FLIGHT LINE
SUPPLEMENT
UH-60L
1-212th Aviation Regiment
Fort Rucker, Alabama
Revised January 2014
UH-60A/L FLIGHT LINE SUPPLEMENT
TABLE OF CONTENTS
Fire Protection Systems .............................................................................................................................. 1
T700 Engine ................................................................................................................................................ 3
Hydromechanical Unit (HMU) ..................................................................................................................... 5
Overspeed and Drain Valve (ODV) 701C 701D/CC................................................................................... 7
Engine Alternator 701C 701D/CC ............................................................................................................... 8
Digital Electronic Control (DEC) 701C 701D/CC ...................................................................................... 9
Fuel System .............................................................................................................................................. 10
Mechanical Mixing Unit (MMU) ................................................................................................................. 13
Collective/Airspeed to Yaw (Electronic Coupling) .................................................................................... 13
Torque Compensation ............................................................................................................................... 13
Automatic Flight Control System (AFCS) ................................................................................................. 15
Stability Augmentation System (SAS) ...................................................................................................... 16
Trim System .............................................................................................................................................. 17
Flight Path Stabilization (FPS) .................................................................................................................. 18
Stabilator System ...................................................................................................................................... 20
Hydraulic System ...................................................................................................................................... 21
Pneumatic System .................................................................................................................................... 27
Powertrain System .................................................................................................................................... 28
Main and Tail Rotor Groups ...................................................................................................................... 30
Utility Systems ........................................................................................................................................... 32
Electrical Power Supply System ............................................................................................................... 33
Electronic Navigation Instrument Display System .................................................................................... 35
Integrated Vehicle Health Management System ...................................................................................... 39
Attitude Heading Reference Unit (AHRU) ................................................................................................ 41
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FIRE PROTECTION SYSTEMS
Fire detection and fire extinguishing systems are installed so that a fire may be detected and put out at
either engine or the APU, without affecting the remaining two.
Detection
The system consists of five radiation-sensing flame detectors. Two detectors are installed in
each main engine compartment and one detector in the APU compartment.
• In case of fire, the detectors react to the infrared radiation and send a signal which causes the
proper fire extinguisher arming lever: #1 ENG EMER OFF, #2 ENG EMER OFF, or APU Thandle to appear. Also, the master FIRE warnings will appear.
• The detector system automatically resets itself with warnings disappearing when the infrared
radiation source ceases to emit.
Extinguishing
A high rate discharge extinguishing system provides a two-shot, main and reserve capability to either
main engine compartment or the APU compartment. Both containers have dual outlets, each with its
own firing mechanism.
• The system has a thermal discharge safety port that will cause a visual indicator on the right side
of the fuselage to rupture, indicating that one or both containers have discharged overboard.
• Electrical power to operate the No. 1 main and No. 2 reserve outlet valves is by the No. 2 dc
primary bus. Power to operate the No. 2 main and No. 1 reserve fire extinguisher container outlet
port valves and the directional control valve is by the battery utility bus.
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WARNING
In case of fire when ac electrical power is not applied to the
helicopter, the reserve fire extinguisher must be discharged. Fire
extinguisher agent cannot be discharged into the No. 2 engine
compartment if ac electrical power is not applied to the helicopter.
The FIRE EXTGH switch has marked positions RESERVE-OFF-MAIN. The switch is operative only
after one of the t-handles has been pulled. The contents of the fire extinguisher container
discharge into the compartment of the last lever pulled.
A crash actuated system is part of the fire extinguisher system. An omnidirectional inertia switch is
hard-mounted to the airframe to sense crash forces.
• Upon impact of a crash of 10 Gs or more, the switch will automatically fire both fire extinguishing
containers into both engine compartments.
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T700 ENGINE
The T700 engine is a front drive, turboshaft engine of modular construction. The engine is divided into
four modules: cold section, hot section, power turbine section, and accessory section.
Cold Section Module: Includes the inlet particle separator (IPS), the compressor, the output shaft
assembly, and line replaceable units (LRUs).
• The IPS removes foreign material from the engine inlet air. Engine inlet air passes through the swirl
vanes, directing air into a rotating pattern. Centrifugal force carries foreign material to the outer
section and is collected in the scroll case. Particles are drawn from the scroll case by the enginedriven blower and discharged through the aircraft discharge duct.
• The compressor has five axial stages and one centrifugal stage that continuously pressurizes the
combustion chamber, cools the hot section components, and provides bleed air from three bleed
ports at the 5th stage of the compressor case. One bleed port provides air to the anti-ice start bleed
valve and remaining two ports provide air to the pneumatic system and engine inlet anti-ice.
• The output shaft assembly consists of a forward support that houses the high speed shaft, which
connects the input module to the main transmission.
• LRUs mounted on the cold section module are the 700 electrical control unit (ECU) or 701C
701D/CC digital electronic control (DEC), anti-ice/ start bleed valve (AISBV), 700 history recorder or
701C 701D/CC history counter, ignition system, and electrical cables.
Hot Section Module: Consists of three subassemblies: the gas generator (GG) turbine, stage 1 nozzle
assembly, and combustion liner. LRUs on the hot section are 701C 701D/CC ignitors and 700 primer
nozzles and ignitors.
Power Turbine Section: Includes a two-stage power turbine, exhaust frame, drive shaft assembly, and
the C-sump housing. The LRUs on the power turbine section are the thermocouple harness, torque
sensor/overspeed sensor, and the Np (% RPM1 or 2) sensor.
• The power turbine converts gas pressure into rotational shaft energy to power the drive train.
• The seven thermocouple probes measure the temperature of the gases between the Ng and Np
turbine sections. The probes feed temperature data to the 700 ECU and 701C 701D/CC DEC for
the TGT limiting function and the TGT TEMP signal to be displayed on the central display unit
(CDU).
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Accessory Module Section: Includes the top mounted accessory gear box (AGB) and a number of
LRUs. The LRUs are the hydromechanical unit (HMU), 700 pressurizing and overspeed unit (POU)
and 701C 701D/CC overspeed and drain valve (ODV), starter, engine driven fuel boost pump, oil
filter, oil cooler, alternator, oil and scavenge pump, IPS blower, fuel filter assembly, chip detector, oil
filter bypass sensor, liquid to liquid cooler, oil temperature sensor, fuel pressure sensor, and oil
pressure sensor.
• LRUs on the front of the AGB.
• LRUs on the back of the AGB.
