flap extension.
Pilot or Copilot yoke-mounted elevator trim switches in the Emergency Stabilizer mode to provide stabilizer movement independent of flap position.
[G450 AOM , §2A-27-50 ¶1.C.] Each flap extends back from the wing along four tracks with rollers attached to the flaps moving within the tracks. The two middle track installations contain the jackscrews that extend or retract the flaps. The jackscrews on each wing are driven by torque tubes installed on the aft section of the wing running inboard to a hydraulically driven gearbox located in the main landing gear wheel well. The gearbox is turned by a fixed displacement hydraulic motor and rotates forward or backward to extend or retract the flaps. The hydraulic motor is powered by either the left or AUX hydraulic system. A shuttle valve will admit fluid to the motor from whichever system that is operating at the highest pressure (although normally the left and AUX systems are not in operation simultaneously) If the left hydraulic system is not available due to a pump malfunction, the Power Transfer Unit (PTU) can be used to pressurize left hydraulic system fluid to operate the motor.
Incorporated into the flap jackscrew actuators are force limiters that prevent structural damage to the flaps from high airspeeds. The force limiters also interrupt flap operation if one of the actuators malfunctions. If the flaps are selected to a setting greater than that allowed by airspeed limits as monitored by the Modular Avionics Units (MAUs), the overspeed condition will first be signalled by the change in color of the airspeed tape on the PFD and the OVERSPEED aural clacker will sound over cockpit speakers. If airspeeds produce high aerodynamic loads that could result in damage to the flaps, the force limiters brake the motion of the jackscrews, preventing further extension of the flaps. Once the force limiters have activated, flap extension must be reversed (selected to a lesser setting) to release the braking action of the force limiters before they may be extended again at a lower airspeed.
The flaps are driven by two hydraulic motors which are normally powered by left system hydraulics, which can be powered by the left engine hydraulic pump or the PTU. The flaps can also be driven by the AUX hydraulic pump, though at a reduced rate.
[G450 AOM , §2A-27-50 ¶1.D.] The FSECU compares flap handle RVDT with flap position resolver data to drive the flap hydraulic motor in the correct direction to extend or retract the flaps. The FSECU also positions the horizontal stabilizer according to the schedule corresponding to flap position. The FSECU processor software monitors performance of the flaps, horizontal stabilizer and flap / stabilizer scheduled positions. If a fault is detected, the operation of the malfunctioning unit (flaps or stabilizer) is interrupted. The condition is signalled through ARINC-429 connections to the MAUs, and the MWS generates the display of the appropriate CAS message(s). For instance, if the FSECU detects an asymmetry during the extension of the wing flaps, the FSECU stops operation of the flap hydraulic motor and the horizontal stabilizer electric motor and signals the condition to the MWS through the MAUs. The MWS then generates an amber “Flap Asymmetry” caution message on the CAS display.
Here's what the flaps look like going from up to 39° and back again, off the aux pump: Flaps Up-Down-Up.
[G450 AOM , §2A-27-50 ¶1.E.] If a malfunction in the FSECU or in the RVDT circuits that communicate cockpit flap handle position to the FSECU or another type of failure prevents normal flap operation, an alternate command path is provided to extend or retract the flaps. A pushbutton switch, labelled ALT FLAP, located on the center console forward of the rudder trim knob, selects the alternate flap mode. Alternate flap operation uses a set of position switches in the flap handle for a command input circuit that bypasses the FSECU. Switches are located at each of the corresponding detents in the flap handle. A matching set of switches is installed on a follow-up mechanism attached to the gear box of flap hydraulic motor. The follow-up mechanism rotates proportional to the flap position driven by the gear box as directed by the alternate solenoids of the Flap Hydraulic Control Module. When the flaps have reached the desired position, rotation of the follow-up mechanism closes the switch matching the flap position commanded by the flap handle switch, and the extend or retract signal from the Flap Hydraulic Control Module is cancelled.
Flap asymmetry protection remains available in alternate mode using the flap position resolvers and the FSECU.
The "ALT FLAP" switch would be better named the "FSECU BYPASS" switch because that is all it does. The FSECU monitors flap handle and flap position using a series of switches in the flaps and the flap gear box. If the FSECU quits for some reason and those switches are still working, pressing "ALT FLAP" on allows those switches to communicate directly.
