Flight Control System

Gulfstream GVII

Eddie sez:


Photo: GVII flight control system, Eddie's drawing)
Click photo for a larger image

For those without fly-by-wire experience, there are some new concepts here. Even if you've dealt with non-Gulfstream fly-by-wire systems, there are some new ideas here. Let's try four approaches to understanding the GVII flight control system:

  1. A Video Primer — The training center approach to this starts with the computers and works their way out, and eventually you piece it all together. I think it might be easier to do it the other way around: start with the familiar and add the new concepts after. I've done that with three videos.

  2. An On-Screen Primer — The videos are quick, around ten-minutes each. If you want more details, I have them here.

  3. The Components — Sometimes it helps to dive deeper into a particular component. Here are a few.

  4. Once you have all that down and need a refrehser, you should look at Ivan Luciani's excellent notes on this: G500 Flight Control System. And there are also a few flash cards here: Quizlet.

Finally there are also sections on:

Everything here is from the references shown below, with a few comments in an alternate color.

Last revision:


A Video Primer

An On-Screen Primer

Learning from bullet points in the PAS might work for you, but it doesn't for me. The tempting thing to do with a fly-by-wire system is to start with the new (the computers) and work your way out to the old (the stick and rudder pedals). While doing that, you are forced to cover a lot of things that don't make a lot of sense if you don't understand the stick and rudder pedals. So my approach here will be to start with the pilot input devices in the cockpit (stick, rudder pedals), skip the computers and move on to the actuators, then cover the trim swtiches, speed brake and flap handles, and only then cover the computers. And I will do that starting with the normal stuff first, getting to increasingly complex abnormals later. So here goes.

The stick


Photo: G500 Active Control Sidestick (ACS), Eddie's drawing
Click photo for a larger image

There are a lot of aircraft out there with electronic sidesticks but, as of the G500's debut, only one other provides feedback to the pilot as well as takes pilot inputs: the Lockheed Martin F-35 Joint Strike Fighter. But for now, just realize that you move this stick left and right for ailerons, fore and aft for elevator. It does this with Rotary Variable Differential Transducers (RVDTs) that turn sidestick movements into analog electrical signals for flight control computers, more on that later. What makes this an "active" sidestick is that it also includes a set of motors that moves the stick to mimic the movements of the other sidestick in the cockpit, as well as movements induced by the autopilot and airloads on the controls.

The rudder pedals


Photo: G500 Rudder pedals, Eddie's drawing
Click photo for a larger image

The rudder pedals are mechanically linked from one side of the cockpit to the other and connected to their own sets of RVDTs to produce an analog electrical signal for the flight control computers.

The flight control actuators


Photo: G500 EHSA schematic, Eddie's drawing
Click photo for a larger image

Movement inputs from the sidesticks and rudder pedals are processed by the flight control computers and eventually end up at electro-hydraulic servo actuators (EHSA), electronic backup hydraulic actuators, and the Horizontal Stabilizer Trim System (HSTS). The EHSAs and EBHAs use aircraft left or right system hydraulic pressure to move the ailerons, spoilers, elevators, and rudder. Each EHSA includes a hydraulic reservoir to keep the actuator and its manifold filled with hydraulic pressure to prevent cavitation and to provide a damping force if the hydraulic pressure is lost. The EHSA is in "active" mode when pressurized and in "damping" mode when not.


Photo: G500 EBHA schematic, Eddie's drawing
Click photo for a larger image

An EBHA does the same thing as the EHSA unless it loses hydraulic pressure. When it has normal left or right system hydraulic pressure, it will be in active or standby mode, just like an EHSA. If it loses hyraulic pressure, it has two additional modes that take advantage of the fluid inside the manifold's reservoir. If the EBHA is attached to a surface that also has an EHSA that has good hydraulic pressure, the EBHA will go into "standby" mode, ready to take over if the opposite hydraulic system is also lost. In that case, it goes into "backup mode," whereby it uses an internal pump to pressurize the reservoir fluid to move the actuator.


Photo: GVII flight control system, Eddie's drawing
Click photo for a larger image

There are 9 EHSAs, one of each aileron, elevator, and inboard and midboard spoilers; as well as one on the rudder. There are 7 EBHAs, one on each aileron, elevator, and outboard spoilers; as well as one on the rudder. Roll control is provided through the ailerons, outboard and midboard spoilers. These spoilers combine with the inboard spoilers to provide speed brakes and ground spoilers. Yaw control is provided by the rudder. Pitch control is provided by the elevators and the horizontal stabiler, more on that later.

Between the Flight Control Computers and the Flight Control Surfaces


Photo: GVII FCC to REU to actuators and HSTS, Eddie's drawing
Click photo for a larger image

There are two Flight Control Computers each with two channels. At any given time, one channel is active and the other is on standby. Within each channels there are two lanes, a command lane and a monitor lane. Each lane has the same hardware but different software as a check. Both flight control computers communicate back and forth with eight Remote Electonics Units, REUs. Each REU communicates with two actuators, but REUs 6 and 7 also communicate with the Horizonal Stablizer. The communications from the FCCs to the actuators and Horizontal Stabilizer System (HSTS) is two way. The FCCs send what they want done, the actuators and HSTS report back what they have done.

Speed Brakes


Photo: GVII Speed brake handle, Eddie's drawing
Click photo for a larger image

There is a conventional handle connected to an RVDT which sends analog electrical signals to the FCCs, which turn this into instructions for the spoilers. (More on that below.)



Photo: GVII Flap handle, Eddie's drawing
Click photo for a larger image

There is a conventional handle with positions for UP, 10°, 20°, and DOWN. It is connected to two RVDTs and a rotary switch that sends commands to the Flaps Electronic Control Unit (FECU) which sends commands to a Power Drive Unit (PDU) which is a left system hydraulic motor that positions the flaps through a jack screw.

Horizontal Stabilizer Trim System


Photo: GVII Horizontal Stabilizer Trim System, Eddie's drawing
Click photo for a larger image

The horizontal stabilzer is positioned by one of two electrical motors which react to commands from the flight control computers through the Horizontal Stabilizer Motor Control Electrics (HSMCE). The HSMCE contains two channels, one of which is active, the other is in standby mode. They are normally controlled by an REU which has two-way communications with the FCCs. If the REUs are unavailable, trim is still available directedtly from the pedestal trim switches.

