United States Army Aviation Warfighting Center
Fort Rucker, Alabama
January 2008
UH-60A
STUDENT HANDOUT
UH-60A Automatic Flight Control System (AFCS)
4748-6
PROPONENT FOR THIS STUDENT HANDOUT IS:
110TH AVIATION BRIGADE
ATTN: ATZQ-ATB-AD-C
FORT RUCKER, ALABAMA 36362-5000
FD5: This product/publication has been reviewed by the product developers in coordination with the
USAAWC, Foreign Disclosure Officer, Fort Rucker, AL foreign disclosure authority. This product is
releasable to students from all requesting foreign countries without restrictions.
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TERMINAL LEARNING OBJECTIVE:
ACTION: Identify operational characteristics of the UH-60A/L Automatic Flight Control System (AFCS).
CONDITION: In a classroom given appropriate training devices, a list of characteristics, TM 1-1520-237-10, TM
1520-237-10CL, Aircrew Training Manual, and the student handout.
STANDARD: In accordance with (IAW ) TM 1-1520-237-10, TM 1-1520-237-10CL, Aircrew Training Manual,
and the student handout.
SAFETY REQUIREMENTS: Use care when operating training aids and/or devices.
RISK ASSESSMENT LEVEL: Low
ENVIRONMENTAL CONSIDERATIONS: It is the responsibility of all soldiers and DA civilians to protect the
environment from damage.
EVALUATION: You must answer 7 out of 10 questions correctly to receive a "GO" on this scoreable unit.
LEARNING STEP/ACTIVITY 1: Identify the functions and purpose of the Automatic Flight Control System
(AFCS).
a. Description. The AFCS is an electromechanical or electrohydro-mechanical servo system.
(1) Electrical signals drive mechanical and hydraulic actuators.
(2) Actuators position flight control linkages to provide automatic correction for unwanted aircraft
movements.
b. Function. The AFCS provides the helicopter with static and dynamic stability.
(1) Static stability (long-term stability) is the tendency to return to the pilot's desired attitude, airspeed, or
heading.
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(2) Dynamic stability (short-term stability) is the tendency to resist oscillation.
c. Reasons for system.
(1) Helicopters are not as stable as fixed-wing aircraft. External forces will cause any aircraft to change
attitude, airspeed, or heading. However, fixed-wing aircraft can be designed to return to the desired attitude
when the external force is removed (static stability).
(2) Helicopters have little or no tendency to return to the desired attitude. Since the rotor head moves with
the aircraft, helicopter will assume a new attitude and not the desired one.
(3) A helicopter hangs under a rotor head (like a pendulum) and swings or oscillates under rotor head.
Dynamic stability prevents porpoise in pitch, rock in roll, and fishtail in yaw.
(4) Pilot workload is much higher in a helicopter than in a fixed-wing aircraft.
(a) Constant correction is required to maintain attitude, airspeed, and heading.
(b) Constant correction is required to minimize oscillation.
(c) Instrument flying is extremely difficult.
(d) Accuracy of weapons is very poor due to lack of stability.
(e) Passengers are uncomfortable because of lack of stability.
d. Purpose of the AFCS is to enhance the stability and handling qualities of the helicopter.
(1) It provides static stability by holding-(a) Airspeed.
(b) Attitude.
(c) Heading.
(d) Coordination while turning
(2) It provides dynamic stability to prevent-(a) Porpoising in pitch axis.
(b) Rocking in roll axis.
(c) Fishtailing in yaw axis.
(3) AFCS Breakdown
(a) The Stabilator control system provides the helicopter with stability in the pitch axis by controlling the
position of the horizontal stabilator.
(b) The SAS enhances dynamic stability in the pitch, roll, and yaw axes.
(c)The trim system maintains flight control position and provides a force gradient which the pilot can
override. This system is essential for FPS system to operate.
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(d) FPS provides limited flight control positioning which assists in maintaining helicopter pitch and roll
attitudes, airspeed, heading, and turn coordination.
LEARNING STEP/ACTIVITY 2: Identify the operational characteristics of the stabilator system.
a. Description. The helicopter has a variable angle of incidence stabilator to enhance handling qualities. The
automatic mode of operation positions the stabilator to the best angle of attack for the existing flight conditions.
After the pilot engages the automatic mode, no further pilot action is required for stabilator operation. Two
stabilator amplifiers receive airspeed, collective stick position, pitch rate, and lateral acceleration information to
program the stabilator through the dual electric actuators.
b. Operation. The stabilator is programmed to-(1) Align stabilator and main rotor downwash in low speed flight to minimize nose up attitude resulting from
downwash.
(2) Provide collective coupling to minimize pitch attitude excursions due to collective inputs from the pilot.