[G450 AOM , §2A-27-50 ¶1.F.] A safety feature is incorporated into the flap / stabilizer system to interrupt any uncommanded motion of either control surface. The autopilot / stall barrier disconnect switch on each control wheel is wired into the control circuits of the flaps and stabilizer. If any uncommanded movement of either surface is encountered, depressing one or both of the autopilot / stall barrier disconnect switches will electrically interrupt the movement of both flaps and stabilizer as long as the switch is held in the depressed position. Stopping the motion of the flaps and stabilizer with the switch provides an interval to enable the flight crew to disable the malfunctioning system by pulling the applicable power circuit breakers. The motion disable function will operate in both normal and alternate modes of flap operation and for normal and emergency stabilizer operating modes.
Power levers at any setting, flaps extended more than 22° and landing gear not down.
[G450 AOM, § 2A-27-60 ¶2.] The SWPS provides the flight crew with visual indications of deteriorating airspeed and high angles of attack, a physical warning of an impending stall by activation of a yoke stick shaker and automatic elevator flight control movement by a stick pusher to decrease angle of attack. The SWPS functions are hosted in the autopilot (AFCS) processor modules of MAUs #1 and #2. Each MAU autopilot module is a fully-redundant independent processor using separate data inputs from AOA sensors, Flap/Stab Electronic Control Unit (FSECU) paths, IRSs, Air Data Systems (ADS) and Display Controllers (DCs).
Operation of the SWPS is controlled by the STALL BARR switch on the cockpit center console. The SWPS must be operative prior to commencing flight, and the switch is required to be selected on for all flight operations unless a malfunction occurs during flight.
[G450 AOM, § 2A-27-60 ¶3.A] AOA data is more accurate than airspeed in determining the onset of a stall. Angle of attack is included on the PFD, but in an abbreviated format. Two AOA sensors are installed, one on each side of the aircraft positioned below and aft of the pilot and copilot windows. The AOA sensors are powered and heated by the respective side essential DC bus (left AOA by left essential DC bus, etc.). The AOA sensors are vanes that measure the relative wind generated by the aircraft flight path. Each AOA vane is mounted on a shaft that allows the vane to rotate, aligning the vane into the relative wind. The amount of vane rotation is measured internally by the AOA sensor and transmitted over redundant ARINC-429 connections to the MAUs (the left AOA sensor has a primary ARINC-429 path to MAU #1 and a secondary ARINC-429 path to MAU #3 - the right AOA sensor has a primary ARINC-429 path to MAU #2 and a secondary ARINC-429 path to MAU #3).
The effective chord of the wing changes with flap extension and is sensed by the SWPS function through the dual paths of the Flap/Stab Electronic Control Unit (FSECU). Position of each wing flap is reported by position resolvers to the respective FSECU control channel. Each of the FSECU paths is connected over ARINC-429 links to the MAUs. The SWPS software is programmed with a set of AOA values for each flap setting adjusted for altitude, and triggers stall warning and prevention measures if the values are exceeded. To ensure that the AOA values used by the SWPS are accurate, each MAU compares inputs with the AOA data from the other MAU processor.
Since the display of actual AOA would not be meaningful to the flight crew due to the change in significance of the values with flap setting and altitude, the software in the SWPS computes a normalized value for AOA. The normalized value uses a format of zero (0) to one (1) with the indication shown to two (2) decimal places, where 0.00 indicates zero lift on the wing and 1.00 indicates the point of maximum wing lift. A normalized AOA value of 1.00 is also the threshold of stick pusher activation, since any further increase in AOA above the maximum lift angle results in a stall. The normalized AOA value is displayed on the PFD below the airspeed tape.
In normal cruise flight, your AOA should be around 0.30 for maximum endurance and perhaps as low as 0.18 for best range. On approach you may see 0.60 with an occasional transient above that. The Pitch Limit Indicator comes on at 0.70 and in gusty conditions you may see that now and then.
Learn more about this at Angle of Attack, Everybody Has One.
[G450 AOM, § 2A-27-60 ¶3.B.] As airspeed slows, the stick shaker threshold point is shown beside the airspeed tape as a red bar descending with airspeed from the shaker initiation point. If airspeed is slowing at a rate that will soon reach the shaker initiation speed, the airspeed trend vector indicator will change to amber. If airspeed reaches the stick shaker threshold, the digital airspeed readout and AOA display will change to amber. Further airspeed loss that initiates activation of the stick pusher is denoted by the digital airspeed and AOA displays changing to red.