FCC Inputs


Photo: GVII Flight Control Computer inputs, Eddie's drawing
Click photo for a larger image

The flight control system uses a set of rules in software to determine what it does; these are called control laws. The usual level is called "Normal Control Law" which provides normal control feel and protections. (More about control laws: GVII Flight Control Laws. The FCCs require the following inputs to allow itself to use normal control laws:

  • Navigation: At least two IRSs or one IRS and on AHRS. There are three Inertial Reference Systems available and the system will prefer to use all three or at least two. There are two Attitude Heading Reference Systems, one of which can be used as a backup, but only if paired with an IRS. Each AHRS derives heading information from a dedicated magnetometer mounted on the tail, about halfway up. The AHRS derive attitude information from three gyroscopes, each connected to an accelerometer.
  • Air Data: There are four air data probes on the nose of the aircraft. Each data probe provides conventional pitot-static information, but also additional angle of attack and sideslip information derived from static ports above, below, and outboard of each probe. The flight control system needs information from at least two of these probes.
  • Horizontal Stabilizer Position: The flight control system needs two-way communications with the HSTS.



Photo: GVII Trim system summary, Eddie's drawing
Click photo for a larger image

There are no trim tabs on the airplane, all trim is normally handled by the flight control computers. The only exception is that the pedestal pitch trim switches can control horizontal stabilizer position in the event the HSTS loses communications with its REUs. The trim system is unlike previous Gulfstreams:

  • Side stick trim button, roll. Moving the side stick four-position trim switch left or right will induce aileron trim with a three second time limit, at which time roll trim freezes. This is part of the preflight check: hold the switch left or right for three seconds, wait for the trim to stop, then reverse switch movement and you will end up with the roll trim centered.
  • Side stick trim button, pitch. Moving the side stick four-position trim switch forward or aft will induce horizontal stabilizer movement with a three second time limit.
  • AP DISC buttons. The Autopilot Disconnect buttons function as Trim Speed Sync (TSS) buttons when the autopilot is not connected. Pressing the TSS will automatically trim the aircraft for its current speed in 1G flight.
  • Pedestal trim switches. There are two pitch trim switches on the pedestal that will induce a pitch trim only when actuated together. They do not have the a 3-second time limit like the side stick trim switches.
  • Yaw trim swtich. The yaw trim switch is spring loaded to the center and the speed at which yaw trim is applies increases with yaw trim switch deflection. If the yaw trim is not centered, moving and holding the yaw trim switch opposite the deflection will cause the trim to stop once centered.

Stabilizer / Elevator Position


Photo: GVII Control position synoptic, Eddie's drawing
Click photo for a larger image

During flight (starting once 10 feet off the ground after takeoff), the flight control computers will position the horizontal stabilizer to offset any sustained elevator deflection. (The drawing shows it independent of the elevator on the ground.)

Pitch Trim Synoptic Numbers


Photo: GVII pitch trim synoptic summary, Eddie's drawing
Click photo for a larger image

On the ground the pitch trim switches will directly move the stabilizer and the number on the synoptic to the right of the pitch trim indication reflects stabilizer position. Once airborne with the autopilot disconnected, starting when 10 feet off the ground, the number reflects the airspeed at which the horinzontal stabilizer is trimmed to hold. Once the autopilot is connected, the number disappears.

Speed Brakes and Ground Spoilers


Photo: GVII rear view, flight controls, Eddie's drawing
Click photo for a larger image

The speedbrakes are the six spoiler panels on top of the wings. Pulling the speedbrake handle to the extend position gives you as much as 30 degrees of spoiler deflection when inflight. On the ground, you can get as much as 55 degrees. They will automatically retract at high power settings, but the handle doesn't move as a result. The ground spoilers will automatically deploy the spoilers to 55 degrees and both ailerons up.

The ground spoilers are triggered:

  • Both throttles at idle and both gear WOW = ground, OR
  • One gear WOW = ground, opposite gear wheels > 47 knots, OR
  • Wheels on both gear > 47 knots, AND
    • Radio altimeter < 10 feet, OR
    • Flaps > 20° OR GPWS flap override inhibit selected

The ground spoilers will automatically stow:

  • Airspeed and wheel speed < 42 knots for 10 seconds, OR
  • Either throttle lever not at idle, OR
  • Two of the following occur:
    • Both main gear WOW = air
    • Wheel speed < 47 knots
    • Radio altimeter > 10 feet

Normal Gain

Under normal control law, the amount of control deflection in response to side stick or rudder deflection is in reverse proportion of the aircraft's airspeed. At lower speeds you will get greater control deflection in response to pilot inputs, at higher speeds you get less.

The Components

Active Control Sidesticks

The Active Control Sidesticks are designed to act like they are connected to pitch and roll control surfaces by cables and pulleys but produce a feel that changes throughout the flight envelope. You end up with an easier to fly airplane that protects you from you, or so they say. These sticks are unlike what you will find in an Airbus or Falcon; Gulfstream has learned from the mistakes of others.


Photo: Active Control Sidesticks, PAS, p. 6-22
Click photo for a larger image

[PAS, pp. 6-22 to 6-24]

  • Located on left and right side of flight deck, each contains an internal computer with 2 channels, one active, one in standby.
  • Electronically controlled to provide feel, centering, and dampening.
  • FCC’s average sidestick position inputs.
  • When sidestick motion is synchronized by cross-cockpit coupling, sidesticks are linked to each other. Input on one stick → Same motion on opposite stick.