A collective position sensor detects pilot collective displacement and programs the stabilator for a
corresponding amount of movement to counteract for pitch changes. This coupling of stabilator input for
collective displacement is automatically phased in between 30 and 60 KIAS
(3) Decrease angle of incidence with increased airspeed to improve static stability.
(4) Provide sideslip to pitch coupling to reduce susceptibility to gusts. When the helicopter is out of trim in a
slip or skid, pitch excursions are also induced as a result of the canted tail rotor and downwash on the stabilator.
Lateral accelerometers sense this out of trim condition and signal the stabilator amplifiers to compensate for the
pitch attitude change (called lateral to sideslip to pitch coupling). Nose left (right slip) results in the trailing edge
programming down. Nose right produces the opposite stabilator reaction.
(5) Provide pitch rate feedback to improve dynamic stability. The rate of pitch attitude change of the
helicopter is sensed by a pitch rate gyro in each of the two stabilator amplifiers and is used to position the
stabilator to help dampen pitch excursions during gusty wind conditions. A sudden pitch up due to gusts would
cause the stabilator to be programmed trailing edge down a small amount to induce a nose-down pitch to
dampen the initial support.
.
c. Actuators (2).
(1) Description. Two identical stabilator actuators connected back to back.
(2) Purpose. Actuators serve to move the stabilator to the required position.
(3) Location. The Number 2 actuator connects to the helicopters tail pylon (top). The number 1 actuator
connects to the stabilators upper surface.
(4) Operation. Each actuator contains an electric servo motor that responds to input signals by driving a
screwjack, which causes the actuator to extend or retract.
d. Limit switch/position transmitter.
(1) Description. This assembly contains four limit switches and a position transmitter.
(2) Purpose. The limit switches limit stabilator travel to about 9 degrees (trailing edge up) and 39 degrees
(trailing edge down). Its transmitter produces a three-wire synchro signal that drives the pilot and copilot's
stabilator position at the indicators.
(3) Location. The limit switch/position transmitter assembly is mounted in the helicopters tail pylon. It
connects through a shaft to the stabilators upper surface.
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e. Position indicators (2).
(1) Description. Two identical stabilator position indicators allow the pilot and copilot to monitor stabilator
position.
(2) Location. The indicators, are located on the instrument panel (just below the pilot's and copilot's
Airspeed Indicators).
(3) Operation. They contain OFF flags that are removed from view when power is applied and an indicator
pointer that displays stabilator angles between 10° up and 45° down.
NOTE: Although the indicators are marked 10° up and 45° down, do not confuse this with the actual stabilator
travel of 9° up and 39° down.
f. Stabilator position placards.
(1) Description. These lighted decals show maximum airspeed in relation to stabilator position for stuck
stabilator.
(2) Location. A lighted decal is located beside each stabilator indicator.
(3) Function. Decals serve to limit airspeed for varying stabilator angles. They are marked with degrees on
the left side and airspeed on the right side.
g. Amplifiers (2).
(1) Description. The two identical stabilator control amplifiers contain electronic components, pitch rate
gyros used in automatic stabilator control, and relays that allow manual control.
(2) Purpose. Stabilator control amplifiers provide lateral acceleration, airspeed discrete signals, and pitch
rate signals to the SAS/FPS computer and automatic stabilator control and provide relays that allow manual
control of the stabilator.
(3) Location. Stabilator control amplifiers are mounted on the aft overhead cabin. The amplifier located on
the left side is identified as Number 1; the amplifier located on the right side is Number 2.
(4) Operation.
(a) The Number 1 amplifier provides control signals to the Number 1 actuator.
(b) The Number 1 stabilator control amplifier provides filtered lateral acceleration and airspeed discrete
signals to the SAS/FPS computer. It also supplies airspeed discrete and filtered pitch rate signals to SAS
amplifier.
(c) The Number 2 stabilator control amplifier provides filtered pitch rate and filtered lateral acceleration
signals to the SAS/FPS computer. Each stabilator amplifier contains a pitch rate gyro that produces a DC signal
proportional to the helicopters rate of pitch attitude change
(d) Bias potentiometers, in both amplifiers, supply signals that command the stabilator to 37 degrees ±2
degrees (trailing edge down) when all other signals are minimum. (Aircraft on ground.)
h. Airspeed sensors
(1) Description. Airspeed and air data transducers produce identical DC electrical signals that
represent the helicopters forward airspeed between 30 and 180 knots.
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(2) Purpose. Airspeed and air data transducers supply airspeed signals to the stabilator amplifiers for the
stabilators automatic mode operation and the SAS/FPS computer.
(3) Location. Airspeed and air data transducers are located on the cockpit bulkheads (forward of and
below the instrument panel). The air data transducer is on the right side of the helicopter; the airspeed
transducer is on the left side.