[G450 AOM, § 2A-27-60 ¶3.C.] Each cockpit yoke is equipped with an electrically powered stick shaker installed under the flight deck floor. The motor for the pilot stick shaker is powered by the left essential DC bus; the copilot shaker is powered by the right main DC bus. The pilot yoke stick shaker is controlled by the SWPS function in MAU #1 and the copilot stick shaker by MAU #2. The SWPS software function activates the stick shakers at a threshold that provides adequate margin for pilot action to avoid a stall. The stick shaker activation point is determined within the SWPS software by comparing the current onside AOA with a threshold table value adjusted for flap setting and altitude. Except for the first five (5) seconds after takeoff, the AOA rate of change (in degrees per second) is also monitored to provide an airspeed margin prior to stick pusher activation and stall onset. Once the stick shakers have been triggered, only a decrease in AOA below the activation point will turn off the stick shakers. If a stick shaker malfunctions and activates below the AOA trigger threshold, the circuit breaker for the faulty stick shaker can be pulled (SHAKER #1 on the pilot overhead panel, SHAKER #2 on the copilot overhead panel). The MAUs monitor their respective data input paths and SWPS activation paths for operational capabilities. If faults or failures are found, the status is reported to the Monitor and Warning System (MWS) that will in turn generate appropriate CAS messages for flight crew notification.
The stick shaker activates at 0.85 AOA during the ground test, that is true, but you cannot predict its arrival while in flight. The software that is deciding to shake your stick will also make the decision to push it, so you might as well heed its warning.
[G450 AOM, § 2A-27-60 ¶3.D.] If the aircraft approaches a normalized AOA of 1.00, each SWPS channel will separately signal activation of a dedicated hydraulic solenoid valve on the stick pusher mechanism in the tail compartment. However, both AOAs must exceed the stall threshold before a stick push will occur, unless one of the AOAs has invalid data. In that case, only the SWPS channel with the valid AOA data will activate.
The stick pusher mechanism is installed just aft of the elevator hydraulic actuator, and employs a linkage connected by a cam arm to provide a nose down movement to the input side of the elevator actuator. Activation of the stick pusher will not interrupt the action of the stick shakers. The SWPS software provides an anticipatory function for the stick pusher like that of the stick shaker that incorporates AOA rate of change to prevent the aircraft from overshooting the 1.00 AOA stall point. The stick pusher will deactivate the nose down input to the elevators when aircraft acceleration drops below 0.75 G or the actual wing AOA decreases two degrees (2°) below the stick pusher activation threshold.
If a malfunction causes a spurious activation of the stick pusher, a manual force on the control column of approximately seventy-five pounds (75 lbs) will overcome the nose down control input. The stick pusher may be electrically disconnected by depressing either of the pilot or copilot yoke A/P DISC BARR DISC switches or by selecting the STALL BARR switch on the cockpit center pedestal to off (the OFF legend in the pushbutton switch will illuminate amber).
Back in the days we were doing things ruled by the thought (safe), legal, ethical — safe is optional because we were in the business of breaking things and killing people — the stick pusher was often deactivated. We would, after all, not stall the airplane.
Those days are over and we've seen many pilots over the years kill their own passengers because they overrode the stick pusher. You too can override the stick pusher, but don't. Let it get the nose down while you assess why it is doing that.
[G450 AOM, § 2A-27-60 ¶3.E.] Although the stall warning and prevention system does not operate with the aircraft on the ground (weight on wheels), a preflight test of the system may be conducted by selecting the system on the TEST menu of the Display Controllers (DCs). Each independent SWPS channel has a dedicated Line Select Key (LSK) on the DC TEST menu. Depressing the LSK next to the STALL 1 or STALL 2 legend will first activate the stick shaker, then activate the stick pusher to the nose down position. The test function of the SWPS is operational only with the aircraft on the ground to preclude inadvertent activation of the system during flight.
The SWPS test requires that the left and right AOA heat be selected ON with the pushbutton switches on the lower right section of the overhead panel. Selecting the AOA heat switches ON will provide a valid AOA indication on the PFD.
When the test function is initiated, an internal electric servo motor in the selected AOA probe will slew the probe at a high speed, approximately 40°/sec to a test start point of +13°. From the start point the AOA probe will increase angle at a slow rate of 3°/sec until reaching an upper limit of 28.5°. As the angle of the AOA probe increases, the stick shaker will first activate, followed by the stick pusher moving the control column forward. When the control column is in the nose down position, the override function may be tested by depressing the autopilot disconnect button.
Once the test is complete, the AOA probe electric servo motor will drive the probe to a down limit of minus forty-five degrees (-45°) the rapidly to an up limit of plus forty-five degrees (+45°) and returning to level or zero degrees (0°) to ensure that the probe has a full range of unimpeded operation and has not become stuck during the previous test step. During this cycle the normalized AOA indication on the PFD will be replaced with the legend “TEST” in magenta letters.
