United States Army Aviation Center of Excellence
Fort Rucker, Alabama
February 2019
H-60M FSXXI HANDOUT
Automatic Flight Control System
011-UH60M-1000
PROPONENT FOR THIS STUDENT HANDOUT IS:
1st BN 212th AVN REGT
ATTN: ATZQ-ATB-L
Fort Rucker, Alabama 36362-5000
FD2 This training product has been reviewed by the training developers in coordination with the USAACE
foreign disclosure officer. This training product can be used to instruct international military students when
the country meets specific criteria. Specify requirement(s) that each country must meet (select all that are
appropriate): 1) Must purchase equipment through FMS UH-60M; 2) Must be a member of a specific group
or coalition valid FMS customer; 3) Must have an accepted clearance (must be authorized under an
identified general security agreement with the US); 4) May not attend FD3 modules Blue Force Tracker
(BFT) of UH-60M Flight/Mission Display System (F/MDS) Multifunction Displays (MFDs) and UH-60M
Common Missile Warning System (CMWS); 5) Other ; FMS students may attend AMPS training, obtain the
student handout and operate the software under supervision in a training environment; however, any further
use or distribution of the software or pocket guide is RESTRICTED.
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ENABLING LEARNING OBJECTIVE:
ACTION:
Identify the operational characteristics that pertain to the H-60M Automatic Flight
Control System (AFCS).
CONDITION:
In a classroom, given a laptop computer/tablet, access to the AFCS student
handout, TM 1-1520-280-10, relevant training aids and multimedia.
STANDARD:
Identify the operational characteristics that pertain to the UH-60M AFCS IAW TM 11520-280-10.
1. Learning Step/Activity: Define the purpose and terminology of the UH-60M Automatic Flight Control
System (AFCS).
a. The Automatic Flight Control System (AFCS) enhances the stability and handling qualities of the
helicopter. It is comprised of five basic subsystems: Stability Augmentation System (SAS), Trim,
Flight Path Stabilization (FPS), Stabilator and the Flight Director Set (FD).
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b. The AFCS subsystems provide the helicopter with static and dynamic stability. AFCS provides
oscillation damping (dynamic stability) and maintains desired attitude, airspeed, and handling
(static stability).
(1) Static stability (Long term stability) is the tendency to return to the pilot’s desired attitude or
heading.
(2) Static stability is provided by the AFCS holding the attitude (pitch and roll), heading and
coordination while turning.
(3) Dynamic stability (Short term stability) is the tendency to resist movements. Dynamic stability
prevents porpoising in the pitch axis, rocking in the roll axis, and fishtailing in the yaw axis.
(a) Porpoising–Pitch axis.
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(b) Rocking – Roll axis.
(c) Fish tailing- Yaw axis.
c.
AFCS provides two types of control, identified as inner-loop and outer-loop. Both inner and outer
loops allow for complete pilot override through the normal use of the flight controls.
(1) The inner-loop (SAS) employs rate damping to improve helicopter stability, this system is fast
in response, limited in authority to 10% maximum, and operates without causing movement
of the flight controls. An inner loop servo (SAS) makes a flight control input with no feedback
into the cockpit controls
(2) The outer-loop (FPS & Trim) provides long-term inputs by trimming the flight controls to the
position required to maintain the selected flight attitude. It is capable of driving the flight
controls throughout their full range of travel (100% authority) at a limited rate of 10% per
second. An outer loop servo (FPS & TRIM) makes a flight control input and has feedback into
the cockpit controls.
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2. Learning Step/Activity: Provide a brief description of the AFCS computer system and operational
overview
a. AFCS is a computer system comprised of the following:
(1) Inputs - Attitude and rate signals from 2 EGIs, airspeed signals from 2 ADCs, signals from 2
CLTV position sensors, 2 lateral accelerometers, 2 longitudinal accelerometers, a vertical
accelerometer, and the stabilator position transmitter and limit switch assembly.
(2) Processors- Number 1 & 2 Flight Control Computers (FCC’s).
(3) Outputs-: TRIM, SAS, and stabilator actuators.
b. AFCS is comprised of the following 5 subsystems:
(1) The Stabilator improves flying qualities in the pitch axis by positioning the stabilator by means
of electromechanical actuators in response to collective, lateral acceleration, airspeed and
pitch rate inputs.