(4) Operation. The airspeed transducer supplies an airspeed signal to the Number 1 stabilator control
amplifier. The air data transducer supplies airspeed signals to the Number 2-stabilator-control amplifier. The
air data transducer also supplies airspeed and pressure altitude signals to the command instrument system.
Both transducers supply airspeed signals to the SAS/FPS computer and receive air pressure inputs from the
instrument pitot-static system.
i. Collective stick position transducer (2).
(1) Description. The two collective stick transducers are identical.
(2) Purpose. They supply DC collective stick position signals to the stabilator control amplifiers and
SAS/FPS computer and from the Number 2 collective stick position transducer to the CIS processor.
(3) Location. Both transducers are mounted on the right side of the flight controls mixer assembly.
(4) Operation.
(a) Number 1 transducer. The (forward) Number 1 transducer supplies signals to the Number 1
amplifier. It also supplies signals to the SAS/FPS computer.
(b) Number 2 transducer. The (aft) Number 2 transducer supplies signals to the Number 2 amplifier.
An additional section of the Number 2 transducer supplies signals to the command instrument system. The first
section also supplies signals to the SAS/FPS computer.
j. Lateral accelerometer (2).
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(1) Description. The two lateral accelerometers are Identical.
(2) Purpose. They provide the stabilator control amplifiers with DC electrical signals that represent the
relationship of the helicopter bank angle to its turn rate.
(3) Location. The lateral accelerometers are mounted in the cabin ceiling. The Number 1 accelerometer is
located on the left side of the helicopter, and the Number 2 accelerometer is located on the right side. On some
aircraft, they are located on the left and right sides of the stabilator amplifiers.
(4) Operation.
(a) The Number 1 accelerometer receives excitation from the left side of the helicopter and supplies
signals to the Number 1 amplifier.
(b) The Number 2 accelerometer receives excitation from the right side of the helicopter and provides
signals to the Number 2 amplifier.
k. Control panel.
(1) Description. The stabilator/flight control panel contains switches and relays used to control the
stabilator system.
(2) Location. The panel is located on the center of the lower console.
(3) Switches.
(a) Cyclic mounted stabilator slew-up switch. The preferred method of manually slewing the stabilator
is to use the cyclic mounted stabilator slew-up switch.
NOTE: Use of the cyclic mounted stabilator slew-up switch should be announced to the crew to minimize
cockpit confusion.
(b) MAN SLEW. This switch is spring loaded to the center (OFF) lever lock switch. It allows the pilot to
position the stabilator to any fixed position within its full range of travel.
NOTE: Use of the MAN SLEW causes the AUTO control mode to disengage.
(c) AUTO control reset. An illuminated push-button switch displays the word ON when the stabilators
automatic control mode is engaged. Pushing the button resets the automatic mode if it fails.
NOTE: Airspeed transducers must be connected before the automatic mode will engage.
(d) Test button. The test button is operational at airspeeds below 60 knots and is used to ground check
the stabilator system. It causes the stabilator to drive up and the AUTO mode to disengage. Relays in the
control panel cause the STABILATOR caution and master caution to illuminate and a beeping tone to be
supplied to the pilot's and copilot's headsets when the AUTO mode is OFF. Pushing the master caution capsule
resets the master caution and tone.
LEARNING STEP/ACTIVITY 3: Identify the operational characteristics of the stabilator system modes of
operation.
a. Automatic mode operation.
(1) Description. The automatic mode positions the stabilator to the best position for existing flight
conditions without any input from the pilot. It engages automatically when power is applied to the aircraft.
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WARNING: MAKE SURE ALL PERSONNEL AND EQUIPMENT ARE CLEAR OF THE STABILATOR BEFORE
APPLYING ELECTRICAL POWER TO THE HELICOPTER.
(a) The automatic mode disengages if positions of the stabilator actuators disagree by an amount that
depends on forward airspeed. Shutdown occurs at 10° of error at 0 knots and 4° of error at 150 knots.
(b) The automatic mode also disengages when the MAN SLEW switch is moved to the UP or DN
position or the cyclic mounted stabilator slew up switch is used.
(2) Operation (general).
(a) Both stabilator amplifiers receive AC and DC power when electrical power is applied to the
helicopter. They, in turn, supply VDC to the airspeed and air data transducers, collective stick position
transducers, lateral accelerometers, and actuator feedback potentiometers.
(b) Airspeed and air data transducers produce DC output signals that increase for forward airspeeds
between 30 and 180 knots.
(c) Collective stick position transducers produce a DC output signal that is proportional to the collective
stick's position.