The ground spoilers have been the most feared system in traditional Gulfstreams. The safety checks in the GII and GIII were repeated just prior to takeoff and system activation in flight. All of that was automated in the GV and yet that is the only Gulfstream that has actually crashed because of inadvertent ground spoiler activation.
The G450 has the GV system with a slight change: the control signal can only be powered by the left hydraulic system or the left system powered by the PTU. The G450 AOM and AMM still have cut and paste errors claiming the control signal can be powered by the AUX pump. It cannot. If Gulfstream can't seem to understand the system what chance have we got? Well, we're going to try.
Back in the old days we would trace electrical schematics and tell ourselves we understood the system because we were smart. Of course many of those schematics had errors. Now the schematics are gone and some of the functions are controlled by software. The best we can do is to understand the ground spoilers can hurt us and we are best served by disarming them as soon as possible after takeoff and only arming them after ensuring the ground spoilers are in the air mode.
[G450 AOM, §2A-27-70 ¶1.] The speed brake and ground spoilers use six panels on the upper wing surface to increase drag and reduce lift allowing the flight crew to increase descent rates and, on landing, to shorten landing roll and increase braking effectiveness by quickly transferring aircraft weight to the main landing gear wheels. The three panels on each wing are hydraulically powered and activated by either manual command through the SPEED BRAKE handle on the pilot side of the cockpit center console, or electrically actuated as an automatic selectable feature during landing. The automatic ground spoilers will also deploy during an aborted takeoff to reduce aircraft stopping distance.
Hydraulic power for speed brake and ground spoiler actuation is provided by both the left and right hydraulic systems, with the pressure of each system reduced from 3,000 psi to 1,500 psi by a pressure control module. Two hydraulic actuators on each wing operate the spoiler panels in response to manual or electrical control inputs. Each actuator has tandem pistons, one for each hydraulic system, that move a single actuator shaft. Failure of one hydraulic system will not effect the actuators - the remaining system pressure reducer will allow a full 3,000 psi to the operational piston to power actuator shaft movement.
The G450 AOM, at this point says the ground spoilers require left or right system pressure to operate as well as control pressure from the left, PTU, or Aux systems. That is not true, in the G450 the control pressure can only come from the left or PTU systems. The AOM also says you will get a blue Single Speed Brake CAS message when down to only one system providing operating pressure. While that is true in the GV/G550, there is no evidence of such a CAS message in the G450.
[G450 AOM, §2A-27-70 ¶2.A.] All six spoiler panels will deploy upward from the wing surface in response to manual commands from the speed brake handle. The speed brake handle is connected by wire cables and linkages to the ground spoiler hydraulic actuators that position the inboard spoiler panels on each wing. The ground spoiler hydraulic actuators are in turn mechanically linked to the flight spoiler actuators on each wing that position the middle and outboard spoiler panels. The amount of spoiler panel deflection is dependent upon the position of the speed brake handle. The handle may be moved to any range from the retract position to fully extended position with no detents in the handle range.
At the fully extended position, the spoilers will deploy to 26°. The position of the SPEED BRAKE handle is monitored by MAU #2. When the handle is out of the retract position, the I/O module communicates with the MWS to generate a blue Speed Brake Extended advisory CAS message. If the power levers are positioned forward of the idle position with the speed brakes extended, an amber Speed Brake Extended caution CAS message is displayed as a reminder to the flight crew.
The two hydraulic actuators and the three spoiler panels on each wing are mechanically linked together; however, the linkage is unidirectional, operating from the inboard ground spoiler actuator outboard to the flight spoiler actuator. All speed brake or ground spoiler commands, either manual or electric, are directed to the ground spoiler actuator that mechanically relays the command to the flight spoiler actuator. The single direction of the mechanical linkage is necessary in order for the two outboard spoiler panels of each wing to assist the ailerons without operating the inboard spoiler panel. The aileron control linkage to the outer panels is a separate command input that is cumulative to the speed brake deployment command.
If the flight crew commences a roll command to the ailerons with the yoke or autopilot while the speed brake handle is extended, the two outboard spoiler panels on each wing will increase displacement over the speed brake commanded spoiler position to aid in aileron roll control. The maximum displacement of the outboard spoiler panels is 55° with full speed brake and aileron commands.