Photo: Cross cockpit coupling, PAS, p. 6-23
Click photo for a larger image

[System Description Manual, §27-00-00, ¶

    Active Mode

  • Active mode is the normal operational mode for the sidesticks. The sidesticks simulate the feel of aerodynamic control surface loading, as well as provide feedback of the opposite pilot and autopilot inputs. The two primary differences between active mode and other conventional sidesticks are that the feedback and damping characteristics are variable, based on flight conditions and that the cross-flight deck sidestick moves to track the stick being moved by a pilot as if the sidesticks were mechanically linked. In active mode, soft stops set the deflection range of the pitch axis at 10 degrees forward and 15 degrees aft. In the roll axis, the deflection is set at 10 degrees for both inboard and outboard lateral motions.
  • In the active mode, the sidesticks provide tactile feedback to the pilot to indicate, but is not limited to the following:
    • Opposite Pilot Inputs (Cross-Flight Deck Coupling)
    • Simulated Aerodynamic Loading (Q-feel)
    • Approaching Stall Angle of Attack (Stick Shaker)
    • Autopilot Commands
    • Pitch and Roll Trim Inputs
  • When the autopilot is engaged and the sidesticks are in active mode, the sidesticks move according to the autopilot commands. In this scenario, the sidesticks are back-driven by the FCCs in order to provide a visual and tactile feedback to the pilots that the autopilot is issuing commands. Because the grips are moving, the position sensors continue to indicate the position of the grip to the FCC, but the FCC ignores these position commands during operation of the autopilot. During autopilot operation, the FCC monitors for pilot force inputs to both active sticks and accept a pilot force in any direction on either stick as an autopilot disconnect command. A trim input or autopilot disconnect pushbutton depression on either stick is accepted as an autopilot disconnect command.
  • From an operational perspective, active mode is transparent to the flight crew in that the pilot applies a force to the sidestick grip and the grip is displaced as a function of the pilot force. The position of the grip is what is actually provided to the FCC or BFCU as the pilot command. The sidesticks provide a stick back-force or Q-feel, to provide aircraft speed stability, as well as a consistent stick force per g, similar to a conventional control system. Additionally, the sidesticks provide situational awareness by simulating mechanical linkage between the sidesticks providing indication of the opposite pilot input to the control system.
  • There are not any single failures that result in loss of active mode. The primary causes for loss of active mode would be a pair of failures in both sidestick control channels. The majority of failures that result in loss of the ability of a stick to drive its motors are absorbed by having redundant motor drives, processors and power supplies. In the majority of these cases, the failure should be transparent to the pilot and the stick remains in active mode. A CAS indication is made by the FCC that sidestick maintenance action is required.
  • Degraded (or Semi-active) Mode

  • A fully functional active mode assumes the FCCs are operating in normal mode. In the event of FCC degradation, cross-flight deck coupling is expected to be available, but other active mode features are not available. This is known as semi-active mode. Semi-active mode is a sub-mode of active mode. So long as both sidesticks have cross-flight deck coupling available, the sidesticks are considered to be in active mode and the loss of other functions considered minor consequences. For example, loss of air data would degrade the FCC to its alternate mode, so soft stops and Q-feel would not be available, but cross-flight deck coupling remains in operation.
  • Passive Mode

  • Passive mode is the reversionary mode used when both sidesticks lose the resources that enable active mode. No active features would be available. Passive mode still provides the FCCs with stick position signals for pitch and roll commands, but does not provide dynamic feedback to the pilot. Passive mode feel and centering is provided by two fixed springs in each axis. Damping is provided by the unpowered SAUs. In passive mode, soft stops are not available so the deflection range of the pitch axis is 12 degrees forward and 17 degrees aft. In the roll axis, the deflection is 12 degrees for both inboard and outboard motions.
  • With one or both sticks in a passive state, the FCC sums stick inputs when the autopilot is off. The passive sticks still provide handling qualities resembling normal operation. In the event both pilots make stick inputs, the unexpected airplane response would alert both pilots that both were trying to fly the airplane. Passive mode is a no-dispatch condition if it occurs on the ground. In flight, the flight is continued to the intended destination.
  • In the event that a failure is encountered on a single sidestick that results in a loss of the active mode resources, that sidestick reverts to passive mode. The opposite sidestick is capable of operation in active mode. This allows the SAU to continue providing Q-feel to the active sidestick, but cannot allow the sticks to be linked to each other. The pilot with the active stick would be expected to assume the pilot flying duties.

AP DISC / TSS button

That little red button appears to do three things in this airplane: disconnect the autopilot, stop runaway trim, and trim for 1G flight.


Photo: AP DISC / TSS button, PAS, p. 6-27
Click photo for a larger image

[PAS, pp. 6-22 to 6-27]

  • AP DISC disengages the autopilot on the first press and silences the autopilot disengagement tone on the second.(Too much pressure on the sidestick also disengages the autopilot.
  • AP DISC also stops runaway trim on all three axis.
  • TSS (Trim Speed Sync) is available with the autopilot off: trims pitch for 1G at current speed. (Should not be used in a level turn.)

Backup Flight Control Unit (BFCU)

The Backup Flight Control Unit (BFCU) is more than just a backup Flight Control Computer, it includes backup Rotary Variable Differential Transducers (RVDTs), electrical power source, and actuators. You lose a number of capabilities, but you will still be flying.

[System Description Manual, §27-00-00, ¶

  • The BFCU is a separate part of the flight control electronics and provides continued safe flight and landing of the aircraft in the event of failure or non-availability of all four FCC channels, even though such a condition would be extremely improbable. The BFCU is located below the cabin floor just aft of the main entrance door. The BFCU is powered by 28 Vdc, derived from the uninterruptable power supply bus.
  • In the extremely improbable event that all four FCC channels lose the capability to provide continued safe flight and landing functionality, the BFCU automatically activates to support minimum aircraft control. This is known as the backup mode of flight control operation.
  • The BFCU interfaces with five dedicated Rotary Variable Differential Transducers (RVDTs): two per pilot sidestick, two per copilot sidestick and one for the rudder pedal assembly. It has seven dedicated standard ARINC 429 bus outputs for communication with the REUs associated with the EBHA of the ailerons, elevators, rudder and outboard spoiler panels. There is no feedback from the REUs to the BFCU. The BFCU does not support control of the mid or inboard spoiler panels and there is no ground spoiler, speed brake, roll or rudder trim capability. Pitch trim is provided by the Pitch Trim switches on the flight deck center pedestal.
  • The BFCU receives one input from inertial reference unit 3 input for the yaw damping function and several discrete signals (FCC, WOW, gear and flap position) used for high / low gain scheduling and activation logic. The BFCU also provides its engagement status through a discrete output that is relayed to the modular avionics unit by the Data Concentration Network (DCN) (remote data concentrator 18) to be used in the logic that sets the CAS message.

[System Description Manual, §27-00-00, ¶ -

The System Description Manual includes a diagram that says Backup Mode is entered for a "Quad-FCC Hard Failure" or "Loss of Min Flightdeck Inputs." To the first point there is a note that says "Hardware Reversion to BFCU, i.e. all FCC channels failed" and to the second there is this, "Software Reversion to BFCU, i.e. loss of minimum pitch or roll control inputs." Neither of these points are explained in the text.