(d) A centered collective equals 0 volts. A down collective results in a positive signal; an up collective
causes the output to go negative. (Output voltage is 1.34 volts per inch of stick displacement from the center.)
(e) Pitch rate signal.
1. Each stabilator amplifier contains a pitch rate gyro. These produce DC signals that represent
the helicopters rate of pitch attitude change.
2. Pitch rate signals also are routed out of each stabilator amplifier through filters.
3. The number 1 signal is used by the SAS amplifier. The number 2 signal is used by the SAS/FPS
computer.
b. Manual mode operation.
(1) Manual mode operation allows the pilot to control stabilator position.
(2) The stabilator control/flight control panel stabilator MAN SLEW switch allows selection of any fixed
stabilator position between 9 degrees (trailing edge up) and 39 degrees (trailing edge down). The switch has
four sets of contacts. Two sets of "hot slew" contacts provide 28-VDC power to the Number 1 and Number 2
stabilator control amplifiers anytime the switch is moved up or down. This output causes amplifier logic circuits
to disengage the automatic mode. The additional two sets of switch contacts supply 28 VDC to either slew up
or slew down relays in both amplifiers. Contacts of the slew up or slew down relays supply the interlocks (28
VDC) power to the actuator motors which cause the stabilator to drive. The pilot must never exceed the limit
airspeed, listed on instrument panel placards, for any selected stabilator angle.
c. Degraded operation.
(1) One actuator (trailing edge down). If one actuator is not functional at +39° (trailing edge down), range
of control with the remaining good actuator is from +39° to +5.51°.
(2) One actuator (trailing edge up). If one actuator is not functional at -9° (trailing edge up), range of control
with the remaining good actuator is from -9° to 22.04°.
NOTE: Long pitot-static streamers are attached to door handles to minimize the chance of operating the aircraft
with covers in place.
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WARNING: COVERS ON PITOT TUBES WOULD CAUSE THE STABILATOR TO REMAIN IN THE TRAILINGEDGE DOWN POSITION WITH NO CAUTION LIGHTS OR AURAL WARNING.
LEARNING STEP/ACTIVITY 4: Identify the operational characteristics of the stability augmentation system
(SAS).
a. Analog stability augmentation system. The analog system (SAS 1) is one of two SAS systems. It operates
independently of the SAS/FPS computer and provides the aviator with redundancy.
b. Digital stability augmentation system (SAS 2). SAS 2 is operated by the SAS/FPS computer. It is
independent of SAS 1 and provides stability in the same axis using the same actuators.
c. Purpose--provides dynamic stability in the pitch, roll, and yaw axes.
d. Components.
(1) Actuators (3).
(a) Description. Three actuators are provided for the pitch, roll, and yaw channels.
(b) Purpose. Actuators link SAS electronic components to the helicopters mechanical flight control
system.
(c) Location. The actuators are mounted on the transmission deck at the pilot assist servo.
(d) Operation.
1. Hydraulic pressure. An electrohydraulic servo control flapper valve allows 3,000 psi hydraulic
pressure from the Number 2 hydraulic system, or backup system, to operate the actuator.
2. Electrical inputs. Electrical inputs are supplied by the analog SAS amplifier and digital SAS/FPS
computer. The actuators respond to electrical inputs and hydraulic pressure by moving control linkages that
change rotor blade angles without moving cockpit controls.
3. Actuator stroke. The actuator stroke is mechanically limited to allow maximum control authority
of 10 percent of the total control available to the pilot.
4. Lock pin. Each actuator contains a centering lock pin that locks the actuator output at midstroke
when hydraulic pressure is removed.
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(2) SAS Amplifier.
(a) Description. It contains a rate gyro that serves as a sensor for yaw SAS.
(b) Purpose. It processes aircraft sensor signals to develop command signals that are applied to SAS
actuators when SAS 1 is engaged.
(c) Location. The SAS amplifier is located on the floor of the electronic compartment's "tunnel" (below
the instrument panel).
(d) Control authority. Amplifier electrical outputs are limited to allow analog SAS (SAS 1) a maximum
of a ±5-percent authority. Gain control circuits double each channel's gain when SAS 2 is switched OFF.
Authority remains at 5 percent.
(3) Sensors.
(a) No. 1 stabilator control amplifier supplies the SAS amplifier with filtered pitch rate, filtered and nulled
lateral acceleration, and airspeed discrete signals. The pitch rate signal originates from a rate gyro inside the
stabilator amplifier. The airspeed discrete signal also is developed by the stabilator amplifier.
(b) The pilot's (Number 2) vertical gyro supplies the SAS amplifier with a signal that represents the
helicopters roll attitude.
(c) An airspeed transducer supplies a signal that determines polarity of the airspeed discrete.
(Determines if the airspeed is above or below 60 knots.)