Both actuator linkages on each wing have dedicated RVDTs that furnish position information to the Modular Avionics Units (MAUs). Two additional contact switches, one on each wing, are closed when the spoilers are flush with the wing surface. The switches, located on the aft wing beam provide position signals to the MAUs confirming that the spoilers are stowed. If a malfunction of the spoiler system occurs, a single pushbutton, labelled LATERAL CONTROL, located just forward of the speed brake handle can be used to interrupt hydraulic power to the spoilers (and to the ailerons).
The flight spoilers and speed brakes are hydraulically powered and mechanically controlled; the only electrons involved are for position monitoring and warning. When you pull the speedbrake handle up, cables telegraph the motion to the inboard panel of each wing which in turn tells the outboard panels to come up as well. When you raise an aileron, the aileron actuator telegraphs the magnitude of the roll to the two outboard panels, which raise proportionally higher until they reach a mechanical limit.
[G450 AOM, §2A-27-70 ¶2.B.] The automatic operation of all six spoiler panels to reduce landing roll distance or decrease stopping distance in the event of an aborted takeoff is an electrically controlled function. Ground spoilers use no manual inputs via control cables or linkages but instead use electro-hydraulic servo valves to control the flow of hydraulic system pressure within the ground spoiler hydraulic actuators to extend the spoiler panels. The servo valves are electrically operated and use left hydraulic system pressure to control the actuators. If left system pressure is not available, PTU system pressure will operate the servo valves using left system fluid.
To provide protection against spurious activation of the ground spoilers, two levels of servo control safeguards are incorporated. The servo valves are plumbed in series and both must be open to command actuation of ground spoilers, and each servo valve has an independent electrical power circuit for operation. When the servo valves operate the ground spoiler actuators, system operation is the same as that in speed brake operation: the hydraulic actuators use both left and right hydraulic system pressure, reduced to 1,500 psi to operate the inboard spoiler panels and the mechanical linkage to the outboard flight spoiler actuators commands operation of the two outboard panels on each wing. All six wing panel spoilers extend to 55°.
For automatic operation, ground spoilers must first be armed using the GND SPLR pushbutton on the cockpit center console and the LATERAL CONTROL switch must be in the ON position.
The ground spoilers are hydraulically powered and electrically controlled. You will not have ground spoilers unless you have electrons and those electrons have to agree the airplane is on the ground. As is true with many things that can harm you, a committee is involved. This committee includes the main landing gear weight on wheel switches, the power levers, flap position, the GPWS/GND SPLR ORIDE switch, main landing gear wheel speed sensors, and the GND SPLR arm switch. What can go wrong?
Here is a diagram only an engineer could love. It comes from the G450 AOM. FlightSafety added some color to make it simple. Simple?
Left or PTU system pressure to send the "servo signal"
Either the flaps > 22° or the GPWS/GND SPLR FLAP ORIDE switch on, and both main gear > 47 knots.
[G450 AOM, §2A-27-70 ¶2.C.] When the automatic ground spoilers are armed for landing or for aborted takeoff protection, the system is monitored by the MWS for correct performance. The MWS verifies all of the control logic parameters, senses uncommanded pressurization of the ground spoiler control servos, and circuit integrity to the servos and power lever position switches. If all of the control logic parameters for the specific aircraft configuration as previously described have been satisfied, but either or both of the ground spoilers do not deploy, both red NO GND SPLRS lights located on either side of the forward windshield center post, will illuminate, a red CAS warning message of “Ground Spoiler” is displayed accompanied by illumination of the Master Warning lights on the cockpit glare shield and annunciation of the warning chime signal over the cockpit speakers.
[G450 AOM, §2A-27-70 ¶2.D.] The uncommanded deployment of a spoiler panel during flight may cause structural damage or seriously degrade the stability of the aircraft. If the SPEED BRAKE handle is in the forward stowed position and the GND SPLR switch is not armed, or if the switch is armed in preparation for landing but the aircraft is not yet on the runway, and one or both of the ground spoiler panels is not in the stowed position (sensed by the contact switches), the red “Ground Spoiler” CAS message will be displayed along with illumination of the Master Warning glare shield lights and aural warning tones. The MWS will also generate a red “Ground Spoiler” warning if conditions are detected that could cause an uncommanded spoiler activation.
Prior to the first flight of the day, a complete check of the automatic ground spoiler system is usually conducted using the test switch and other controls to verify the operation of the system activation logic.
The test procedure is ten steps long and is a rite of passage for all Gulfstream pilots. But just because you have the motions down do you really understand what you should be looking at and why? Here is the test procedure taken from the G450 AOM, §04-03-10.