  • The BFCU has three operational states: disarmed, armed and engaged.
  • Disarmed. At power up on the ground, the BFCU enters the disarmed state and cannot enter the armed state before having seen at least three FCC channels as valid (minimum dispatch condition). In the disarmed state, the BFCU transmits a discrete to the REUs. This discrete keeps the REUs and EBHAs from being commanded to center the flight controls, which could pose a safety hazard for personnel around the flight control surfaces.
  • Armed. The BFCU enters the armed state on the ground when a discrete is transmitted from three or more FCC channels, indicating that the FCCs are in their operational mode. The BFCU enters the armed state if it is powered-up in flight following a power transient. The armed state corresponds to the state that the BFCU is in for most of the ground and flight operation. After the BFCU has entered the armed state, it cannot disarm until power is removed from the aircraft.
  • Engaged. The BFCU enters the engaged state from the armed state if the discrete signals from all four FCC channels are lost, indicating that all FCC channels are failed. In the engaged state, the BFCU transmits valid data to the REUs. At this point, the REUs select the BFCU inputs only if they no longer receive valid FCC commands. This provides a double protection against inadvertent BFCU engagement so that no mixing of FCC and BFCU commands can occur. Once the BFCU is in the engaged state during flight, it cannot relinquish control even if the FCCs come back on-line. The only time the BFCU can switch from the engaged state back to the armed state is on the ground and at least one of the FCC channels re-instate their discrete. This on-ground interlock prevents any possibility of reversion from BFCU control of the flight control actuators to FCC control in flight.
  • The BFCU implements control laws that are somewhat similar to the control laws used for the FCC alternate and direct modes. The BFCU control laws use an average of the pilot and copilot sidestick position for pitch and roll axis commands. The BFCU commands consist of high and low gain settings based on the position of the flap and gear handles for all three axes, with rate limited changes when transitioning between the two gains. The low gain setting corresponds to a cruise configuration when the flaps and landing gear are fully retracted. The high gain setting corresponds to the approach configuration when either the flaps or the landing gear are extended. The rate limiting when changing between the high and low gain is to prevent abrupt changes in the response of the flight control surfaces. Additional functionality is provided in the rudder law that consists of a yaw damping function using yaw rate data from a single inertial reference unit.
  • The BFCU does not require any electronic rigging when installed in the aircraft. The cockpit control RVDTs that provide inputs to the BFCU are pre-rigged such that no electronic rigging correction is required. The actuator rigging corrections are stored directly in the REUs.

Photo: G500 Backup Flight Control Unit (BFCU), PAS, p. 6-9
Click photo for a larger image

[PAS, p. 6-9]

  • Designed for “get home capability” if both FCCs fail: Chance of that happening → 1 in 1 billion per flight hour
  • Powered by UPS bus
  • Has own set of RVDTs for sidesticks and rudder pedals
  • Outboard spoiler roll control is provided; Speedbrake and ground spoiler functions are inop
  • Once active
    • Communicates directly with REUs that control EBHAs
      • EBHAs not in electrical backup mode in this case; Unless normal hyd power not available
      • Utilizes separate, single-direction backup data buses, so REUs can’t communicate back to the cockpit
      • Actuators and surface positions unknown thus not displayed
      • Synoptic indications
        • Control surface positions black and neutral
        • Non-EBHA REUs amber
        • Inop spoilers and trim displayed with amber X
    • Active for duration of flight (can’t be reset to FCC ops)
    • EBHAs inop < 47 kts

Flap Electronic Control Unit (FECU)

You can think of the Flap Electronic Control Unit (FECU) as the Flight Control Computer for the flaps.


Photo: G500 FECU AMM, §27-55-02, fig. 3
Click photo for a larger image

[PAS, p. 6-37]

  • 2 power sources for redundancy: Left and Right Ess DC
  • Receives info from 2 Flap Handle RVDT’s and a rotary switch (Provides redundant info to FECU to determine valid flap commands)
  • FECU has command and monitor lanes: Command lane sends info from FECU to Hydraulic Control Module (HCM). Monitor lane monitors commanded flap position, Flap position, flap speed, and the direction of movement.
  • Command and Monitor lanes must agree on commanded flap position before flaps are allowed to move.
  • Flap malfunctions (jam, asymmetry, runaway, etc.) will interrupts flap motion by stopping hydraulic pressure at the HCM.

Flap Handle


Photo: G500 Flap handle, PAS, p. 6-36
Click photo for a larger image

[PAS, p. 6-36] The "flap control module" is located on the center pedestal aft of the throttle quadrant. It electrically controls the Flap Electronic Control Unit (FECU).

Speedbrake Handle


Photo: G500 Speed brake handle, PAS, p. 6-40
Click photo for a larger image

[PAS, p. 6-40] The speed brakes are electrically controlled.


The flaps are controlled by a handle that is little more than a switch and an RVDT that sends electrons to a computer (FECU) which send signals to a hydraulic motor (PDU). Everything is monitored and there is no "backup" or "alternate" as found on the GIV and earlier.


Photo: G500 Flaps hydraulic control module, PAS, p. 6-37
Click photo for a larger image

[PAS, pp. 6-36 to 6-38]

  • Fowler type flaps track rearward and downward, mounted in four flap tracks attached to rear wing spar.
  • A Flap Electronic Control Unit (FECU) receives information from 2 Flap Handle RVDT and a rotary switch.
  • The FECU has control and monitor lanes which prevent flap movement in the event of a disagreement.
  • A Hydraulic Control Module (HCM) provides hydraulic system pressure to a Power Drive Unit (PDU) as commanded by the FECU.
  • Hydraulically powered by left, PTU, or Aux systems.
  • The PDU turn flap actuators through torque tubes to position flaps.
  • The elevator is biased trailing edge up when flaps are extended to compensate for a nose-down pitching moment.

Photo: G500 Flaps indications, PAS, p. 6-39
Click photo for a larger image

Flight Control Computers (FCCs)

The Flight Control Computers (FCCs) take in all the inputs (including what your hands and feet are doing, what the atmosphere is doing, and what the various flight control components are doing) and output instructions to those components and report back to you what is going on. Those magical Flight Control Laws are nothing more than a list of rules the FCC follows under various conditions. Each FCC includes a lot of redundancy and backup systems, and there are two of them. If they both fail, you also have a Backup Flight Control Unit (BFCU).

[PAS, p. 6-4]

  • FCC #1 located in the LEER
  • FCC #2 located in the REER
  • Backup Flight Control Unit (BFCU) located under cabin floor just aft of EERs

Photo: G500 Flight Control Computers (FCCs), PAS, p. 6-7
Click photo for a larger image


[PAS, p. 6-7]

  • Each FCC contains
    • 2 Channels (A and B)
    • Cooling fan between channels
  • Power Sources (3)
    • UPS → FCC 1A & FCC 2B
    • Left Ess → FCC 1B
    • Right Ess → FCC 2A
    • UPS → BFCU

    The reason you can dispatch with 1A or 2B inoperative is that they share the same power source (the UPS) and the rule is you need 3 independent power sources.