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HYDROMECHANICAL UNIT (HMU)
The HMU is mounted to and driven by the AGB. Various parameters are sensed by the HMU and
influence fuel flow, variable geometry position, and engine AISBV operation. Fuel from the HMU flows
to the 700 POU or 701C 701D/CC ODV.
The functions of the HMU are as follows:
• High pressure fuel pump - Delivers fuel to the combustion chamber. The high pressure provides a
precise spray pattern for good atomization.
• Fuel flow metering/scheduling – The three primary HMU functions affecting adjustment of fuel flow
via the metering valve are as follows:
LDS input – The load demand schedule is manually adjusted by collective movements through the
LDS. The collective link to this schedule proactively attempts to match engine power output to the
main rotor load. This attempt is approximate and must be further adjusted by the ECU/DEC, via the
torque motor.
ECU/DEC input – The torque motor acts as an interface between the ECU/DEC and the load
demand schedule of the HMU. ECU/DEC trimming of the HMU reduces pilot load by electrically
keeping Np and Nr constant (Np governing), keeping power outputs closely matched (load sharing)
and preventing TGT from reaching destructive levels (TGT limiting).
PAS input – The maximum power available schedule is set by the PCL through the PAS. The PAS
sets the maximum power limits that may never be exceeded by collective movement.
• Additional fuel flow metering features of the HMU are as follows:
Acceleration/deceleration fuel flow limiting – An automatic mechanical function that prevents
flameout due to rapid power changes (power changes called upon by PAS, LDS, or ECU/DEC).
This limits the metering valve movement rate to prevent compressor stalls and engine flame out.
PAS override function with ECU/DEC malfunction - The pilot may sever the ECU/DEC to HMU
interface by locking out the torque motor. Advancing the PCL to the lockout position (ECU/DEC
LOCKOUT) mechanically blocks off the torque motors fuel control passage within the HMU. The
three ECU/DEC functions have been locked out.
Ng limiting – There are two types of Ng limiting that limit the overall Ng based on environmental
conditions.
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The HMU can either mechanically compute/set the maximum Ng speed (Ng governing) or it
induces a physical limitation in which the metering valve is full open (maximum fuel flow). The
type of limiting that is in effect depends on engine inlet temperature.
Ng is limited to 103% Ng and in extreme cold conditions, less Ng is allowed. Ng governing may be entered
prior to TGT limiting.
Ng shutdown is the engine shutdown (flameout) at 110% Ng to prevent destruction on Ng section
and occurs when a mechanical Ng speed sensing mechanism (centrifugal flyweights) stops all
fuel to the engine when Ng speed reaches 110%.
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• HMU functions that do not affect fuel flow metering/scheduling are as follows:
Variable geometry positioning – A variable geometry actuator, on the HMU, varies the angle of
attack of the inlet guide vanes, the variable stator vanes, and opens and closes the AISBV. These
actions permit optimum performance over a wide range of operating conditions.
Vapor vent valve – Allows the pilot to manually prime the fuel system passages from the fuel cell to
the HMU. See Chap 8 of the -10 for operation in cold conditions and manual prime procedures.
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OVERSPEED AND DRAIN VALVE (ODV) 701C 701D/CC
The ODV controls the sequencing of fuel for the engine, fuel purge functions, and also the flow schedule
for overspeed protection.
The functions of the ODV are as follows:
• Sequences fuel through the main fuel manifold to the injectors for starting acceleration and engine
operation.
• Purges fuel from the main fuel manifold on shutdown to prevent coking.
• It shuts off fuel flow to prevent an engine overspeed (120% ±1%) when the overspeed system is
sensed by the DEC.
• Shuts off fuel to prevent hot starts when activated by the hot start preventor (HSP).
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ENGINE ALTERNATOR 701C 701D/CC
Provides the same basic functions as the 700 engine alternator, except for the following
additions/modifications:
When the alternator Ng signal is interrupted, a loss of Ng cockpit indication will occur with
a corresponding ENG OUT warnings and audio (Ng is >55%).
When the alternator power supply to the DEC is interrupted, 400 Hz 120 vac aircraft power is utilized
to prevent engine (high side) failure.
A complete loss of engine alternator power results in:
• ENG OUT warning and audio will occur.
• There will be no loss of associated % RPM 1 or 2, % TRQ indications, because the DEC can utilize
400 Hz 120 vac aircraft power.
• Overspeed protection is still available.
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DIGITAL ELECTRONIC CONTROL 701C 701D/CC
The DEC replaces the analogue ECU. The DEC contains a micro-computer processor and provides the
same basic functions of the ECU, except the following additions and modifications:
The DEC can be fully powered by either the engine alternator (primary) or by the 400 Hz, 120 vac
(alternate) aircraft power.
The DEC provides the following signals:
• The initial Np indication is at 4,000 RPM vs 6,000 RPM with the ECU.
• Torque signal is locked at zero until Np reaches 35%, which eliminates the torque spike, on the
PDU, during engine start and shutdown.
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• TGT has a 71 C bias (the engine is running hotter than indicated on the CDU).
Signal fault validation.
• Signals are continuously validated when the engine is operating at idle or above. If a failure
occurred, the failed component or related circuit will be identified by a preselected fault code
displayed on the engine torque meter digital readout.
• Fault codes are displayed for 4 seconds on, 2 seconds off, starting with the lowest code and
rotating through all applicable codes.
• Codes will be displayed 30 seconds after both engines have been shut down with 400 Hz, 120 vac
power applied.
• Pilots can suppress the fault code display of an engine by depressing the associated cockpit
overspeed test button (TEST A/B).
DIAGNOSTIC INDICATION
ON TORQUE METER (±3%)
15%
SIGNAL FAILED
DEC
Np Demand Channel
25%
Load Share Channel
TGT Channel
35%
45%
Alternator Power
55%
Ng Channel
65%
Np Channel
75%
Torque and Overspeed Channel
85%
Hot Start Prevention Channel
95%
Aircraft 400 Hz Power
105%
Collective Channel
115%
Nr
125%
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The TGT limiting system limits fuel flow when the TGT TEMP reaches the dual engine 10 minute
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limiting value of approximately 866 C. The automatic contingency power limiting will switch to a
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higher single engine 2 ½ minute temperature limiting value of approximately 891 C when the
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opposite %TRQ is less than 50%. It is normal to see a transient increase above 903 C TGT TEMP
when pilot demands maximum power.