(2) Stability Augmentation System (SAS) provides short-term rate damping in the pitch, roll, and
yaw axis.
(3) Trim – provides control positioning and force gradient functions.
(4) Flight Path Stabilization (FPS) commands the pitch, roll, and yaw trim actuators to provide
control positioning/ basic autopilot functions.
(5) The Flight Director (FD) provides steering commands from pilot initiated inputs; or directly
coupled for autopilot flight.
c.
The following systems/components interface with the AFCS:
(1) Pilot and copilot Flight Director/Display Control Panels (FD/DCPs).
(2) No. 1 and No. 2 Embedded Global Position System/Inertial Navigation Sets (EGIs).
(3) No.1 and No. 2 Air Data Computers (ADCs).
(4) Radar Altimeter (RADALT).
(5) No. 1 and No. 2 Data Concentrator Units (DCUs).
(6) No. 1 and No. 2 Flight Management Systems (FMS).
(7) Cyclic Sticks Trim Release (TRIM REL) and Trim (TRIM) switches.
(8) Collective Sticks Trim Release (TRIM REL) and Trim (TRIM DN/UP L/R) switches.
(9) Pedal trim release switches (micro switches on pedals)
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d. The central components of the AFCS systems are the two redundant FCCs. Each computer
receives signals from the pilot flight controls, motions sensors, control panels, and avionics
systems to compute commands which are sent to the cockpit, trim actuators, SAS actuators, and
the stabilator actuators. The No. 2 FCC is the only FCC that sends signals to the Trim actuators
for dynamic flight control inputs. The computers also provide self monitoring, fault isolation, failure
alerts, and advisory and caution indications. Some sensor inputs to each FCC are compared via
the Cross Channel Data Link (CCDL).
e. Operational Characteristics of the Flight Control Computers (FCC):
(1) Each FCC receives signals from sensors, which are processed and commands are sent to
the respective actuators. Both FCC’s are identical and interchangeable, the difference is what
actuator they send their signal to.
(2) The No. 1 FCC signals are sent to SAS 1 and No.1 Stabilator actuator.
(3) The No. 2 FCC signals are sent to SAS 2, No. 2 stabilator actuator, and the Trim actuators.
(4) The FCC’s provide self monitoring, IBIT, fault isolation, failure alerts, and caution and
advisory indications.
(5) The FCC’s compare information via the Cross Channel Data Link (CCDL).
f.
The No. 1 FCC is located on the middle shelf in the transition section and the No. 2 FCC is
located in the nose compartment.
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g. Stabilator Control/Auto Flight Control Panel – contains button type, push on switches that provide
the ability to turn SAS 1, SAS 2, TRIM, FPS, and SAS/BOOST ON or off. It also provides the
ability to reset the CPTR 1, CPTR 2, and the Stabilator to the automatic mode through RESET
switches. Stabilator MAN SLEW (UP, OFF, DOWN) switch and TEST button are also located on
the panel.
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3.
Learning Step/ Activity: Describe the operational characteristics and components of the stabilator
system.
a. Description: The stabilator system drives a variable angle of incidence airfoil which enhances
handling qualities in the pitch axis.
(1) The stabilator is a three section stabilator. It is comprised of two outboard foldable sections
and a center box. The panels can be folded upright for ease of shipping the helkicopter
during deployments. The stabilator center box provides the mount to the tail pylon. The
outboard stabilator panels consist of aluminum hinge fittings, integral composite ribs, end
ribs, leading edge, tip cap, composite upper and lower skins, and internal spars. The
outboard stabilator panels provide provisions for the mounting of electrostatic discharge
wicks, IR and formation lighting, balance weights, hand holds, and hinge fittings. The FMS
provide a means of status indication, initiation of Built-In Testing (BIT), stabilator rigging,
parameters, and FCC testing.
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(a) “STAB UNLOCKED” caution will be indicated when outboard panels are not locked into
flight position.