  • FCC Channels (4)
    • Contain the Control Laws (CLAWS)
      • Programming for each flight control surface
      • Housed within each FCC channel
      • Normally one channel from each FCC sends inputs to REU; REU averages the 2 inputs then commands its actuator
      • If needed, any single channel can control all flight controls; Quadruple redundancy
    • Each channel contains 2 lanes
      • Command Lane; Sends position commands to:
        • Control surface actuators
        • Monitor Lane
      • Monitor Lane
        • Performs same computation as Command Lane
        • Compares calculations to Command Lanes
      • Each lane utilizes
        • Same type of hardware (A or B)
        • Different software; Act as self-checking pair for error detection
    • FCC 2/3 Synoptic indications
      • Green = Powered
      • Green and boxed Green = Active
      • Green and boxed Gray = Standby
      • White = Inactive

Flight Control Actuators

There are 9 actuators for normal operations called Electro-Hydraulic Servo Actuators (EHSAs) that take signals from an REU and fluid from either the left or right hydraulic systems and turn that into a physical movement of the associated aileron (2), elevator (2), rudder (1), and midboard and inboard spoilers (4).

There are also 7 backup actuators for abnormal operations called Electrical Backup Hydraulic Actuators (EBHAs) to drive the ailerons (2), elevator (2), rudder (1), and the outboard spoilers (2). The EBHAs can use left or right hydraulic system pressure, if available. If not available, the EBHA has a pump that uses trapped hydraulic fluid to provide its own pressure. The EBHA can use the associated REU but if that has failed, it has a Electric Backu Motor Control Electronic (EBMCE) that receives FCC commands but does not report back control surface position.

Here is a drawing of an aileron EBHA, but it looks identical to the EHSA to my eye. There are some differences between these and those used for the elevators, rudder, and spoilers. But we can look at the aileron EBHA and EHSA to get an idea of how they work.


Photo: G500 Aileron EBHA, System Description Manual, §27-00-00, fig. 24
Click photo for a larger image

[System Description Manual, §27-00-00, ¶] In normal operation, the FCCs transmit the commanded position on digital data buses to the REUs located near each actuator. The REUs command the electrically controlled hydraulic actuators that move the aileron surfaces. There are two separate and independent hydraulic actuators at each aileron surface. The inboard actuator is the EHSA and the outboard is the EBHA. [. . .] If both left and right hydraulic systems fail, left and right outboard aileron actuators maintain control through the EBHA. The Electric Backup Motor Control Electronic (EBMCE) in the EBHA activates a hydraulic pump that draws fluid from a self-contained compensator reservoir (manifold). The fluid allows the actuator to continue functioning as it did when fluid was provided by the aircraft hydraulic system.


[System Description Manual, §27-00-00, ¶] The EHSA manifolds contain an Electro-Hydraulic Servo Valve (EHSV) and a Mode Select Valve (MSV) which receive signals from the controlling REU. The actuator is a dual sided actuating cylinder with piston movement and velocity controlled by the EHSV. The EHSV uses commands from the REU to route hydraulic pressure to the extend or retract chambers of the associated actuator. The MSV is controlled by a solenoid-operated pilot valve and is used to switch the actuator from the active state to the damped state. A single Linear Variable Differential Transducer (LVDT) connected to the actuator ram reports actuator position information back to the controlling REU. The EHSAs have two states of operation, active and damped. During normal operation with hydraulic pressure and a controlling REU available, the EHSA is in active state. The supply and return pressure sensors allow the REU to determine if there is sufficient hydraulic pressure available to operate the actuator.

[System Description Manual, §27-00-00, ¶] In the active state, signals from the REU energizes the MSV to allow hydraulic pressure from the EHSV to be routed to the actuator cylinder. The EHSV receives electronic signals from the REU to extend or retract the actuator and control actuator velocity based on signals from the FCC. A LVDT attached to the EHSV reports EHSV position back to the REU. A separate LVDT internal to the actuator ram provides actuator position and velocity back to the REU which relays the signal back to the FCC. A bi-directional relief valve relieves any hydraulic over-pressurization within the manifold.

[System Description Manual, §27-00-00, ¶] The EHSA system is placed in damped bypass state when MSV is de-energized. Damped bypass state is entered when an REU or aircraft hydraulic system failure occurs. Since actuator is no longer actively controlled, a damping force is used for flutter suppression. The adjacent EBHA actuator is also actively controlling the same surface and overcomes the EHSA actuator damping force. In EHSA damped bypass state, actuator is commanded to shut down through a solenoid valve driving the MSV. The REU de-energizes the three-way MSV shutoff valve that moves the MSV so that cylinder ports are disconnected from EHSV and interconnected through Variable Damping Orifice (VDO). Upon loss of electrical or hydraulic power actuator MSV automatically reverts to damped state.


[System Description Manual, §27-00-00, ¶] The electric backup hydraulic actuator can operate in the active or damped states, similar to the EHSA, but can operate in the electric backup state in the event of a dual hydraulic system or REU failure. Active State: The actuator is a dual sided actuating cylinder with piston movement and velocity controlled by the EHSV. The EHSV ports hydraulic pressure through the energized Power Select Valve (PSV) and MSV to the extend or retract side of the actuator. The movement of the actuator piston is reported by two LVDTs internal to the actuator body. Damped State: If aircraft supplied hydraulic pressure to the EB manifold is lost, detected by an internal inlet pressure sensor or the controlling REU fails, the EBHA reverts to damped state. With the EBHA in the damped state, the flight control surface is driven by the EHSA. In the event of a dual hydraulic system failure, the EBHA reverts to electric backup state where the EHSV and PSV in the EB manifold as well as the MSV in the EBHA actuator are de-energized. The Integrated Motor Pump Assembly (IMPA) supplies hydraulic pressure through the pump up circuit check valves to reposition the MSV and allow pressure to retract or extend the actuator and operate the control surface. Directional and velocity commands for the IMPA are sent by the REU through the EBMCE or by the FCC through the EBMCE if the REU is not available.

[System Description Manual, §27-00-00, ¶] Internally, the EB manifold contains the following components: Electro-Hydraulic Servo Valve (EHSV), Solenoid Operated Power Select Valve (PSV), Supply and Return Pressure Sensors, Temperature Transducer, and Integrated Motor Pump Assembly (IMPA). During all states of operation, the EB manifold is commanded by its electrical interface with the EBMCE and REU. During normal operation, aircraft hydraulic pressure is controlled by the EHSV to extend or retract the actuator. During the electrical backup state, the IMPA generates hydraulic pressure to extend or retract the actuator based on the direction of motor rotation. [. . .] The EBMCE is mounted near the EB manifold. Each EBMCE has its own 70 amp circuit breaker on the 28 Vdc EBHA Power distribution box. The EBMCE uses the power for its internal electronics and to power the IMPA.