Verify ground spoilers are stowed, NO GND SPLRS lights are extinguished, Ground Spoilers not displayed, Ground Spoiler Unarm CAS is displayed.
This verifies the GND SPLR switch works because the Unarm CAS is displayed with the power levers at idle and the WOW in ground mode.
Verify ground spoilers deploy, Ground Spoilers not displayed, Ground Spoiler Unarm CAS is not displayed.
This verifies the GND SPLR switch works, the primary and secondary valve work in the open position.
Verify ground spoilers stow, NO GND SPOILERS lights remain extinguished, and Ground Spoilers not displayed.
The verifies the left throttle switch works and that the primary valve does close.
This verifies the right throttle switch works.
Verify Ground Spoiler Unarm CAS is displayed.
Verify the ground spoilers remain stowed, the NO GND SPLR lights illuminate, the Master Warn lights come on with a triple chime, and the Ground Spoilers is displayed.
This electrically satisfies the conditions for ground spoiler deployment while forcing only the primary valve open, thereby testing the warning system because the spoilers did not deploy when called for and testing the secondary valve for the same reason.
Verify the ground spoilers remain stowed, the NO GND SPLR lights are extinguished, Ground Spoilers is no longer displayed, and that the warning chime ends.
Verify the Ground Spoiler Unarm CAS is displayed.
Your flight controls may be hydraulic but they have cables attached to them and they can be whipped about by the winds when the hydraulics are not pressurized. In that state, the wind can damage the flight control stops and internal components, so the gust lock is used to secure the rudder and ailerons in a neutral position and the elevator in a nose down position. The throttle is locked as well to prevent you from attempting to fly the airplane with the gust lock on.
While the gust lock is treated with hardly a few paragraphs in the Aircraft Operating Manual, it needs further research. Applying lessons from previous, non-Gulfstream, aircraft to flight control checks in your Gulfstream can be a big mistake. The worst I've heard so far: before taking the runway pulling the yoke full aft and allowing it to fall full forward in lieu of a 60 knot elevator free check. First: it is against flight manual procedure. Second: if the gust lock springs dedicated to the elevator and rudder should break, this could end up locking the elevator prior to takeoff. Yes, this scenario is extremely unlikely, but we often protect against the extremely unlikely. More about this below.
The better you understand the gust lock system, the better you will understand the need to set it when the hydraulics are depowered, the need to "unset" it before hydraulics are powered, the need to do a full flight control check prior to flight, and the need to do an "elevator free" check during the takeoff roll at 60 knots.
Unfortunately the pilot manuals we have are of little use when it comes to understanding the gust lock. I think we can piece together a better understanding with what we have at hand and maybe a little guess work too.
Engage the gust lock when leaving the aircraft without hydraulic power and test each axis and the throttles to ensure it is working.
There are a lot of poor techniques out there regarding what to do if you start the engines before disengaging the gust lock. Older Gulfstreams (GIV and earlier) had a method of removing hydraulic power from the flight controls. Some pilots would do this, disengage the gust lock, and then restore power to the flight controls. I've never seen this done but see how it can work. But it ignores the fact you could actually damage the lock, actuators, and the controls themselves. The Air Force GIII manual, Technical Order 1C-20B-1, p. 2-53 has this to say: "If engines are started with the rudder locked, hydraulic pressure surge may break the lock and cause damage to the actuator or rudder." G450 AOM, §2A-27-80, ¶2B says "FAILURE TO ALLOW HYDRAULIC PRESSURE TO DISSIPATE PRIOR TO ENGAGING THE GUST LOCK MAY CAUSE DAMAGE TO AIRPLANE STRUCTURE." We all know that. So why wouldn't the same circumstances cause damage right after engine start? You need to have the airplane inspected.
One final note, if you have any kind of speed on the tail the airloads will push up on the elevator and prevent the gust lock from disengaging. Who would be so stupid as to try this? See: Gulfstream IV N121JM. My advice: follow the checklist using a challenge-do-response method. If you find yourself with the engines running and the gust lock engaged, shut the engines down, cancel the flight, and have a Gulfstream mechanic inspect the airplane.
The flight controls gust lock secures the aircraft moveable surfaces in stabilized positions to prevent damage to the surfaces and attached control linkages from high winds or jet blast. A GUST LOCK handle on the copilot side of the cockpit center console is connected through a system of cables, pulleys and rods to hooks or latches that engage the mechanical linkages of flight controls to prevent movement. The control surfaces are locked in positions that offer the best protection to the aircraft. The ailerons and rudder are locked in neutral faired positions, and the elevators are locked in a slightly trailing edge down position so that any wind force would tend to maintain weight on the nosewheel to preserve aircraft stability.