(b) The folding stabilator is equipped with locking pin switches. These switches are to
provide an indication of unlocked condition. They are one of the interlocks that allow the
stabilator system to initiate the Auto Mode upon initial power application, or after pressing
the AUTO CONTROL RESET switch after a failure, to reengage the auto mode. If either
switch opens for any reason a STAB UNLOCKED caution will illuminate on the EICAS
display.
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Stabilator Hinge Pins
Stabilator Lock Pins
Stabilator Lighting Connector
Keeper
Stabilator Lock Pin Switch
(c) Lock pin engagement can be visually checked. The lock pin switches must be rigged to
ensure proper engagement and secured with safety wire. The spring loaded keeper
should be visually checked to ensure it fully engages the nut.
(d) Stabilator Position Transmitter/Limit Switch assembly – The position transmitter/limit
switch assembly, located in the tail pylon, is mechanically connected to the stabilator by a
rod. The position transmitter assembly is a 3-wire synchro device that supplies a position
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signal to the FCCs and to the DCUs for display on the PFD. The limit switches should
prevent the stabilator from contacting the fuselage in any mode of operation.
(e) Stabilator Actuators- The stabilator is moved using two electromechanical actuators in
response to collective position, lateral acceleration, airspeed, and pitch rate inputs.
1) No. 1 and No. 2 Stabilator Actuators – Two identical actuators, mounted back to back
and located in the tail pylon assembly are used to position the stabilator. Each
actuator contains an electric motor which is geared to a jackscrew, limit switches, and
a position potentiometer. The potentiometer provides actuator position to each FCC.
The actuators extend or retract as necessary to position the stabilator in accordance
with outputs from the FCCs or manual slew signals.
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2) Cyclic-mounted, Stabilator Slew-Up Switches – Cyclic-mounted, Stabilator Slew-Up
switch provides the ability to manually slew the stabilator up. Auto mode will be
disengaged as soon as this switch is moved.
a)
As a result, the STAB MANUAL MODE caution will illuminate on the EICAS
display and a beeping tone will be heard in the headsets.
(f) Stabilator Control/Auto Flight Control Panel – The following switches are located on the
AFCS Control Panel:
1) The Manual Slew Switch (Up, Off, Down) Triangular shaped three position switch
indented so the user cannot accidently move the switch. This switch will manually
slew the stabilator up or down.
a) Use of the switch, when the automatic mode is engaged, will disengage the
automatic mode.
b) As a result, the STAB MANUAL MODE caution will illuminate on the EICAS
display and a beeping tone will be heard in the headsets.
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2) Stabilator Test Switch - A TEST pushbutton provides a check of the automatic mode
fault detector feature.
a) The No. 2 FCC commands the audible warning tone for a short duration, and
then the No. 1 FCC commands it for a short duration.
b) The stabilator drives to approximately 31˚ trailing edge from the full down position
of 39˚ down, using both actuators (a miscompare is not induced).
c) Both FCCs go into manual mode and command a constant beeping warning
tone, requiring the pilot to press the AUTO CONTROL RESET pushbutton to
return to the auto mode.
d) The test can only be performed when the helicopter is weight-on-wheels and Nr <
10%.
3) Auto Control Reset Button- The auto control reset button is used to reset the
stabilator to the automatic mode following a malfunction, a test procedure or when
the stabilator is in the manual mode. Should a malfunction occur in the stabilator, the
automatic mode will disengage. Upon automatic disengagement or manual slew
switch activation, the AUTO CONTROL RESET pushbutton ON legend will
extinguish.
4) When the malfunction has been removed, the set may be reset by pressing and
releasing the AUTO CONTROL RESET pushbutton switch. The ON legend
illuminates and remains on as long as there are no longer any malfunctions.
g) The stabilator has three modes of operation:
1) Stabilator angular position is displayed in the upper left corner on the MFDs PFD. It
provides an angular indicator and digital readout. The digital readout is displayed
above the indicator and ranges between -10 (up) and 45 (down). If the stabilator
angle exceeds these limits, the indicator remains parked at the nearest limiting value.