(d) The yaw rate gyro inside the SAS/ amplifier supplies SAS with signals that represents the
helicopters yaw rate.
(4) Controls. SAS 1 and SAS 2 receive ON and OFF control through two switches, labeled SAS 1 and SAS
2, on the flight control panel.
(a) Depressing either switch ON removes 28 VDC from the SAS shutoff valve allowing the valve to
supply pressure to SAS actuators.
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(b) SAS 1 switch energizes the SAS amplifier for SAS 1 coils, and the SAS 2 switch energizes circuitry
in the SAS/FPS computer for SAS 2 coils.
(c) SAS 1 and SAS 2 each have a 5-percent authority for a total of a 10- percent authority.
(d) Failure of SAS 1 will not be indicated visually but may be known by the lack of aircraft stability.
(e) Failure of sensors controlling SAS 2 will cause the failure advisory panel on the flight control panel
to illuminate.
(5) Indicators. In case of loss of actuator pressure, or if both SAS 1 and SAS 2 are off, the SAS OFF
caution will appear.
LEARNING STEP/ACTIVITY 5: Identify the operational characteristics of the digital automatic flight controls
system (AFCS/SAS2).
a. Description. The digital AFCS will provide the following:
(1) Cyclic stick and pedal trim.
(2) Stability augmentation (SAS 2).
(3) Autopilot functions (FPS).
b. Components.
(1) Computer.
(a) Purpose. The computer processes signals used for the digital AFCS.
(b) Location. It is located below the center console and accessible from the cabin.
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(c) Power. The computer is powered by the Number 2 AC primary bus and Number1 and Number 2
primary DC busses.
(2) Vertical gyros. Two vertical gyros, located in the nose electronic compartment, produce attitude signals
for pitch and roll.
(3) ASN 43 compass system directional gyro. This gyro is located in the nose electronic compartment and
produces a heading signal.
(4) Airspeed and air data transducers.
(a) Location. These transducers are located in the forward section of the cockpit, above the tail rotor
pedals, with airspeed on the left side and air data on the right side.
(b) Operation. They operate from the pitotstatic system and are powered by the stabilator system.
(c) Signal use.
1. Yaw trim.
2. Pitch FPS--airspeed hold.
3. Yaw FPS--automatic turn coordination logic
4. Yaw SAS 2.
(5) Rate gyros. Rate gyros in the pitch, roll, and yaw channels control the rate of any change.
(6) Lateral accelerometers. These are used to produce signals when the helicopter slips or skids.
(7) Collective stick position transducers. These are located at the flight controls mixer. They produce a
signal that represents collective stick's position. The signals are used for yaw trim, collective to yaw coupling,
and pitch FPS airspeed hold.
(8) Drag beam switch/Weight on Wheels switch (WOW). This switch -(a) Holds pitch, roll, and yaw integrators when helicopter is on ground.
(b) Prevents FPS from continuing to drive controls if aircraft is taxied to spot that is not level.
c. System's operation.
(1) Electrical supply. The SAS amplifier receives 115-VAC power from the AC essential bus. The SAS
amplifier and AUTO flight control panel engage circuits are supplied with 28-VDC essential bus. Operation of
roll SAS and yaw turn coordination is dependent on the pilot's (Number 2) vertical gyro. It is powered from the
AC essential bus. Pitch and roll channels receive signals from the Number 1 stabilator amplifier.
(2) Hydraulic supply. Hydraulic pressure, required for SAS actuator operation, is supplied by the Number 2
or backup hydraulic pump. Pressure at 3,000 psi is supplied through the SAS shutoff valve on the pilot's assist
module. The valve may be shut OFF by turning OFF SAS 1 and SAS 2.
(3) Pitch channel. The SAS pitch channel causes the main rotor tip path to tilt in the direction required to
oppose pitch attitude changes. This provides the helicopter with dynamic stability by minimizing oscillations in
the pitch axis.
(a) The SAS amplifier receives an electrical signal that represents the helicopters rate of pitch attitude
change from the Number I stabilator amplifier.
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(b) SAS amplifier circuits modify the rate signal to provide desired aircraft and system response.
(c) SAS 1 gain depends on the engage condition of SAS 2. If SAS 2 is ON, SAS 1 operates at a normal
gain. SAS 1 and SAS 2 supply the actuator with signals to stabilize the helicopter. If SAS 2 is OFF, the SAS 1
amplifier doubles its gain. This provides larger signals for a given aircraft movement to help compensate for the
loss of SAS 2.
(d) The servo valve driver supplies current to operate the SAS actuator.
LEARNING STEP/ACTIVITY 6: Identify the operational characteristics of the trim system.
a. Trim actuators are connected to flight control linkages in a manner that allows them to move cockpit controls.