That last statement is wrong: it locks the elevator full nose down.
You need to actively move each control surface into the lock. In our GV this wasn't done with the rudder and it spent so much time being beaten by the wind it eventually wore down the latch of the rudder gust lock to the point it didn't work.
As the GUST LOCK handle is rotated aft and up, the motion is translated by a cable, rods and pulleys to the locking mechanisms on the flight control linkages. The lateral motion of the single cable rotates pulleys that in turn move latching arms into position to engage the command input links of the flight controls. The latches will not lock in the engaged position until the flight control is placed in the specified position necessary to minimize the effect of wind forces on the aircraft.
NOTE: The gust lock is effective in preventing flight control movement during wind speeds up to sixty knots (60 kts). If weather conditions are forecast to include stronger winds, consideration should be given to securing the aircraft within a hangar or other suitable shelter.
The gust lock handle, located on the right side of the control pedestal, operates the system. The ailerons and rudder are locked in the neutral position and the elevators are locked in the 13° down position.
Each gust lock consists of a mechanical latch, springs and a bungee rod. The aileron gust lock mechanism is located at FS 283 below the cabin flooring.
The elevator and rudder gust lock mechanism is located at FS 775 in the tail compartment.
Moving the gust lock handle to the aft position transmits motion to the gust locks through a 3/32-inch diameter 7 x 7 cable. Final input to the locks is made through a spring bungee. The aileron and rudder gust locks engage the flight control system linkage when they are in the neutral position. If a surface is not in the neutral position when the gust lock handle is engaged, a cammed lead-in at the gust lock will guide the control linkage into the lock as the control reaches neutral position. (This action occurs when the controls in the cockpit initially attempt to pass through neutral with the gust lock handle in the ON position.) As the surface reaches neutral, the control linkage progresses along the cam, deflecting the spring bungee to engage the gust lock. The elevator lock detent is at the upper end of the mechanical latch and engages when the control columns are moved forward, deflecting a spring bungee.
With the gust lock handle in the OFF position, the bungee acts as a fixed rod to minimize any possibility of the locks being engaged or jammed. An additional safety feature is a set of two springs at each gust lock which will unlock the surfaces in the event of a cable disconnect.
There are indeed two springs for the rudder and elevator lock and two more for the ailerons. That should minimize the chance of inadvertent control locking if one spring fails, but not if the parts securing the springs fail. You could have a single aileron spring failed and not know it until the second spring goes. But you always have the rudder should this happen. It is absolutely critical that you inspect the rudder and elevator springs during every preflight.
NOTE: The gust lock system should not be engaged during taxi operations or ground runs, and for that reason there is an interlock between the gust lock handle and the throttle levers that restricts throttle lever movement with the gust lock system engaged.
Moving the handle aft to the ON position locks the ailerons and the rudder in the neutral position and the elevators in the 13° down position. The gust lock handle locks in the forward and aft positions. A spring latch at the lever knob must be unlocked before the control can be moved in either direction.
A mechanical interlock incorporated in the gust lock handle mechanism is a safety feature which prevents aircraft takeoff with the flight controls locked. With the gust lock handle in the ON position, movement of the throttle levers is restricted to a minimal amount. Force applied to advance either or both throttle levers cannot override the interlocks.
The gust lock cables run the length of the airplane from the handle in the cockpit to the rudder and elevator assembly in the tail. Along the way they lock the ailerons with an assembly forward of the wings that intercepts movement of the aileron cables. The throttle lock is handled by an axle and lever assembly from the gust lock lever to the throttle quadrant.
The gust lock cables are connected to two levers mounted to an axle which is common to the elevator and rudder gust locks. (Only the bottom cable is visible in this picture.) A pair of springs pull this axle from above to the off position, the bottom cable pulls it to the on position. If for any reason the cable is broken, the spring pulls the rudder and elevator locks to the off position.
The handle in the cockpit locks into place with an internal latch and places tension on the "on side" cable. This places tension on the aileron lock, elevator lock, and rudder lock. The throttle lock is activated with levers on the gust lock axle. In each case, the individual lock is not engaged until each control is manipulated into the locked position and checked for proper locking.
CAUTION: Failing to ensure each control is locked can result in damage to the lock itself. For example, if the rudder is not locked it can be whipped about left and right through the lock itself, gradually wearing it away to the point it is no longer effective. This caution does not appear in any manuals, but we've seen this in actual practice on several airplanes.