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2) The FCCs are designed to detect mode failures and display the appropriate condition
on the MFD PFD and EICAS displays. By communication over the Cross Channel
Data Link (CCDL), each FCC will know the status of the other. Should an FCC
receive signals which fall outside programming parameters, it will independently
command the system to the Degraded or Manual Mode.
a) AUTO –automatically positions the stabilator to the best angle of attack for the
existing flight conditions. Once the stabilator engages in the automatic mode, no
further pilot action is required for stabilator operation. In the auto mode, the
stabilator position indicator and digital readout is white. Use of the manual slew
switch or the cyclic mounted manual slew-up switch, when the automatic mode is
engaged, will disengage the automatic mode. As a result, the STAB MANUAL
MODE caution legend on the EICAS display will illuminate and a beeping tone
will be heard in the headsets.
b) LIMITED AUTO MODE- If a signal from one of the collective stick position
sensors, lateral accelerometers, or pitch rate data is lost, or if a failure occurs
preventing an FCC from moving its associated actuator the stabilator reverts to
the Limited Automatic Mode. As long as the FCC can identify the position of the
failed actuator, the other FCC will continue to program its actuator automatically,
but in a Limited Auto Mode. In the Limited Auto Mode, the stabilator position
indicator and digital readout is white. In this situation, Stabilator programming will
only be half of its normal maximum rate due to the loss of the other actuator. The
available range of stabilator movement will depend on the position of the actuator
on the failed side; about 35˚ if failure occurs full down, or about 30˚ if failure
occurs full up. If for any reason the FCC cannot identify the position of the failed
actuator, the result will be Manual Mode.
1
The STAB DEGRADED advisory legend on the EICAS display will illuminate,
but no beeping tone will be heard in the headsets. The AUTO CONTROL
RESET push button ON legend will remain on.
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2
If the stabilator stops operating in the automatic mode, the Vne becomes
variable to indicate the airspeed limit by extending the red arc in a counterclockwise direction to the applicable airspeed for the current stabilator
position. Pointer and readout turn red with a black outline when the airspeed
is at or above Vne.
c) STAB MANUAL MODE – Manual mode will occur when neither channel can
operate in Auto Mode. If the automatic mode disengages and cannot be reset
due to a malfunction, the MAN SLEW switch is used to manually position the
stabilator. The Stabilator may or may not move in response to direct pilot input.
STAB MANUAL MODE caution will illuminate on the EICAS display and the
Stabilator position indicator and digital readout will be displayed in yellow. The
red Vne indication, normally placed at 193 KIAS, will adjust according to the
manually adjusted Stabilator position. Manual Mode will occur under the following
circumstances:
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1
Any ADC airspeed sensor becomes invalid.
2
Miscompare of both ADCs airspeeds.
3
FCC failure.
4
Stabilator actuator position becomes unknown.
5
Stabilator Position Transmitter and Limit Switch Assembly fails.
6
Pilot or copilot chooses to override Auto Mode.
b Manual control available:
1) The stabilator shall be set full down at speeds below 40 KIAS.
2) The stabilator shall be set at 0° at speeds above 40 KIAS.
3) Autorotation airspeed shall be limited to 120 KIAS
c Manual control not available:
1) The Vne becomes variable to indicate the airspeed limit by extending
the red arc in a counter-clockwise direction to the applicable airspeed
for the current stabilator position. Pointer and readout turn red with a
black outline when the airspeed is at or above is at or above the Vne
d) When the stabilator input status has failed, the stabilator indicator and digital
readout are removed and a red STAB flag will be displayed in lieu of the angular
and digital readout.
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b. The stabilator will program according to (5 functional characteristics (SCALP)) while in Auto or
Limited Auto Modes of Operation.
(1) Streamlines - Align stabilator and main rotor downwash in low speed flight to minimize nose
up attitude resulting from rotor downwash.
(2) Collective coupling- 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) Decreasing angle of incidence- Angle of incidence will decrease as airspeed increases to
improve static stability.
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(4) Lateral to sideslip to pitch coupling- 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) Pitch rate feedback- Provide pitch rate feedback to improve dynamic stability. The rate of
pitch attitude change of the helicopter is sensed by the EGI’s and a signal is sent to each
FCC. The FCC’s determine the position the stabilator will move 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.
(6) Signals used when the stabilator is in AUTO/Limited Auto Mode (CLAP).