(1) Pitch--to fore and aft cyclic stick linkage.
(2) Roll--to lateral cyclic stick linkage.
(3) Yaw--to pedal linkage.
b. Pitch actuator is hydraulic.
(1) It receives 1,000 psi from pilot assist module.
(2) Pressure is supplied only if trim is ON and computer detects no pitch trim malfunction.
c. Roll and yaw actuators are electromechanical.
d. FPS provides 100 percent control authority using trim actuators.
e. All actuators operate at a limited rate (about 10 percent per second). They are self-limited and also limited
by computer.
f. All actuators contain override springs which-(1) Allow pilot a 100-percent override of trim.
(2) Provide breakout and gradient forces at controls.
g. All actuators provide dampening.
(1) Force at controls is proportional to rate of movement.
(2) Small force is present during pitch and roll.
h. Roll and yaw actuators contain override clutches which-(1) Allow override if electromechanical actuator jams.
(2) Slip when sufficient force is applied to controls.
i. Computer controls trim actuators when trim is ON.
(1) Feedback signal changes if actuator moves.
(2) Change in feedback signal causes computer to output command to actuator.
(a) Command causes actuator to drive back toward trimmed position.
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(b) Actuator remains very close to trimmed position.
(c) Actuator holds the cyclic stick fixed.
j. Cyclic stick trim release button causes computer to release trim.
(1) Hydraulic pressure is removed from pitch trim actuator.
(2) Feedback signal is zeroed.
(3) Stick is free to move.
(4) Trimmed or referenced position becomes changed.
(5) Stick is trimmed to position where button is released.
(6) Cyclic stick trim switch changes computers command signal to actuator.
(a) Actuator will drive.
(b) Stick will drive at about 0.4 inches per second.
k. Roll trim actuator has a built-in servo system.
(1) System holds actuator and cyclic stick fixed when trim is ON and not released.
(2) It is referenced through clutches and centering springs in actuators when trim is OFF or released.
(3) Computer supplies command to the actuator when cyclic stick trim switch is used.
NOTE: When the cyclic trim switch is slewed left and right, while on the ground and with FPS on, the cyclic will
return to center.
l. Yaw trim actuator has a built-in servo system.
(1) System holds actuator and pedals fixed when trim is ON and not released.
(a) It is referenced when yaw trim is released.
(b) Pedal microswitches release yaw trim at airspeeds less than 60 knots.
(c) Pedal microswitch and cyclic trim (pressed at the same time), is required to release yaw trim at
airspeeds above 60 Kts. This disengages yaw trim and the turn coordination feature.
(2) Computer supplies command signal to yaw trim actuator and electronic collective to airspeed to yaw
coupling (maximum effect below 40 knots, no effect above 100 knots).
(a) Boost servo pressure must be ON for proper operation.
(b) Up collective causes left pedal to move forward.
(c) There is a maximum coupling at airspeeds below 40 knots.
(d) There is no coupling at airspeeds greater than 100 knots.
(e) Computer compensates for changes in lateral thrust provided by cambered vertical stabilizer.
m. Actuator failure causes computer to disable a failed trim system.
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(1) Pedals or cyclic stick become free in affected axis trim and FPS is inoperative.
(2) During pitch failure, actuator hydraulic pressure is removed.
(3) During roll or yaw failure, the actuator clutch is disengaged.
(4) Computer senses failure when actuator position signal (feedback) does not agree with command signal.
(a) Trim failure advisory light will be ON.
(b) Trim failure and flight path stabilization caution lights will be ON.
n. Computer failure cause affected trim axis to shut down.
(1) Pitch channel. Flight indications are the same as for an actuator failure.
(2) Roll or yaw.
(a) Stick or pedals will be free.
(b) Trim failure and flight path stabilization caution lights will be ON.
(c) Computer (CPTR) failure advisory light will be ON.
(d) Trim and FPS in affected axis will be inoperative.
LEARNING STEP/ACTIVITY 7: Identify the operational characteristics of the flight path stabilization system
(FPS).
a. It provides the autopilot functions, which are to -(1) Hold roll attitude.
(2) Hold pitch attitude/airspeed.
(3) Hold heading or automatic turn coordination.
b. The flight path stabilization operates by trimming cockpit controls.
(1) Trim must be ON.
(2) Boost servos must be ON for proper yaw channel operation.
(3) There is a 100-percent control authority.
(4) There is a limited rate of response.
(a) The maximum response is limited to 10 percent per second.
(b) At least one SAS must be ON to provide the rapid response required for short term stability.
(5) Complete pilot override is available through trim force gradient.
c. Roll channel provides attitude hold by trimming cyclic stick to position required to maintain the pilot's desired
attitude.