The Np overspeed system is set to trigger at 120% ±1 % RPM 1 or 2 and will result in a fuel flow
shutoff causing the engine to flame out. When % RPM is below the limit, fuel flow is returned to the
engine and engine ignition will come on to provide a relight.
Hot Start Prevention (HSP): Prevents overtemperature during engine starts.
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• When Np and Ng are below their respective hot start references and TGT TEMP exceeds 900 C,
the system activates a solenoid in the ODV. This shuts off fuel flow and causes the engine to shut
down.
• The pilot can disable the HSP for emergency starts by pressing and holding the overspeed test
button (TEST A/B) for the engine being started during the start sequence.
Transient Droop Improvement (TSI): Provides significant droop improvement during some
maneuvers and is 4:1 over the 700.
• Engines are more responsive to aircraft maneuvers by providing a collective and Nr rate
compensation.
• Signals required: Adds Nr sensor located on left accessory module, new collective position
transducer in the mixing unit, and engine torque signal.
The history counter replaces the history recorder.
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FUEL SYSTEM
A separate suction fuel system is provided for each engine and supplies fuel to both engines and the
APU. The engines and APU are suction fed. The APU is fed from the #1 main fuel tank by a separate
fuel line.
The fuel system consists of two interchangeable, crashworthy, ballistic-resistant (7.62 mm) tanks, self
sealing fuel lines (except APU), firewall-mounted selector valves, prime/boost pump and fuel tanks,
engine-driven pumps, quantity gauging and low level warning system.
Each fuel system has a selector valve which is manually operated through the ENG FUEL
SYS selector lever. Each lever can be actuated to three positions: OFF, DIR, and XFD.
• OFF – The control valves are closed, allowing no fuel into the engines.
• DIR – The selector valves are opened, providing fuel flow for each engine from its individual fuel
tank.
• XFD – This connects the engine to the other fuel tank through the crossfeed system.
Fuel is drawn from the main tank up through the fuel selector valve and proceeds through the low
pressure suction engine driven pump, fuel filter, HMU, liquid to liquid cooler, the 700 POU or 701C
701D/CC ODV, out to either the 700 primer nozzles and/or the main fuel nozzles.
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An automatic engine fuel prime feature is activated during engine start and stops when the engine
starter drops out. During single engine starts with the engine fuel system selector in cross feed
the automatic prime feature is unable to prime the engine being started.
Main fuel tanks useable fuel is approximately 360 gallons/2412 lbs. (JP-8). There are three
methods to refuel the aircraft: gravity, pressure, and closed circuit.
• Gravity – Main fuel tanks useable fuel is approximately 360 gallons (180 per/tank).
• Pressure – Main fuel tanks useable capacity is approximately 359 gallons (179.5 per/tank).
Maximum fuel pressure is 55 psi, at 300 GPM.
• Closed circuit – Main fuel tanks useable capacity is approximately 356 gallons (178 per/tank).
Maximum fuel pressure is 15 psi, at 110 GPM).
Two low-level sensors provide signals which activate the #1 FUEL LOW or #2 FUEL LOW caution(s).
These cautions flash when the fuel level decreases to approximately 172 lbs in each tank.
Even though cold weather does not particularly affect the engine itself, it still causes the usual
problems of ice in the fuel lines, control valves, and fuel sumps.
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• When starting in cold weather below -40 C, if engine light-off does not occur within 45
seconds after initial indication of Ng SPEED, abort the start and refer to Chapter 8 of the -10.
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MECHANICAL MIXING UNIT (MMU)
A mechanical mixing unit provides control mixing functions which minimizes inherent control coupling.
The four types of mechanical mixing and their functions are as follows:
• Collective to Pitch – Compensates for the effects of changes in rotor downwash on the stabilator
caused by collective pitch changes. The mixing unit provides forward input to the main rotor as
collective is increased and aft input as the collective is decreased.
• Collective to Roll – Compensates for the rolling moments and translating tendency caused by tail
rotor thrust. The mixing unit provides left lateral input to the main rotor system as collective is
increased and right lateral input as collective is decreased.
• Collective to Yaw – Compensates for changes in torque effect caused by changes in collective
position. The mixing unit increases tail rotor pitch as collective is increased and decreases tail rotor
pitch as collective is decreased.
• Yaw to Pitch – Compensates for changes in the vertical thrust component of the canted tail rotor as
tail rotor pitch is changed. The mixing unit provides aft input to the main rotor system as the tail
rotor pitch in increased and forward input as tail rotor pitch is decreased.
COLLECTIVE/AIRSPEED TO YAW (ELECTRONIC COUPLING)
This mixing is in addition to collective to yaw mixing. It helps compensate for the torque effect caused by
changes in collective position and is a function of the SAS/FPS computer.
It has the ability to decrease tail rotor pitch as airspeed increases and the tail rotor and cambered fin
become more efficient. As airspeed decreases the opposite occurs.
The SAS/FPS computer commands the yaw trim actuator to change tail rotor pitch as collective
position changes. The amount of tail rotor pitch change is proportional to airspeed.
Maximum mixing occurs from 0 to 40 knots. As airspeed increases above 40 knots, the amount
of mixing decreases until 100 knots, after which no mixing occurs.
TORQUE COMPENSATION
The UH-60 attempts to compensate for the torque produced by main rotor with three distinct design
features working in concert to reduce pilot workload. The first two features, the MMU and electronic
coupling, are discussed above and the last feature is the cambered fin. In addition, changes to the
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nominal pitch setting in the tail rotor from 7.5 to 10 have had an effect on the efficiency of the
compensation. The effects of Wind, PA and Temp are not considered.
The cambered fin assists in streamlining the fuselage as airspeed is increased. The fin does little
to assist at lower airspeeds.
• Increased efficiency of the cambered fin tapers off at greater speeds not quite compensating for all
the torque required to cruise.
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• The cambered fin is notched to unblock the tail rotor allowing for more tail rotor efficiency at
a hover. This increases the sideward flight airspeed limit but reduces the streamlining effect.
The MMU is a fixed ratio and would quickly overcompensate as the aircraft accelerates with a
constant torque setting. Therefore, it is designed to add the torque compensation needed at higher
airspeeds when the increased efficiency of the cambered fin has tapered off.
Collective/Airspeed to Yaw Electronic Coupling provides the torque compensation necessary to cover
the gap at lower airspeeds where the cambered fin is less than optimally efficient.