(a) Collective signal from the No. 1 and No. 2 Collective Stick Position Sensors.
(b) Lateral Acceleration from the No. 1 and No. 2 Lateral Accelerometers.
(c) Airspeed signal from the No. 1 and No. 2 ADCs.
(d) Pitch Rate signal from the No. 1 and No. 2 EGIs.
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1) No. 1 and No. 2 Collective Stick Position Sensors – located on the mixer assembly,
reflect the stabilator position by generating a signal proportional to the collective stick
position.
2) Lateral Accelerometers – The No. 1 and No. 2 accelerometers, located in the cabin
overhead, produce an output signal proportional to the helicopters lateral acceleration. Each
signal is fed to its related FCC where it is conditioned. The lateral acceleration signal is used
to position the stabilator to compensate for sideslip to pitch coupling.
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3) No. 1 and No. 2 ADCs – The ADCs are connected to the Pitot-Static system and
provide ARINC serial data that contains information of the helicopter airspeed,
altitude, and altitude rate. The stabilator system only uses the airspeed data. Each
FCC uses its respective onside ADC for airspeed data. The ADCs are located on
either side of the aircraft just forward of the gunner windows.
4) No. 1 and No. 2 EGIs – The No. 1 and No. 2 EGIs provide helicopter pitch rate
information via ARINC serial data lines to the No. 1 and No.2 FCCs respectively. The
pitch rate data is used by the stabilator system to enhance the AFCS system’s ability
to correct short term pitch disturbances.
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4. Learning Step/Activity: Describe the operational characteristics of the Stability Augmentation System
(SAS).
a. Description: Stability Augmentation System (SAS) – The SAS responds to short-term
aerodynamic disturbances and effectively dampens any helicopter movement. Since response to
disturbances is almost instantaneous, the SAS control is limited in authority (amplitude) to
prevent SAS malfunctions from causing undesirable helicopter response before the pilot has a
chance to react and take control.
b. The purpose of SAS is to enhance the dynamic stability of the aircraft in the pitch, roll, and yaw
axis. Both SAS 1 and SAS 2 are controlled by their respective FCC based off of signals from
sensors throughout the aircraft.
c.
SAS 1 and SAS 2 both have a maximum of 5% control authority over the aircraft for a maximum
of 10% when working together. If one SAS should be turned off during flight the other SAS is able
to double its gain for a maximum control authority of only 5 %. Control authority is defined as the
amount of input a system can make to the rotor system compared to how much the pilot moves
the controls.
(1) For example: If SAS 1 is working at 3% authority and SAS 2 is working at 3% authority, then
they are working at a total of 6% control authority. If you should happen to turn off SAS 2,
then SAS 1 can double its gain up to a maximum of 5% control authority.
d. FCCs. The Flight Control Computers (FCC) control the inputs to the SAS actuators by receiving
signals from sensors throughout the aircraft. There are no cross-side sensor comparisons for
SAS inputs to the actuators. Each FCC sends their own signal to the actuators based off of the
signals from the respective sensors.
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(1) Sensors. SAS works in the Pitch, Roll and Yaw axis. Various sensors throughout the aircraft
will pick up the dynamic disturbances and the signals are sent to the FCC’s for processing.
The FCC’s determine how to correct for the dynamic movement. FCCs receive their
respective EGI pitch, roll and yaw rate signals.
(2) SAS1 and SAS2 enhance turn coordination by deriving aircraft accelerations from the lateral
accelerometers which together with roll rate signals, are computed by the FCCs and sent to
their respective yaw channels automatically at airspeeds greater than 50 knots. Longitudinal,
lateral, and vertical accelerations are provided by individual accelerometers in each axis.
(a) Two additional features of SAS are hover augmentation and gust alleviation.
(b) Hover augmentation reduces pilot workload to maintain a hover. Hover augmentation is only active
at airspeeds less than 50 knots.
(c) Gust alleviation improves disturbance response by responding to changes in pitch and roll attitude.
This mode is most effective when used in conjunction with FPS attitude hold.
(3) Malfunctions. A malfunction of the SAS system may be detected by the pilot as an erratic
motion in the helicopter without a corresponding failure advisory indication.