[PAS, pp. 6-5 to 6-6]

  • Two actuators for each primary flight control surface
    • Aileron (4)
    • Elevator (4)
    • Rudder (2)
  • One actuator for each spoiler panel (6)
  • Two types:
    • (9) Electro-Hydraulic Servo Actuators (EHSA’s)
      • One for each primary flight control surface (5)
      • One for each Inboard and Midboard spoiler panel (4)
      • Uses Left or Right Hydraulic System pressure
      • Commanded by an REU
    • (7) Electrical Backup Hydraulic Actuators (EBHA’s)
      • One for each primary flight control surface (5)
      • One for each outboard (multifunction) spoiler (2)
      • Normally uses left or right hydraulic system pressure
      • Normally commanded by an REU
      • If normal hydraulic pressure not available
        • Reverts to Electric Backup (EB) mode
        • Utilizes electric power to drive a pump at the actuator
          • Pressurizes trapped hydraulic fluid
          • Acts as a third hydraulic system
      • Each EBHA has a EBMCE with dual roles for that actuator
        • Powers the pump for trapped hyd fluid
        • Backup REU
          • Able to receive FCC commands
          • Unable to report back the control surface position
  • Normal operation of primary flight control surface actuators
    • EHSA & EBHA on each single surface
      • Active
      • Powered by their respective Left or Right hydraulic system
    • REU’s compare commanded to actual surface position; If error detected → New command sent to correct error
  • Actuators operate in one of the following modes or states
    • EHSA’s
      • Active
      • Damped Bypass
    • EBHA’s
      • Active
      • Damped Bypass
      • Electric Backup (EB)
  • Active Mode
    • Normal state of operation
    • Powered by Left or Right Hydraulic System pressure
  • Damped Bypass Mode
    • If actuator fails (hydraulic loss, REU fail, etc): Reverts to a damped condition; Hyd pressure trapped within actuator; Suppresses surface flutter in flight
    • For surfaces with dual actuators (primary flight controls): Damped actuator will passively follow the working actuator
  • EB Mode
    • EBHA’s only
    • When normal hydraulic pressure not available for that surface; Utilizes electric power to drive a motor pump at the actuator
      • Pressurizes trapped hydraulic fluid
      • Acts as a third hydraulic system

Flight Control System Batteries


Photo: G500 FCS Batteries, PAS, p. 6-45 and 6-46
Click photo for a larger image

[PAS, p. 6-45]

  • Electric Battery Hydraulic Actuators (EBHA) battery powers Electric Backup Hydraulic Actuators' Motor Control Electronics (MCEs).
  • UPS Battery powers FCC Channels 1A and 2B, Backup Flight Control Unit (BFCU), and provides a secondary power source for Remote Electronic Units (REU’s).
  • Select ON in any order (Checks volts per Airplane Power-Up checklist, powers actuators before the computers, initiates a System Power-On Self-Test; takes about 45 seconds.

Horizontal Stabilizer Trim Actuator (HSTA)

The Horizontal Stabilizer Trim Actuator (HSTA) is a conventional jack screw connected to two motors which are controlled by Motor Control Electronics (MCE).

[System Description Manual, §27-0-01, ¶1.1.4]

  • The aircraft pitch trim control is accomplished through a conventional movable horizontal stabilizer installed in the tail of the aircraft. The Horizontal Stabilizer Trim System (HSTS) is electronically controlled and electrically powered.
  • The HSTS electronic control is integrated with the primary flight control system in that some of the REUs are common to the primary flight control system and HSTS and the FCCs control both systems over the same data buses.
  • The Horizontal Stabilizer Trim Actuator (HSTA) provides the functions of positioning the stabilizer surface, providing position feedback, protection of the mechanical system and aircraft structure from overload and irreversibility under load. The actuator is dual load path, concentric design, where a normally unloaded secondary load path is contained within the primary load path ballscrew. The horizontal stabilizer surface position is determined by the position of the actuator.
  • The HSTA is electrically driven by dual redundant motors, which is controlled by a dual channel HSMCE. The HSMCE operates in an active-standby configuration, i.e., only one HSMCE channel and one HSTA motor is active at a given time.
  • The pitch trim input is provided through a switch in either pilot or copilot ACS or through the pitch trim switch in the center pedestal. All switch inputs is fed into the FCCs. The pitch trim switch in the center pedestal provides input to the HSMCE.
  • While the primary flight control system is in normal mode and in-air, these switch inputs are fed into the FCCs and used by the fly-by-wire control law to move the horizontal stabilizer and elevators. When the primary flight control system is in alternate or direct mode operation, these switch inputs move the stabilizer only (through the FCC). In the unlikely event of loss of all valid inputs from all FCCs or loss of all associated REUs, the HSMCE provides trim control capability using the trim inputs from the center pedestal.
  • Horizontal Stabilizer Trim Actuator

  • The HSTA has two electromechanical channels interfacing with the HSMCE that are intended to be operated in active / standby arrangement. The HSTA motors are arranged in a tandem configuration (two motor windings on a single shaft) with a separate dual channel brake located downstream of the slip-clutch.
  • The HSTA provides the functions of positioning the stabilizer surface, providing position feedback, protection of the mechanical system and aircraft structure from overload and irreversibility when under aerodynamic load.
  • The electric limit of the stab is 0.5° -0/+.25° leading edge up and 10.5° -0/+.25° leading edge down. Mechanical stops of the HSTA are 1.0° leading edge up and 11.0° leading edge down.
  • The HSTA incorporates a primary and a secondary structural load path between the aircraft fixed vertical stabilizer structure and the moveable horizontal stabilizer leading edge. Each load path independently interfaces to the aircraft through matching redundant structure. The secondary load path is unloaded in normal operation. In the event of an HSTA primary load path failure, the HSTA detects the failure as an actuator jam and air loads are transferred to the secondary load path (failsafe).
  • The lower HSTA body (fixed portion), primary load path, interface is of the universal gimbal type providing degrees of freedom in both the fore-aft direction and the lateral direction.
  • The lower HSTA secondary load path interface is through a clevis arrangement and spherical type self-aligning bearing on the structural side. This arrangement also provides the same degrees of freedom as for the primary load path attachments.
  • Horizontal Stabilizer Trim Actuator Motor

  • Each motor is a four-pole, slotted, three-phase, wye-wound, permanent magnet dc brushless motor with commutation resolver. The major subcomponents include the stator assembly, rotor assembly, commutation resolver and bearings. The motor is supplied with trapezoidal commutated three-phase 270 Vdc power by the HSMCE. The two motors share a common rotor shaft.
  • Horizontal Stabilizer RVDT Position Sensor