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AUTOMATIC FLIGHT CONTROL SYSTEM (AFCS)
The AFCS enhances the stability and handling qualities of the helicopter. It is comprised of four
basic subsystems:
• SAS 1 and/or 2
• Trim
• FPS
• Stabilator
AFCS provides oscillation damping (dynamic stability) and maintains desired attitude, speed,
and handling (static stability).
Static stability (long term stability) is the tendency to return to the pilot’s desired attitude, heading,
or airspeed.
Dynamic stability (short term stability) is the tendency to resist movement. It prevents
porpoising, rocking, and fishtailing.
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STABILITY AUGMENTATION SYSTEM (SAS)
SAS enhances dynamic stability by providing short term rate damping in the pitch, roll, and yaw axis. The
UH60 A/L incorporates two SAS systems: SAS 1 (analog) and SAS 2 (digital). Their operation is
essentially the same, except that SAS 2 has self-diagnostic capabilities.
SAS responds to short term aerodynamic disturbances and effectively dampens the helicopter
movement. Since the responses are almost instantaneous (2 times per/sec), SAS has a limited control
authority (10%) to prevent SAS malfunctions from causing undesirable helicopter response before the
pilot can react.
SAS 1 and SAS 2 utilize the same SAS actuators and their hydraulic pressure is monitored. In case of
loss of actuator pressure or if both SAS 1 and SAS 2 are turned off, the SAS OFF caution will appear.
SAS 1
• Control authority of 5% and is controlled by the SAS 1 amplifier. The SAS 1 amplifier uses the
vertical gyro, pitch rate gyro or inputs from AHRS to derive roll attitude and rate for the roll SAS
commands.
• Malfunction of the SAS 1 system may be detected as an erratic motion in the helicopter without a
corresponding failure advisory indication. Erratic electrical input to a SAS actuator can result in
moderate rotor tip path oscillations that are often accompanied by a knocking sound which may be
felt in the cyclic or pedals. If this is experienced, SAS 1 should be turned off.
SAS 2
• Control authority of 5% control authority and is controlled by the SAS/FPS computer.
• SAS/FPS computer provides fault monitoring.
• In case of a malfunction of SAS 2, the input will normally be removed from the actuator and the
SAS 2 fail advisory light on the AUTO FLIGHT CONTROL panel will illuminate. If it is intermittent
the indication can be cleared with a POWER ON RESET. If it is continuous, the SAS 2 should be
turned off.
With SAS 1 or SAS 2 off, the remaining SAS will double its gain and the control authority of the SAS is
reduced by one-half (5% control authority).
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TRIM SYSTEM
When the TRIM is engaged on the AUTO FLIGHT CONTROL panel, the pitch, roll, and yaw trim
systems are activated to maintain position of the cyclic and tail rotor controls. Proper operation of the
yaw trim requires that the BOOST be on.
The tail rotor and lateral cyclic forces are developed in the electromechanical yaw and roll trim
actuators. Both actuators incorporate a slip clutches to allow pilot control inputs if either
actuator becomes jammed.
• The force required to break through the clutch are 80 lbs maximum in yaw and 13 lbs maximum in
roll.
The longitudinal (pitch) force is developed by an electrohydromechanical actuator operated in
conjunction with the SAS/FPS computer.
The trim system provides a force gradient in the pitch, roll, and yaw axis.
• The pilot may remove the longitudinal and lateral cyclic gradient by pressing the thumb operated
TRIM REL switch on the cyclic.
• The pedal gradient maintains pedal position whenever the trim is engaged. By placing feet on the
pedals, the pedal switches are depressed and the gradient force is removed.
Operation of the trim system is continuously monitored by the SAS/FPS computer. If a malfunction
occurs, the SAS/FPS computer will shut off the trim actuator(s) driving the affected axis, and the TRIM
FAIL and FLT PATH STAB cautions will appear. If it is intermittent the indication can be cleared with a
POWER ON RESET.
In addition to the trim release switch, a four-way trim switch (STICK TRIM) on the cyclic establishes a
trim position without releasing trim.
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FLIGHT PATH STABILIZATION (FPS)
FPS enhances static stability by providing long term rate damping in the pitch, roll, and yaw axis. FPS
provides basic autopilot functions using the trim actuators to maintain attitude in the pitch and roll axis
and heading hold/turn coordination in the yaw axis. When FPS is coupled with trim it has 100% control
authority.
Proper operation requires that the BOOST, TRIM, SAS 1 and/or SAS 2 functions have been selected on
the AUTO FLIGHT CONTROL panel. Although not required for operation, the FPS performance will be
improved with proper operation of the stabilator in the auto mode.
FPS will not function is TRIM is not engaged.
The trim attitude, once established, will be automatically held until changed by the pilot.
• When pitch attitude is changed by means of STICK TRIM switch, there is a delay from the time the
input is applied until the new reference speed is acquired. This is to allow the aircraft to accelerate
and decelerate.
• At airspeeds greater than 60 knots, heading hold is automatically disengaged and turn coordination
engaged under these conditions:
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STICK TRIM switch is actuated in the lateral direction.
TRIM REL switch is pressed and roll attitude is greater than prescribed limits.
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About ½ inch cyclic displacement and a roll attitude of about 1.5 . Heading hold is automatically
reengaged and turn coordination disengaged upon recovery.
• To make a coordinated turn, the pilot enters a turn in one of the following ways:
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Changing reference roll attitude by pressing the STICK TRIM switch in the desired lateral direction.
Pressing TRIM REL switch on the cyclic and establishing the desired bank angle with feet off pedal
switched.
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Ø
Ø
Exerting a lateral force on the cyclic to achieve the desired bank angle, and then neutralizing the force with the
STICK TRIM switch.
Keeping a lateral force on the cyclic for the duration of the turn.
FPS monitoring is automatic. If a malfunction is detected, the FLT PATH STAB caution will appear and
the FPS will either continue to operate is a degraded mode or may cease to function altogether.
The pilot must take over manual flight, and may either turn the FPS off or evaluate performance
to determine the degree and type of degradation, and continue flight with the remaining features.
To help evaluate the nature of the degradation, eight failure advisory indicators are displayed on two
FAILURE ADVISORY switches on the flight control panel.
• If a light goes on, it may be turned off by pressing the lighted switch.
• The pilot may attempt to clear the indication of temporary malfunction by conducting a POWER ON
RESET. If the FLT PATH STAB caution disappears, it may be assumed that normal operation is
restored.
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STABILATOR SYSTEM
The helicopter has a variable angle of incidence stabilator to enhance the 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.
The stabilator is programmed to:
• Align stabilator and main rotor downwash in low speed flight to minimize nose-up attitude resulting
from downwash.