(a) If the malfunction is of an intermittent nature, the indication can be cleared by pressing
FAILURE RESET CPTR 1 and/or FAILURE RESET CPTR 2 switches.
(b) If the malfunction is continuous, the appropriate SAS should be turned off, in which case
the SAS DEGRADED advisory will appear on the EICAS display.
(c) SAS actuator hydraulic pressure is monitored and in case of loss of actuator pressure, or
if both SAS 1 and SAS 2 are turned off, the SAS OFF caution will appear on the MFD. In
the latter case, FPS will automatically turn off and the FPS ON segment light will
extinguish.
(d) The SAS DEGRADED advisory will also appear when the pitch, roll and/or yaw SAS
channels from SAS 1 or SAS 2 has disengaged due to a fault.
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5.
Learning Step/ Activity: Describe the operational characteristics of the Trim System.
a. Description: Trim is comprised of four actuators; the collective, pitch, roll, and the yaw trim
actuators.
b. When TRIM is engaged on the control panel, the collective, pitch, roll, and yaw trim actuators are
activated to maintain position of the cyclic, collective, and tail rotor controls. When FPS is
engaged, Trim performs basic autopilot functions in the pitch, roll, and yaw axis by trimming the
flight controls to the position required to maintain the selected flight attitude. When the Flight
Director is coupled (and desired modes captured), Trim positions the flight controls as required to
maintain the aircraft within the desired coupled flight parameters. Trim is capable of driving the
flight controls throughout their full range of travel (100% authority) at a limited rate of 10% per
second.
(1) The collective, pitch, roll, and yaw trim servos are located on the hydraulic deck.
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c.
Operational Aspects of Trim
(1) The Trim systems are engaged with the TRIM button located on the Stabilator Control/Auto
Flight Control Panel.
(2) Trim is implemented solely by the No. 2 FCC through three independent collective, roll, and
yaw electro-mechanical actuators and an independent pitch electro-hydraulic servo/actuator.
Proper operation of yaw trim requires that SAS/BOOST on the Stabilator Control/Auto Flight
Control Panel be on and operational.
(a) The lateral cyclic (roll), tail rotor (yaw), and collective stick forces are developed through
electro-mechanical Trim servos. Longitudinal (pitch) forces are developed through an
electro-hydraulic servo actuator in conjunction with inputs from the FCCs. With TRIM on,
the Trim subsystem provides gradient and detent holding force for collective, pitch, roll,
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and yaw. When turned off, the Trim subsystem is released and light cyclic control forces
are present.
(b) When the pilot applies a longitudinal or lateral force to the cyclic stick with TRIM
engaged, a combination detent and gradient force is felt. The pilot normally removes the
force by pressing the thumb-operated TRIM REL switch on the pilot/copilot cyclic grip.
The TRIM switch on the control panel may also be used to turn Trim off entirely.
(c) In addition to the TRIM REL switch, a four-way trim switch on each cyclic stick
establishes a Trim position without releasing Trim. With TRIM engaged, the Trim position
is moved in the direction of switch movement. The cyclic is moved by the switch in one
direction at a time; diagonal movements, such as forward and left simultaneously, are not
possible. When FPS is engaged (FPS ON illuminated), the Trim switch changes the pitch
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and roll attitude reference instead of the cyclic stick position reference. The Trim release
feature permits the pilot or copilot to fly the helicopter with light stick forces.
(d) The pedal gradient maintains pedal position whenever Trim is engaged. The pedal
switches provide a means of disengaging yaw Trim. At airspeeds below 50 knots, the
pilot can change the heading reference by pressing the pedal switches as the yaw pedals
are moved, positioning the helicopter to the desired heading then releasing the pedal
switches.
1) When the pedal switches are pressed, the 28 VDC yaw Trim engage signal to the
clutch mechanically disengages and the gradient force is removed from the flight
control linkage. The pedals may then be moved to the desired position and released.
The pedals will then be held at this new position by the Trim gradient setting the new
helicopter heading reference.
2) The pedal Trim gradient actuator also includes a pedal damper. The pedal damper is
engaged continuously, independent of electrical power and the TRIM switch on the
control panel.