  • Two rotary variable differential transducer position sensors are incorporated into the HSTA system and provide actuator position feedback to the controlling REUs. The sensors are independent, dual channel, brushless sensors. The output of each HSTA position resolver is compared during initial power up to verify agreement between the channels of the HSMCE. Once the initial power up is complete and the active channel of the HSMCE is defined, the resolver outputs are not compared during the operation of the aircraft.
  • Horizontal Stabilizer Trim Actuator Brake Assembly

  • The brake is located in the gear train (not on motor) downstream of the slip clutch and is sized to hold aerodynamic loads. The brake has dual coils; one each is driven from each HSMCE channel. It only requires one coil to be energized to activate (disengage) the brake. The brake is a power-off spring-loaded design and is released during normal flight to minimize cycling, to reduce wear and tear, to improve HSTA response time and enable constant monitoring of no-backs.
  • Horizontal Stabilizer Motor Control Electronics

  • The HSMCE is a dual channel (active / standby) unit that provides power and velocity control for the HSTA motor. Each channel of the HSMCE is powered by a 115 Vac bus. Channel one of the HSMCE is powered by the 115 Vac emergency ac bus. Channel two is powered by the 115 Vac right main ac bus. The HSMCE is located in the tail compartment.
  • The two HSMCE channels are physically isolated and separated by a metal plate, but a common enclosure is used to reduce weight and volume. Each channel takes commands from a REU and drives one winding of the HSTA motor. Each channel provides its respective HSTA motor winding with pulse width modulated 270 Vdc through an internal transformer rectifier unit.
  • A dual channel resolver connected to the motor is also excited and demodulated by each channel of the HSMCE for closed loop motor velocity control. Channel one of the HSMCE provides a discrete signal to RIU 53 whenever the HSTS is in secondary mode.
  • Normal Mode

  • When there are no faults, the HSTS operates in normal mode and the pitch trim commands are sent from the FCCs to the REUs. REU-6 and REU-7 control HSMCE channel 2 and channel 1, respectively. In turn, HSMCE channels 1 and 2 control HSTA channels 1 and 2, respectively. The FCC is operating in normal, alternate or direct mode while the HSTS is in normal mode. In HSTS normal mode, the FCC generates pitch trim commands based on an automatic elevator off-load function, sidestick pitch trim input or pedestal pitch trim input.
  • Secondary Mode

  • If both REU-6 and REU-7 lose valid communication with both FCCs, the HSTS operates in Secondary Mode. In this mode the pedestal pitch trim inputs are used to control the horizontal stabilizer.

Photo: G500 Horizontal Stabilizer Control, Eddie's drawing
Click photo for a larger image

[PAS, p. 6-33] Horizontal Stabilizer Trim System (HSTS)

  • Horizontal Stabilizer Trim Actuator (HSTA) is located in the vertical stabilizer and has dual and identical electric motor-brake assemblies. One motor is capable of full HSTA performance. One motor is active while the other is in standby mode. The motor moves a jack screw → Moves the stab control surface.
  • Located on top of vertical stabilizer, fully trimmable
  • Horizontal Stabilizer Trim Actuator (HSTA) located in the vertical stabilizer, two motors (one active, one standby), moves a jack screw.
  • Trim accomplished via switches on either Active Control Sidestick or the pitch trim switch on the pedestal.
  • Under normal conditions the pilot only trims the stabilizer on the ground
  • The stabilizer automatically offset any persistent elevator offset. The stabilizer and elevators move simultaneously in opposite directions with a rate depending on airspeed. The stabilizer moves to a new trim position while the elevator moves to a "faired" position.
  • FCCs command stabilizer to 0° position after landing (approximately 20 seconds after ground spoiler retraction, 10 seconds after speed drops below 42 knots.
  • Takeoff pitch trim is automatically set if FMS PERF and Takeoff Init done prior to Initiating FCS test.

Photo: G500 Horizontal Stabilizer Synoptics, PAS, p. 6-35
Click photo for a larger image

Remote Electronics Units (REUs)

There are seven Remote Electronics Units (REU) located between the FCC and the associated actuator, and an eighth REU between the FCC and the Horizontal Stabilizer Trim System (HSTS). Each REU translates FCC signals into control movement, monitors that movement and reports back to the FCC, and disables actuation channels if anomalies occur.


Photo: Remote Electronic Units (REUs), GVII-G500 MM, §27-04-01, figure 6
Click photo for a larger image

[PAS, p. 6-3] 8 Remote Electronics Units (REUs)

  • Multichannel
  • Provide control and monitoring of Primary Flight Control Actuation System (PFCAS) and Horizontal Stabilizer Trim System (HSTS)

Photo: Remote Electronic Units (REUs), PAS, p. 6-4
Click photo for a larger image

[PAS, p. 6-4]

  • Control actuators and HSTS based on FCC commands
  • Report control surface positions back to FCC’s
  • Located in multiple locations
    • Wings (4)
    • Tail (3)
    • Main gear well (1)
  • Each REU has 2 DC power sources for redundancy
  • Multi-channel (2) each with command and monitor lanes
    • Independently receive and process: FCC signals, Sensor data
    • Verifies proper actuator response: If anomalies occur, can disable actuation channels and data buses, reverts to a fail-safe state to prevent erroneous outputs
    • Each actuator has its own REU channel
  • Auto re-engagement of actuators in flight
    • With loss of and subsequent restoration of: REU electric power, Hydraulic pressure
    • FLT CTRL RESET Switch not required

Rotary Variable Differential Transducers (RVDTs)

When you get rid of pulleys, cables, and levers in a flight control system you end up needing a way to measure rotational movements, such as the angle of the power levers, position of the ailerons, etc. A Rotary Variable Differential Transducer (RVDT) is something like a very precise potentiometer. It takes an electrical input and varies the output according to the angle that is being measured. It makes for an inaccurate measurement if done purely using the analog output. But coupled to a computer and turned digital, it can be very precise.


Photo: G500 Spoiler RVDT (Item C), MM, §27-64-01, fig. 2 sheet 2
Click photo for a larger image

This is just an example, from one of the spoiler panels. The RVDT is the round part of Item C in the drawing.

Rudder Pedals

The rudder pedals feel "normal" though there are only electrons between you and the rudder.