• Provide collective coupling to minimize pitch attitude excursions due to collective inputs. The
coupling of stabilator position to collective displacement is automatically phased in beginning at 30
KIAS. An increase of collective results in the trailing edge programming down. A decrease in
collective produces the opposite stabilator reaction.
• Decrease angle of incidence with increased airspeed to improve static stability.
• Provide lateral sideslip to pitch coupling to reduce susceptibility to wind gusts. When the helicopter
is out of trim in a slip or a skid, pitch excursion are induced as a result of main rotor downwash on
the stabilator. Nose left (right slip) results in the trailing edge programming down. Nose left
produces the opposite stabilator reaction.
• Provide pitch rate feedback to improve dynamic stability. The rate of pitch attitude change of the
helicopter is sensed and used to position the stabilator to help dampen pitch excursions during
gusty wind conditions. A sudden pitch up would cause the stabilator to be programmed trailing
edge down to induce a nose-down pitch to dampen the initial reaction.
When initial power is supplied the stabilator system, it will be in the automatic mode.
It is possible to manually slew one actuator using battery power only. If the stabilator is slewed
up, regain automatic control by manually slewing stabilator full down, then push AUTO
CONTROL RESET twice.
If the automatic mode shuts down during flight because of a failure, the helicopter shall be slowed
to 80 KIAS before power is restored.
The TEST switch is used to check the AUTO mode fault detector feature and is inoperable above
60 KIAS.
WARNING
If the airspeed fault advisory light is illuminated, continued flight
above above 70 KIAS with the stabilator in the AUTO MODE is unsafe
since a loss of the airspeed signal from the remaining airspeed sensor
would result in the stabilator slewing full-down.
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HYDRAULIC SYSTEM
The three hydraulic systems are designed to provide full flight control pressure. There are three hydraulic
pressure systems: No. 1, No. 2, and backup. All are completely independent and each is fully capable of
providing essential flight control pressure for maximum system redundancy.
• Complete redundancy is accomplished by the backup pump providing hydraulic power to both No.
1 and/or No. 2 systems if one or both fail.
• If the No. 1 and No. 2 systems lose pressure, there will be a slight restriction in the maximum rate
of flight control movement due to only the backup pump supplying both systems with hydraulic
power.
Hydraulic pump modules
The No. 1, No. 2, and backup pump modules are identical and interchangeable with each other and
a constant pressure, variable volume pump. They maintain a required 3,000 psi for their own system.
• The No. 1 and No. 2 are mounted on driven by their respective accessory module of the main
transmission. The backup pump is mounted on and driven by an ac electric motor.
• Each pump has two filters: a pressure and a return filter. A red indicator button will pop out when
pressure goes up 70 ±10 psi above normal. The pressure filter has no bypass. The return filter has
a bypass valve that opens at 100 ±10 psi above normal. This prevents a hydrostatic lock of the
flight controls.
• A fluid quantity switch senses fluid loss for that system. When the piston is the pump module move
down to the REFILL mark (60% fluid remaining), the piston closes the switch and activates the
RSVR LOW caution.
• Each pump has two temperature sensitive labels. When the temperature label indicates
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132 C/ 135 C has been exceeded, an entry shall be made on the DA Form 2408-13-1.
Number 1 hydraulic system
The No. 1 hydraulic system supplies the first stage of all primary servos and the first stage of the
tail rotor servo.
• First stage tail rotor servo can be manually turned off by the TAIL SERVO switch on the
miscellaneous switch panel or by hydraulic logic’s LDI feature.
• The tail rotor servo is a two-stage servo but, unlike the primary servos, only one stage
is pressurized at a time.
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Number 2 hydraulic system
The No. 2 hydraulic system supplies the second stage primary servos and the pilot assist servos.
• The pilot assist servos cannot be turned off collectively, by the pilot; however SAS, TRIM, and
BOOST can be manually turned off by the switches on the AUTO FLIGHT CONTROL panel.
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Backup hydraulic system
The backup hydraulic system supplies emergency pressure to the No. 1 and or No. 2 hydraulic
systems whenever a pressure loss occurs. It supplies pressure to the No. 2 stage of the tail rotor
servo. The system provides a hydraulic pressure for automatic recharging of the APU accumulator.
• The internal depressurization valve reduces the output pressure of the backup pump upon startup
of the electric motor. This valve unloads the electric motor by reducing torque requirements at low
rpm. After about 0.5 sec (main generator) or 4 sec (APU generator or external power), the valve
is closed and 3,000 psi pressure is supplied.
• A weight on wheels (WOW) switch on the left main landing gear provides automatic operation of
the backup pump when the helicopter is in the air, regardless of the BACKUP HYD PUMP switch
position, and disables the backup pump ac thermal switch.
• There are 5 conditions that automatically turn on the backup pump when ac power is available:
Hydraulic leak detection/isolation (LDI) system
LDI protects the flight control hydraulic system by preventing the further loss of hydraulic fluid in
case of a leak. It uses pressure switches and fluid level sensors for monitoring hydraulic pump fluid
level, and pressure for primary servos, tail rotor servos, and pilot-assist servos.
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When a pump module reservoir fluid level switch detects a fluid loss, the logic module follows the LDI
sequence in the following pages, to isolate the leak.
• The logic module operates the required shutoff valve(s) to isolate the leak and hydraulic logic turns
on the backup pump, when required. The RSVR LOW caution for that system will appear.
• Backup pump and shutoff valve(s) operation is automatic through the logic module.
• If after the isolation sequence, the leak continues, the leakage is in the stage 1 or 2 primary servos
and the appropriate SVO OFF switch must be moved to the off position by the pilot.