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d. Pitch channel provides attitude hold (below 60 Kts) and attitude & airspeed hold (above 60 Kts)(airspeed hold
overrides attitude hold), by trimming cyclic stick to position required to maintain pilot’s desired attitude and
airspeed..
NOTE: Airspeed hold goes OFF for 30 seconds after collective stick is moved through the 60-percent position
below 100 knots.
e. Yaw channel provides heading hold or automatic turn coordination/automatic turn logic by driving pedals as
required to maintain pilots desired heading or balanced flight. It holds heading unless pedal trim is released or
automatic turn coordination/automatic turn logic is engaged.
(1) Pedal trim is released when airspeeds are less than 60 knots by pedal microswitches or when
airspeeds are greater than 60 knots by pedal microswitches and the cyclic trim release.
(2) Automatic turn coordination/automatic turn logic operates only at airspeeds above 60 knots and
engages when the pilot slews (by use of the cyclic stick trim switch) left or right about ½ inch and a roll attitude
of about 1.5 degrees or more.
f. FPS failures.
(1) Directional gyro failure disables heading hold.
(a) Pedal trim and automatic turn coordination remains functional.
(b) Flight path stabilization caution light is ON.
(c) Gyro failure advisory light is ON.
(2) Airspeed sensor failure disables automatic turn coordination.
(a) Flight path stabilization caution light is ON.
(b) Airspeed failure advisory light is ON.
(3) Lateral accelerometer failure disables automatic turn coordination.
(a) Flight path stabilization caution light is ON.
(b) ACCL failure advisory light is ON.
(4) Vertical gyro (roll) failure.
(a) Flight path stabilization caution light is ON.
(b) Gyro failure advisory light is ON.
(5) Yaw rate gyro failure.
(a) SAS 2 failure advisory light is ON.
(b) RGYR failure advisory light is ON.
(6) Lateral stick force sensor failure.
(a) Flight path stabilization caution light is ON.
(b) Trim failure advisory light is ON.
D-17
D-18
DIGITAL AUTOMATIC FLIGHT CONTROL UNITS EFFECTS OF MALFUNCTIONS
D-19
STABILATOR SYSTEM
R/S
pitot
tube
ver 1.1
11/00
COLLECTIVE STICK
POSITION TRANSDUCERS
(on Mechanical Mixing Unit)
USE OF EXCHANGED A/S SIGNALS
#2 Lateral
Accelerometer
AIR DATA
TRANSDUCER
80 KTS or LESS: Amps use higher of the two A/S signals.
80 KTS & ABOVE: Each amp uses it's own A/S signal.
(if one amp is sensing below 80 with the other sensing
above 80, the low one programs for 80 KTS)
STABILATOR CONTROLS
MAN SLEW
UP
PNL
LTS
AUTO
CONTROL
TEST
O
F
F
ON
DN
SAS 1
AUTO FLIGHT CONTROL
SAS 2
TRIM
R
E
S
E
T
ON
ON
BOOST
ON
Stab position
Limit & A/S
signals
#1 Stab.
Amp.
pitch rate
gyro
FPS
PNL
LTS
ON
#2 Stab.
Amp.
pitch rate
gyro
to
both
stab.
amps.
ON
NO.2 ACTUATOR
(moves only when converted power is
present via #2 DC PRI bus)
position commands
2
FAILURE ADVISORY
R
E
S
E
T
CPTR SAS 2
ACCL CLTV
TRIM RGYR
A/S GYRD
POWER ON RESET
R
E
S
E
T
position & limit feedback
NO.1 ACTUATOR
(can be moved with battery
power via DC ESS bus)
1
MAN SLEW: Sprung to off position, selects manual mode
when moved up or down, alows manual positioning of
stabilator to any position.
L/S
pitot
tube
#1Lateral
Accelerometer
9 deg
TEST: Functions @ 60 KTS & below, signals #1amp.
to retracts #1 actuator, fault monitoring detects miscompare
(IAW range chart) between #1/#2 actuators and switches
to manual mode (audible & visual warnings activate).
STABILATOR
AUTO CONTROL: Switches from manual to automatic mode.
AIRSPEED
TRANSDUCER
RANGE OF MOTION
both actuators: @ 48 deg.
CYCLIC SLEW: Selects manual mode & moves stab. up only.
(stabilator up movement stops when switch is released)
single actuator: @ 35 deg.