Photo: G500 Rudder pedals, PAS, p. 6-30
Click photo for a larger image

[PAS, pp. 6-29 to 6-30]

  • Pilot and copilot rudder assemblies are mechanically connected.
  • The pedals will move +/- 3" from neutral.
  • There are position sensors attached to the rudder pedals that are read by the Flight Control Computers (FCCs). The FCCs send these signals to the rudder REUs which command their respective actuators to move the rudder. The maximum deflection is 25° at low speeds, 3.6° and high speeds.
  • Artificial feel is provided by damper and spring, rudder force is proportional to pedal displacement.

Spoilers / Speed Brakes

There are 3 panels on each wing: inboard, midboard, and outboard. The inboards and outboards are powered by the right hydraulic system, the midboards by the left hydraulic system. Only the outboards have a non-hydraulic backup through Electronic Backup Hydraulic Actuators (EBHAs). The Byzantine valve system of earlier Gulfstreams is gone, all the protection now is from the Flight Control Computer.



Photo: G500 Spoilers synoptic, PAS, p. 6-41
Click photo for a larger image

[PAS, p. 6-41] Ground Spoilers

  • During rejected takeoff or at touchdown, all spoiler panels deploy to 55°, ailerons fully deploy trailing edge up (full roll authority still available for crosswinds), increases drag, spoils lift which improves brake effectiveness
  • Triggers automatically when both throttles are at idle and one of the following criteria met: both main gear WOW = Ground, one main gear WOW = G and opposite wheel spin > 47 kts, or both main wheels spin and radar alt < 10’ and either flaps > 21° or GPWS Flap Inhibit selected
  • Auto stow when:
    • Airspeed and wheel speed < 42 kts for 10 secs, or
    • Either throttle not at idle, or if 2 of following 3 occur: both MLG WOW in air mode, wheel speed < 47 kts, or RA > 10'

Speed Brakes


Photo: G500 Speed Brakes synoptic, PAS, p. 6-43
Click photo for a larger image

[PAS, p. 6-43] Speed Brakes

  • Manually operated with handle on center pedestal; to operate first move to left and then pull aft.
  • Proportional to lever displacement; max extension in flight is 30°, on ground 55° with flaps ≥ 10° (with radio altimeter < 10' and WOW from either main gear, or 30° with flaps < 10°.
  • Electrically controlled from RVDT in handle, sends signal to FCCs.
  • When handle moved aft, CAS: Speed Brakes Extended
  • If thrust lever angle > 26° CAS: Speed Brakes Extended
  • If thrust lever angle increased > 90° the spoiler panels auto retract (handle remains extended), CAS: Speed Brakes Auto Retract.

Roll Spoilers


Photo: G500 Roll Spoilers synoptic, PAS, p. 6-44
Click photo for a larger image

[PAS, p. 6-44] Roll Spoilers

  • Midboard and outboard spoilers work in conjunction with ailerons to improve roll response
  • Fully automatic; sidestick roll signals from RVDTs are sent to FCCs which send commands to REUs that control the spoiler actuators
  • Spoiler extension varies based on sidestick inputs, 55° maximum

Synoptics Legend


Photo: Synoptics Legend (Sheet 1), G500 AFM, §03-13-190, sheet 1
Click photo for a larger image


Photo: Synoptics Legend (Sheet 2), G500 AFM, §03-13-190, sheet 2
Click photo for a larger image


Photo: Synoptics Legend (Sheet 3), G500 AFM, §03-13-190, sheet 3
Click photo for a larger image

Trim Switches

Pitch trim is automatic on two levels. If you simply push or pull the nose to where it needs to be, the stabilizer moves to center the elevator, relieving you of the need to apply pressure. But if you don't want to wait for that, pressing the AP DISC/TSS button the stabilizer trims for 1G at the current speed. But you can trim the old fashion way if you wish.

[PAS, p. 6-26]

  • Forward / Aft = NOSE DN / NOSE UP (split switch to prevent inadvertent activation)
  • Left / Right = LWD / RWD
  • Trims entire control surface (no trim tabs)
  • FCCs move stabilizer to "offload" elevator to keep it faired
  • Trim stops if switch activated for more than 3 seconds
  • Trim range: 100 knots to VMO / MMO
  • Will not trim to less than AOA limiting
  • If airspeed greater than 250 knots, will not trim to less than 187 knots (will require 0.5G back pressure to maintain level flight)

Photo: Pitch and roll trim control, PAS, p. 6-26
Click photo for a larger image

[PAS, p. 6-26, 6-31]

  • Pitch and Roll Trim
  • The primary switches are on the Active Control Sidestick (ACS). A secondary way is with the pedestal pitch trim switches. If you trim using the ACS switches for more than 3 seconds the trim stops. When between 100 knots and VMO/MMO, trim will go no lower than AOA Limiting. Mistrim → Can't trim below 187 kts; requires 0.5G of force to maintain level flight at 250 kts if trimmed for 187 kts.

    I think what this means is that if you try to trim the airplane to fly too slowly, the trim refuses and you will have to apply 0.5G back pressure to fly that slowly.

  • Yaw Trim
  • The yaw trim switch is on the center console. When turned, a signal is sent to the Flight Control Computer to move the entire rudder. Spring loaded to center position.

Yaw Trim

The Yaw Trim switch sends signals to the Flight Control Computer with instruction to trim the entire surface.


Photo: Yaw trim controls, PAS, p. 6-31
Click photo for a larger image

[PAS, p. 6-31] The yaw trim switch is on the center console. When turned, a signal is sent to the Flight Control Computer to move the entire rudder. Spring loaded to center position.


Photo: Yaw trim controls, PAS, p. 6-32
Click photo for a larger image


[AFM, §01-27-10] Normal Control Laws

  • Continued flight at or below stick shaker activation speed is prohibited.

  • NOTE

    The AOA limiting / stall protection system is only available in the normal flight control mode. Stick shaker/stall warning is provided in Alternate mode at 0.85 AOA.

  • Speed brake extension with flaps 39 or with landing gear extended is prohibited.

[AFM, §01-27-20] Degraded Control Laws

  • Flight into known icing conditions is prohibited when operating in a flight control law mode other than normal (Alternate, Direct or Backup). If the flight control law mode degrades from normal while in icing conditions, exit icing conditions as soon as possible.

  • NOTE

    The AOA limiting / stall protection system is only available in the normal flight control mode. Stick shaker/stall warning is provided in Alternate mode at 0.85 AOA.

  • Intentional degradation from normal mode or disabling of any flight control system is prohibited.

See Also:

Gulfstream GVII-G500 Airplane Flight Manual, Revision 4, August 29, 2019

Gulfstream GVII-G500 Production Aircraft Systems, Revision 3, July 15, 2019

Gulfstream GVII-G500 System Description Manual, Revision 2, December 15/19