24
LEAKAGE IN #1 HYDRAULIC SYSTEM – 3 STEP PROCESS
LEAK DETECTED
#1 RSVR
LOW
STEP #1
HYD LOGIC/LDI
ISOLATES #1 TAIL
RTR SERVO
PRESSURIZES #2
TAIL RTR SERVO
NO OTHER LIGHTS
LEAK IN #1 TAIL
RTR SERVO
IF THE LEAK
CONTINUES
#1 TAIL RTR
SERVO
BACK-UP
PUMP ON
#2 TAIL RTR
SERVO ON
#1 HYD
PUMP
STEP #2
TRANSFER VALVE
ISOLATES
#1 HYD PUMP
HYD LOGIC
RESTORES #1
TAIL RTR SERVO
#1 HYD
PUMP
BACK-UP PUMP
PRESSURIZES
#1 HYD SYSTEM
NO OTHER LIGHTS
LEAK UPSTREAM
OF #1 TRANSFER
MODULE
IF THE LEAK
CONTINUES
BACK-UP
PUMP ON
BACK-UP
RSVR LOW
STEP #3
PILOT (SVO OFF)
ISOLATED
#1 PRIM SERVO
DID NOT ISOLATE
#1 PRIM SERVO
#1 PRI
SERVO PRESS
BACK-UP
RSVR LOW
#1 PRI
SERVO PRESS
BACK-UP
RSVR LOW
#1 RSVR
LOW
#1 HYD
PUMP
#1 RSVR
LOW
#1 HYD
PUMP
BACK-UP
PUMP ON
#1 TAIL RTR
SERVO
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LEAKAGE IN #2 HYDRAULIC SYSTEM – 3 STEP PROCESS
LEAK DETECTED
#2 RSVR
LOW
STEP #1
HYD LOGIC/LDI
ISOLATES PILOT
ASSIST AREA
NO OTHER LIGHTS
LEAK IN PILOT
ASSIST AREA
IF THE LEAK
CONTINUES
BOOST SERVO
OFF
SAS OFF
FLT PATH
STAB
TRIM FAIL
#2 HYD
PUMP
STEP #2
TRANSFER VALVE
ISOLATES
#2 HYD PUMP
HYD LOGIC
RESTORES PILOT
ASSIST AREA
#2 HYD
PUMP
BACK-UP PUMP
PRESSURIZES
#2 HYD SYSTEM
NO OTHER LIGHTS
LEAK UPSTREAM
OF #2 TRANSFER
MODULE
IF THE LEAK
CONTINUES
BACK-UP
PUMP ON
BACK-UP
RSVR LOW
STEP #3
PILOT (SVO OFF)
ISOLATED
#2 PRIM SERVO
DID NOT ISOLATE
#2 PRIM SERVO
#2 PRI
SERVO PRESS
BACK-UP
RSVR LOW
#2 PRI
SERVO PRESS
#2 RSVR
LOW
#2 HYD
PUMP
#2 RSVR
LOW
BACK-UP
PUMP ON
BOOST SERVO
OFF
FLT PATH
STAB
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BACK-UP
RSVR LOW
#2 HYD
PUMP
SAS OFF
TRIM FAIL
PNEUMATIC SYSTEM
The pneumatic system operates from bleed air furnished by the main engines, the APU, or an external
pneumatic power source, is used to provide the following:
• Drive the #1 and #2 engine starter.
• Heating system operation.
• External extended range tank fuel transfer.
Bleed air from the main engines is used for engine inlet anti-icing subsystem operation.
With engine bleed air turned on, the maximum available torque is reduced as follows:
• Engine anti-ice on: Reduce torque available by a constant 16%.
• Cockpit/gunner heater on: Reduce torque available by 4%.
The heating subsystem and the extended range fuel tanks use bleed air supplied by the main engines
during flight and on the ground, APU, or external source.
27
POWERTRAIN SYSTEM
The powertrain consists of inputs from two engines, a main transmission, intermediate gear box, tail gear
box, and connecting drive shafting. The main transmission consists of five modules: two input modules,
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the main module, and two accessory modules. The main transmission has a built in 3 forward tilt.
Input module
The input modules are mounted on the left and right front of the main module and support the front
of the engines.
The freewheeling unit allows engine disengagement during autorotation, or in case of a
nonoperating engine, the accessory module will continue to be driven by the main rotor.
The input module provides the first gear reduction between the engine and main module.
Accessory module
One accessory module is mounted on the forward section of each input module and provides
mounting and drive for an electrical generator and a hydraulic pump.
A rotor speed sensor is mounted on the right accessory module and provides the Nr signal.
On the UH-60L/UH-60A+, an additional rotor speed signal is mounted on the left accessory module
which provides input signals for improved transient droop response.
Main module
The main module contains the necessary gearing to drive the main rotor and tail rotor systems. It
provides a reduction in speed from the input module to the main module and the tail drive shaft.
The tail drive system consists of six sections of drive shaft connect the main module to the tail rotor gear
box. The shafts drive the oil cooler blower and transmit torque to the tail rotor. The shafts are ballistically
tolerant if hit by a projectile and are suspended at four points in viscous damped bearings.
28
Intermediate gear box (IGB)
The IGB is mounted at the base of the tail pylon. It transmits torque and reduces shaft speed from
the main gear box to the tail gear box.
The IGB may run at cruise flight for 30 minutes, with loss of all oil.
Tail gear box
The tail gear box is mounted at the top of the tail pylon and transmits torque to the tail rotor head.
The gear box mounts the tail rotor, changes angle of drive, and gives a gear reduction. It also
enables pitch changes to the tail rotor blades through the flight control system.
The tail gear box may run at cruise flight for 30 minutes, with loss of all oil.
29
MAIN AND TAIL ROTOR GROUPS
The rotor system consists of a main rotor and tail rotor. Both systems are driven by the engines through
the transmission system, with pitch controlled by the flight controls system.
Main rotor system (Fully-Articulated)
The main rotor system consists of four subsystems: main rotor blades, hub, flight controls, and the
bifilar vibration absorber.
Four titanium-spar main rotor blades attach to spindles which are retained by elastomeric
bearings contained within the hub.
The elastomeric bearing permits the blade to flap, lead, and lag. Lag motion is controlled by
hydraulic dampers and blade pitch is controlled through adjustable control rods which are moved by
the swashplate.
• Each damper is supplied with pressurized hydraulic fluid.
• Each reservoir has an indicator that monitors the reserve fluid. The reservoir is fully serviced when
the indicator shows a full gold band.
When rotor is not turning, the blades and spindles rest on hub mounted droop stops.
Upper restraints called antiflapping stops retain flapping motion caused by the wind.
Main rotor blades
Four main rotor blades use a titanium spar for their structural member. The leading edge of each
blade has an abrasion strip, electro-thermal blankets are bonded into the blades leading edge for
deicing, and wire mesh for lightning strike.
A Blade Inspection Method (BIM) indicator is installed on each blade at the root end trailing edge to
visually inspect the blade spar structural integrity. If is spar crack occurs or a seal leaks, nitrogen
will escape. When the pressure drops below minimum, the indicator will show red bands.
Main rotor gust lock
The gust lock prevents the blades from rotating when the helicopter is parked. It is designed to
withstand torque from one engine, at IDLE, and allows engine maintenance checks without drive
train rotation.