(may be less depending
on stab position when
other actuator was lost)
10
STAB
POS
STAB
POS
DEG
DEG
Copilot
Pilot
(Used for checks & rigging)
0 deg
0
39 deg
stabilator
actuator
position
0
4
0
30
airspeed
150
Miscompare Range Chart
D-27
FLIGHT PATH STABILIZATION SYSTEM
MAIN SLEW
UP
O
F
F
AUTO
CONTROL
TEST
R
E
S
E
T
ON
SAS 1
SAS 2
ON
ON
ON
TRIM
ON
R
E
S
E
T
FPS
ON
FAILURE
BOOST
ver 1.1
10/97
Input sensor B,I,M,P, and Q
STABILATOR CONTROLS
CPTR SAS 2
TRIM RGYR
ADVISORY
ACCL CLTV
A/S
GYRO
R
E
S
E
T
TRIM FAIL
FLT PATH
STAB
POWER ON RESET
fault monitoring / advisory
SAS 2
TRIM
FPS
S
E
N
S
O
R
S
1
GYRO
COMP
RATE
VERT
AIR
SPEED
TRIM
ACT
COLL
STICK
SAS
VALVE
DIR
2
1 SAS2 sensor signals: Q,G,J,K,N,O,H,
and R
2 TRIM sensor signals: E and B
3
LAT
ACCEL
FAN
FAIL
PROC A
3 FPS sensor signals: B,E,J,K,N, and O
GND
FAN
TEST
NORM
PROC B
NOTE: signals A,F,L,M, and R are
for fault monitoring only
J139
J12
J111
J121
Automatic Flight Control Input signals
A
B
C
D
E
F
G
H
I
air data transducer signal
air speed transducer signal
#1 lateral accelerometer signal
#2 lateral accelerometer signal
#1 collective position sensor
#2 collective position sensor
roll rate gyro signal
#2 yaw rate gyro signal
#1 pitch rate gyro signal (#1 stab amp)
D-27
J #2 pitch rate gyro signal (#2 stab amp)
K heading signal (gyro magnetic compass)
L pilots vert. gyro pitch signal (att. ind. sys.)
M pilots vert. gyro roll signal (att. ind.sys.)
N copilots vert. gyro pitch signal (att. ind. sys.)
O copilots vert. gyro roll signal (att.ind.sys.)
P #1 yaw rate gyro signal (from sas amplifer)
Q #1 filtered lateral acceleromtere signal (#1 stab amp)
R #2 filtered lateral acceleromtere signal (#2 stab amp)
STABILATOR SYSTEM
SIGNALS IN
# 1 CLTV
# 1 ACCL
A/S X-DUCTER
#1 Stab.
Amp.
pitch rate
gyro
SIGNALS OUT
SIGNALS OUT
#1 CLTV
#1 ACCL
A/S X-DUCER
PITCH RATE
GYRO
#2 CLTV
#2 ACCL
AIR DATA
PITCH RATE
GYRO
2
TEST FAULT
MONITORING
CIRCUIT
STABILATOR CONTROL PANEL
STAB. MODE SELECT (AUTO)
STAB. MAN. SLEW
STAB. TEST
1
SIGNALS IN
#1 PITCH RATE GYRO
#2 VERTICAL GYRO
#1 ACCL.
A/S X-DUCER
TEST
DRIVE / ACT ONLY
STABILATOR
SIGNALS OUT
PITCH RATE GYRO
VERTICAL GYRO
YAW RATE
GYRO
#2 Stab.
Amp.
pitch rate
gyro
TEST FAULT
MONITORING
CIRCUIT
SAS / FPS SYSTEM
SAS 1
AMP.
SIGNALS IN
# 2 CLTV
# 2 ACCL
A/D X-DUCER
#1 ACCL.
A/S X-DUCER
YAW RATE GYRO
SAS
ACTUATORS
TRIM ACTUATORS
PITCH
ROLL
YAW
SIGNAL OUT
PITCH
fault monitoring / advisory
ROLL
TRIM
SIGNALS IN
SAS 2
YAW
PILOT'S
CYCLIC AND
PEDAL TRIM
CONTROL
SWITCHES
#1 CLTV
A/S X-DUCER
TRIM
(COPILOT'S) #1 VERTICAL GYRO
NOTE:
#1 AND #2 ACCL
FILTERED SIGNALS ARE
PROCESSED IN THE STABILATOR AMP
#1 AND #2 ACCL (FILTERED)
AIR DATA X-DUCER
YAW RATE GYRO
A/S X-DUCER
#1 CLTV
#1 ACCL.
#1 PITCH RATE GYRO
FPS
ROLL RATE GYRO
#2 PITCH RATE GYRO
ASN 43 COMPASS
(HEADING)
GYRO
COMP
RATE
VERT
AIR
SPEED
TRIM
ACT
COLL
STICK
SAS
VALVE
DIR
LAT
ACCEL
FAN
FAIL
PROC A
GND
FAN
TEST
NORM
PROC B
J139
J12
J111
D-31
J121