(e) When the pilot applies an up or down force to the collective stick with TRIM on, a gradient
force is felt. The collective gradient maintains collective stick position whenever the Trim
is engaged. By pressing the TRIM REL trigger switch on the pilot/copilot collective, the
gradient force is removed. The collective stick may then be moved to the desired position
and TRIM REL trigger switch released. The collective sticks will be held at this position
by the Trim gradient. In addition to the collective TRIM REL trigger switch, an additional
four-way Trim switch on each collective stick establishes a Trim position without releasing
Trim. TRIM DN or TRIM UP increases the collective setting down or up. TRIM L and
TRIM R positions for yaw pedal trim heading adjustments below 50 KIAS. Above 50
KIAS, TRIM L and TRIM R will place the aircraft in a turn up to a standard rate of turn.
The L and R inputs are disabled for airspeeds greater than 50 KIAS with the flight director
coupled and heading hold selected.
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1) All trim actuators incorporate slip clutches to allow pilot and copilot control inputs if an
actuator should jam. The maximum forces required to break through the clutch are 80
pounds in yaw, 13 pounds in roll, and 22 pounds in collective.
2) The four trim servo actuators are specific in design for their respective axis and are
not interchangeable.
FPS FAIL
(f) Operational and ground/flight checks, and emergency procedures associated with the
Trim system.
1) Operation of the Trim system is continuously monitored by the FCCs. If a malfunction
occurs, the FCCs, (through FCC2) shut off the servo/actuator driving the affected
axis, and the TRIM FAIL and FPS FAIL legends are displayed on the EICAS display.
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If the malfunction is of an intermittent nature, the indication may be cleared by
pressing the FAILURE RESET CPTR 2 switch on the control panel.
2) AFCS status (and subsequent Trim system status) can be monitored from the AFCS
STATUS (X/Y) page accessed through the FFK STS and then SK-7 AFCS keys.
6.
Learning Step/ Activity: Identify the operational characteristics of the Flight Path Stabilization
System (FPS).
a. The yaw axis of the FPS provides heading hold at airspeeds less than 50 knots and heading hold
or turn coordination at airspeeds greater than 50 knots. For heading hold operation at airspeeds
less than 50 knots, the helicopter is maneuvered to the desired heading with feet on pedals.
When trimmed at the desired heading, the pilot may remove feet from pedals, at which time the
existing heading becomes the reference, which is automatically held. To change heading, the
pilot may activate one or both pedal switches, trim to the desired heading, and remove feet from
pedals. Alternatively, below 50 KIAS the 4-way collective TRIM beeper may be used to slew the
reference heading to the desired heading.
b. At airspeeds greater than 50 KIAS heading hold will be automatically disengaged and turn
coordination engaged under these conditions:
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(1) TRIM beeper is actuated in the lateral direction.
(2) TRIM REL switch is pressed and roll attitude is greater than 2.5 degrees.
(3) About 1⁄2 inch cyclic displacement and a roll attitude of about 1.0°.
(4) Heading hold is automatically re-engaged and turn coordination disengaged upon recovery
from the turn when the lateral stick force, roll attitude, and yaw rate are within prescribed
limits.
c.
The FPS monitoring is automatic. If a malfunction is detected, the FPS FAIL legends on the
EICAS display shall go on and FPS will either continue to operate in a degraded mode, such as
without heading hold, or may cease to function altogether. In this situation, the pilot must take
over manual flight of the helicopter and may either turn the mode off or evaluate performance to
determine the degree and type of degradation and continue flight with the remaining features.
d. To help evaluate the nature of the degradation, failure advisory legends are displayed on the
AFCS STATUS page of the Flight Management System (FMS). These status indications show
the pilot which mode has experienced the failure. If a CPTR 1 or CPTR 2 light illuminates, the
pilot may attempt to clear the indication of temporary malfunctions by pressing the illuminated
CPTR switch on the AUTO FLIGHT CONTROL panel. If the FPS FAIL caution on the EICAS
display extinguishes and the CPTR light extinguishes, it may be assumed that normal operation is
restored.
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(1) The overall status of the AFCS and FCCs can be obtained from the FMS through the AFCS
STATUS (x/y) page. The page can be accessed by pressing the FFK STS and the SK-7
AFCS keys.