Tail rotor system (Rigid)
A cross-beam tail rotor blade system provides anti-torque action and directional control. Blade flap
and pitch change motion is provided by deflection of the flexible graphite fiber spar. The spar is a
continuous member running from the tip of one blade to the tip of the opposite blade.
Electro-thermal blankets are bonded into the blade leading edge for deicing.
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The tail rotor head and blades are installed on the right side of the tail pylon, canted 20 upward.
In addition, the tail rotor provides 2.5% of the total lifting force in a hover.
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A spring loaded feature of the tail rotor servo will provide a setting of the tail rotor blades for
balance flight at cruise power setting in case of complete loss of tail rotor control.
Tail rotor quadrant/warning
The tail rotor quadrant contains microswitches to activate the TAIL ROTOR QUADRANT caution if a
tail rotor cable becomes severed.
Spring tension allows the quadrant to operate in a normal manner.
If the helicopter is shut down and/or hydraulic power is removed with one tail rotor cable feature,
disconnection of the other tail rotor cable will occur when force of the boost servo cannot react against
the control cable quadrant spring tension.
31
UTILITY SYSTEMS
Windshield anti-ice/defogging system
Pilot’s, copilot’s and center windshield (if equipped) are electrically anti-iced and defogged.
Transparent conductors imbedded between the laminations provide heat when electrical power is
applied
The windshield anti-ice system fault monitoring circuit should prevent windshield burnout when the
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windshield surface heat is above 43 C. If heat increases, the monitor circuit will turn off the system.
• Continued use of a faulty windshield anti-ice system may result in structural damage (delamination,
cracking, and/or fire) to the windshield.
• Do not allow ice to accumulate on the windshield, as ice shedding can cause engine FOD.
If the APU generator is the sole source of ac power, the backup pump and the windshield antiice cannot be used simultaneously.
Wire strike protection system
The system is a simple, lightweight, passive system with no motorized or pyrotechnic components
used to cut, break, or deflect wires that may strike the helicopter in the frontal area.
The system consists of nine cutters/deflectors located on the fuselage and landing
gear/support fairing.
Rotor blade deice kit
The blade deice system provides improved mission performance in icing conditions by applying
controlled electrical power to heating elements in the main and tail rotor blades, causing the ice bond
layer to melt allowing symmetrical ice shedding.
Droop stop heaters apply heat to the droop stops hinge pins, to prevent icing and permit
proper operation.
The blade deice system, excluding element-on-time (EOT) failure, may be ground checked using the
APU generator.
The OAT sensor, installed below the windshield, provides a signal to the controller for heating EOT
of the rotor blades. The lower the OAT, the longer the EOT will be.
When one main generator is inoperative, deice power can be supplied by the APU generator. The
APU GEN ON advisory light will not appear.
Flight data recorder
If installed, it is located in the aft transition avionics compartment and is a crash survivable digital
tape recorder providing 25 hours of recorded data on a continuous loop magnetic tape. The recorder
begins to record data as soon as ac and dc essential power is supplied. There are no controls
provided to the pilot or copilot for control of the recorder.
32
ELECTRICAL POWER SUPPLY SYSTEM
Alternating current (ac) is the primary source of power. The primary electrical system consists of two
identical/independent systems, each capable of supplying the total helicopter power requirements
The prime source for each system is a 115/200 vac generator. A subsystem feeds two independent
as primary buses and an ac essential bus.
• A portion of each ac primary bus load is converted to 28 volts direct current (vdc) by two ac/dc
converters.
• The 28 vdc is distributed by two independent dc primary buses and a dc essential bus.
Emergency power is provided by the APU generator. The APU generator is capable of supplying most
flight-essential ac and dc bus loads, excluding rotor blade de-ice system.
• In addition, the APU generator can supply power to the blade deice system if one main generator
should fail.
• Should a second main generator fail, the blade deice load will be dropped and the APU generator
will power the remaining ac bus loads.
DC power supply system
Primary dc power is obtained from two converters with a battery as the secondary power source.
Converters
• Two 200 ampere converters, each powered by the No. 1 and No. 2 ac primary buses, turn ac
power into dc power and reduce it to 28 volts.
• If one converter’s output is lost, the converter load will be transferred to the operating system
and the #1 CONV or #2 CONV caution will appear.
Sealed Lead Acid Battery (SLAB)
• Supplies dc power to the battery bus, battery utility bus, and dc essential bus for operating dc
essential equipment during primary dc malfunction.
• When only battery power is available, the battery life is about 38 minutes day and 24 minutes night
for a battery 80% charged.
• The BATT switch should be ON when external, APU generator, or main generator power is
applied. This will recharge the battery.
• The BATT LOW CHARGE caution appears when voltage on the battery utility bus drops below 23
vdc. The BATT switch should be turned OFF immediately, so that battery power can be conserved
for an APU start.
AC power supply system
Delivers regulated three phase, 115/200 vac, 400 Hz. Each system contains a generator mounted on
the accessory module, a current transformer, a generator control unit, and current limiters.
33
A generator main bearing caution system is installed on each main generator to activate the #1 GEN
BRG or #2 GEN BRG to indicate a worn or failed bearing.
• The auxiliary bearing will allow 10 additional hours of operation after the caution appears.
• When the GEN BRG caution appears for more than 1 minute, make an entry on the DA
Form 2408-13-1.
The generator control units monitor voltage from the No. 1, No. 2, and APU generators and take
the generator(s) off-line when malfunctions occur.
Auxiliary AC power system
The auxiliary ac power system is a backup ac power source that provides electrical power for
ground checkouts. The system consists of the generator driven by the APU, a current transformer,
and a generator control unit.
• An APU GEN ON advisory will appear when the APU generator is operating and the APU
generator switch is ON. The APU GEN ON advisory will appear only when supplying power to the
system; it will not appear when either the No. 1 or No. 2 generator is supplying power.
• External power will be introduced into the system if acceptable external power is connected, the
EXT PWR switch is ON, and no other generating source is operating. An EXT PWR CONNECTED
advisory will appear whenever external power is connected.
34
ELECTRONIC NAVIGATION INSTRUMENT DISPLAY SYSTEM
The system provides displays for navigation and command signals on the VSI and HSI for pilot
visual reference. It consists of two VSIs and two HSIs on the instrument panel. The system has a
common command instrument system processor (CISP), two HSI/VSI mode select panels and one
CIS mode select panel.