CAUTION: You should check each axis for secure locking, including the throttles. There does not appear to be a regular inspection interval for the gust lock system, so it is up to you to make sure it works.
[G450 Maintenance Manual, §27-70-00, ¶1.A.] The aileron gust lock mechanism is located at FS 283 below the cabin flooring.
The aileron lock is about the midpoint (location "K" in the drawing above and part number 355 here). The springs (part number 385 in the drawing) pull the lock off if the tension in the cable is relaxed for any reason.
The gust lock aileron latch (shown above, part number 355) is at FS 283 and appears to lock the right aileron primary cable at a pulley located at FS 280, shown in this diagram.
Figure: G450 right aileron primary cable assembly, (G450 Maintenance Manual, §27-11-07, figure 409, sheet 2.
The elevator lock is in back (location "L" in the drawing above and part number 455 here).
The ailerons and rudder are locked in the neutral position and the elevators are locked in the 13° down position.
Each gust lock consists of a mechanical latch, springs and a bungee rod.
[G450 Aircraft Operating Manual, §2A-27-20, ¶1.] The deflection range of the elevator is twenty- four degrees (24°) up and thirteen degrees (13°) down.
CAUTION: The locked position of the elevator, 13° down, is the same as its natural position with no air loads. That is, without any airspeed the weight of the elevator itself drives the elevator to the down position. If the gust lock springs are broken, vibration could conceivably move the elevator gust lock to the on position, even after a valid flight control check. A poorly conceived technique is to pull the yoke full aft and then allow it to fall forward just prior to taking the runway for takeoff. If the spring is broken this could conceivably lock the elevator just prior to takeoff. For these reasons, it is absolutely critical the elevator be checked for freedom of moment once normal air loads have been established. The AFM requires this check be done at 60 knots.
Note: our manuals do not provide any information on what to suspect in the event both these springs fail. I have given a "worse case scenario" here just to emphasize that the lock could engage itself without these springs. You really ought to check for the presence of both springs during your external preflight and you really ought to check the elevator free at 60 knots.
The rudder hook is in back (location "L" in the drawing above and part number 580 here).
ENSURE HYDRAULIC PRESSURE IS DEPLETED PRIOR TO ENGAGING GUST LOCK. IF IT IS NOT POSSIBLE TO READ HYDRAULIC PRESSURES AS THE AIRPLANE IS POWERED DOWN, CYCLE THE CONTROLS WITH THE CONTROL COLUMN, CONTROL YOKE AND RUDDER PEDALS TO DEPLETE THE RESIDUAL PRESSURE. FAILURE TO ALLOW HYDRAULIC PRESSURE TO DISSIPATE PRIOR TO ENGAGING THE GUST LOCK MAY CAUSE DAMAGE TO AIRPLANE STRUCTURE.
CAUTION: You should inspect both gust lock springs in the aft equipment compartment to ensure you do not have an inadvertent rudder lock engagement caused by vibration of this lever. There is no aircraft manual warning about this and it may not be a problem. But in the absence of more information from Gulfstream, we do not know how the rudder lock will behave in the event both springs fail.
I believe the rudder hook (part number 580 above) is meant to capture the paddle on the radius shown here.
Figure: Electrical throttle connections, (FlightSafety Maintenance Training Manual, figure 76-4.
[G450 Maintenance Manual, §27-70-00, ¶1.A.] A mechanical interlock within the cockpit center pedestal prevents advancing the power levers more than six percent (6%) forward of ground idle if the GUST LOCK handle is engaged. The blocking action of the interlock cannot be overcome with manual force to advance the power levers.
Figure: Throttle gust lock spring, from "Some guy" in tech ops.
The G450 throttle quadrant has one and only one mechanical linkage with the rest of the airplane and that is for the gust lock. While the system isn't described in any of the pilot or maintenance manuals, it appears the gust lock handle is directly linked to an axle just forward of the throttle quadrant with a torsion spring that pulls the throttle gust lock, and therefore the gust lock handle itself, to the off position if tension on the gust lock cable is lost for any reason.
Figure" Throttle gust lock internals from "Some guy" in tech ops.
An arm on the axle forward of the throttle quadrant lifts a spring lever arm which locks the throttle from inside the throttle quadrant.
Davies, D. P., Handling the Big Jets, Civil Aviation Authority, Kingsway, London, 1985.
Gulfstream GV, GV-SP, GV-SP (G550), GV-SP (G500), GIV-X, GIV-X (G450), GIV-X (G350) Master Minimum Equipment List, Revision 07, February 4, 2010.

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