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(3) Individual FCC statuses are capable of being displayed and can be obtained through the
FMS FCC-x STATUS (1/x) page by pressing the FFK STS, SK-7 AFCS, and then SK-1
FCC1 or SK-6 FCC2 keys.
(4) Testing of the AFCS (and FCCs) can be accomplished with the FMS’ through the AFCS
TEST (1/x) page. The test can be initiated by pressing the FFK TST, SK-7 AFCS, and then
pressing SK-1 TEST key. Overall test results will be displayed on this screen. A test can be
performed on each individual FCC through the FMS FCC-x TEST (1/x) page by pressing the
FFK TEST, SK-7 AFCS, and then SK-7 FCC1 or SK-8 FCC2 keys.
e. Attitude Hold
(a) Attitude hold can be engaged with FPS at all airspeeds and is commanded by changing
the cyclic stick position with the TRIM switch. This causes the cyclic stick to move and
the helicopter attitude to change approximately 5° per second in pitch and approximately
6° per second in roll.
(b) When cyclic movement is stopped, the autopilot stabilizes the helicopter at the new
attitude. The roll channel autopilot holds roll attitude of the helicopter. Attitude
information is supplied to the computer from the No. 1 and No. 2 EGIs.
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f. Heading Hold
NOTE
Heading hold will capture in a hover when yaw rate reaches zero, not when pedal
switches are engaged.
(a). The yaw channel of the autopilot provides the heading hold feature for hover and forward
flight and is engaged when ON is illuminated on the FPS switch. Heading hold is an outer-loop
function operating through the yaw trim actuator, and therefore will only be operational when the
yaw trim is engaged. Releasing all pedal switches at a given heading synchronizes the trim
system to the established heading. The yaw autopilot also uses a collective stick position sensor
to hold reference heading for yaw excursions caused by main rotor torque changes.
(b), heading hold is reengaged following a turn when the following conditions are maintained for 2
seconds:
Helicopter roll attitude is within 2° of wings level.
Yaw rate is less than 2° per second. The WOW switch disengages the heading hold
when the helicopter is on the ground. During coupled flight operation above 50 KIAS,
heading hold is achieved by the use of the roll axis (Trim Release).
g. Turn Coordination. Automatic turn coordination is provided at airspeeds greater than 50 knots. Turn
coordination allows the pilot to fly a coordinated turn with directional control provided by the AFCS. The
AFCS uses lateral acceleration and roll rate to determine if the helicopter is out of balanced flight, and
provides the yaw SAS and yaw trim with the inputs necessary to maintain an automatic coordinated turn.
Automatic turn coordination is engaged and heading hold disengaged when roll attitude is greater than 1°
for any of the following conditions:
a. Lateral cyclic greater than 1⁄3 inch stick displacement from trim.
b. Cyclic TRIM REL is pressed.
c. Roll attitude is beeped beyond 2.5° bank angle. Actuation of the
Collective TRIM switch above 50 KIAS uncoupled in the lateral axis
For more than one second provides a 1° per second coordinated turn.
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Flight Director (FD)
Provides pitch, roll, and collective steering cues on the PFD.
All FD modes are engaged by the operator using either flight director/display control
panel (FD/DCP).
The coupled Flight Director set consists of two independent FCCs and Flight
Director/Display Control Panels
.
When either FCC is operating in the coupled mode, it provides pitch, roll, and
collective commands to the AFCS to drive the trim servo actuators for autopilot flight.
If one or more of the FD functions have disengaged, the FLT DIR FAIL caution will
appear.
The FD COUPLE FAIL caution appears when a coupled FD automatically decouples
due to a failure in the FD or related subsystems.
Check on Learning
1.
What are the five basic subsystems that comprise AFCS?
2. Dynamic stability is provided by?
3. Basic auto pilot functions are provided by?
4. The ______________ caution indicates the majority of AFCS has shut
down.
5.
What does the SAS DEGRADED advisory mean?
6.
What does the STAB DEGRADED advisory mean?
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7. To attempt to clear AFCS fault indications, the pilot can_________.
8. The collective trim beeper can be used below 50 KIAS to ________?
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