Steering command bars and pointer
The roll and pitch command bars and the collective position pointer operate in conjunction with the
CISP and the CIS MODE SEL.
• Selection of HDG on the CIS MODE SEL panel provides a display of a roll signal by the roll
command bar.
• Selecting the CIS MODE SEL switch NAV and the MODE SEL switch VOR ILS, the roll command
bar will display roll commands from the CISP.
• If an ILS (LOC) frequency is tuned in, the pitch command bar and the collective command pointer
will also display CISP signals.
• If a VOR frequency is tuned-in, the pitch command bar and the collective command pointer will
be held from view.
Heading mode
The heading mode processes the heading error and roll attitude signals to supply a limited cyclic
roll command which, when followed, causes the helicopter to acquire and track the heading
selected of the HSI.
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• The processor gain provides 1 of roll command for each degree of heading error up to a
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roll command limit of approximately 20 .
• When properly followed, the command results in not more than one overshoot and a tracking error
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of not more than 2 .
Altitude hold mode
The altitude hold mode processes barometric pressure signals from the air data transducer in addition
to the collective stick position signal.
• The CISP provides collective command signals which, when properly followed, cause the helicopter
to maintain altitude to within ±50 feet.
• The mode synchronizes on the engagement altitude for vertical rates up to 200 fpm and provides
for altitudes between -1000 and 10,000 feet at airspeeds from 70 to 150 KIAS.
VOR NAV mode
The CISP processes heading and course signals to provide a limited cyclic roll command which, when
followed, shall cause the helicopter to acquire and track the course setting selected on the HSI.
35
O
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Engagement of the VOR NAV when the helicopter position is in excess of 10 to 20 from the
selected radial will cause the initial course intersection to be made in the heading mode.
• The CISP will light the CIS mode selector HDG switch ON legend during the initial course
intersection.
O
O
• When the helicopter is within 10 to 20 of the selected course, the CISP will capture the VOR
lateral beam. CISP will turn off the HDG switch ON legend and the final course interception, about
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45 . This causes the roll command pointer to deflect in the direction of the required control
response.
• When properly followed, the command will result in not more than one over shoot at a range of 10
NM at a cruise speed of 100 ±10 knots and not more than 2 overshoots at ranges between 5 and
40 NM at speeds from 70 – 140 knots.
When passing over the VOR station, the CISP reverts to a station passage submode for 30 seconds.
Cyclic roll commands will be obtained from the HSI course datum signal.
• Course changes to a new radial or identification of VOR intersections may be made before
station passage by setting the HSI HDG control to the present heading and actuating HDG.
• This will disengage the NAV mode and allow the pilot to continue on the original radial in the
heading mode. A VOR intersection fix of selection of a new radial course may be made without
affecting the CIS steering commands.
• Actuating the NAV switch reengages the VOR NAV mode to either continue on the original radial or
to initiate and intercept to a newly selected radial.
ILS NAV mode
The mode is established by selecting the VOR/ILS switch on the VSI/HSI mode selector, tuning a
localizer frequency, and selecting the NAV on the CIS MODE SEL panel.
• The indicated airspeed and pitch attitude signals are processed to provide a limited cyclic pitch
command which, when properly followed, will result in maintain an airspeed that should not
deviate more the 5 knots from the IAS at the time the mode was engaged.
• The barometric altitude and collective stick position signals are processed to provide a limited
collective position indication which, when properly followed, will cause the helicopter to maintain the
altitude existing at the time the mode was engaged. The CISP will cause the ALT hold switch ON
legend to light.
• Desired approach runway course must be set on the course window of the HSI selected by the
PLT/CPLT indication of the CRS HDG switch.
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• The course interception, about 45 , acquisition, and tracking will be done as described in VOR
NAV except that not more than one overshoot at a range of 10 NM at 100 ±10 knots and not more
than 2 overshoots at ranges between 5 and 20 NM at speeds from 70 – 130 knots.
Approach mode
The approach mode, a submode of the ILS NAV mode, will automatically engaged when the
helicopter captures the glide slope.
36
When the glide slope is intercepted:
• The CISP disengages the altitude hold mode and causes the ON legend of the ALT hold switch to
go off.
• The CISP provides a down movement of the collective position indicator to advise pilot of the
transition from altitude hold to glide slope path acquisition and tracking. When properly followed,
will result in not more than one overshoot in acquiring the glide path and have a glide path tracking
free of oscillations.
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• The cyclic roll commands are limited to ±15 during the approach submode. When properly
followed, the roll commands will result in the helicopter tracking the localizer to an approach.
• The cyclic roll and collective steering performance is applicable for approach airspeed from 130
KIAS down to 50 KIAS.
BACK CRS mode
The back course mode is a submode of the ILS NAV mode and is engaged by concurrent ILS ON and
BACK CRS ON signal from the HSI/VSI mode selector.
• The CISP provides cyclic roll commands which, when properly followed, will allow the pilots to
complete back course localizer approach in the same manner as the front course.
• The desired final approach course should be set on the selected HSI course window.
Level-Off mode
The level-off mode will be activated when either the VOR NAV or ILS NAV modes are engaged and
will be deactivated by selection of another mode or when the radar altitude valid signal is not present.
• During ILS and VOR approaches, the barometric altimeter must be used to determine arrival at the
minimum altitude.
• This mode is automatically engaged when the radar altitude goes below either the pilot’s or
copilot’s radar altimeter low altitude warning bug setting, whichever is set higher.
• The CISP provides a collective pointer command which, when properly followed, will cause the
helicopter to maintain the altitude with 10 feet of the low bug setting for setting below 250 feet and
20 feet for settings above 250 feet.
• The CISP causes the ALT switch ON legend to light and engage the altitude hold mode.
Go-Around mode
The go-around mode processes several inputs to provide cyclic roll, cyclic pitch, and collective position
indication. The go-around mode is engaged when either pilot presses the GA switch on the cyclic control
grip.
• Once engaged, the CISP provides collective position indications which, when followed, will result in
a 500 ±50 fpm rate of climb at zero back angle.
37
• Five seconds after the GA switch is pressed, the CISP will provide cyclic pitch bar commands
which, when properly followed, will result in an 80 KIAS for the climbout.
• The go-around mode is disengaged by changing to any other mode on the pilot’s CIS mode
selector.
Doppler, Doppler/GPS mode
The mode is engaged by selecting the DPLR, DPLR/GPS switch on the VSI/HSI mode selector and
the NAV switch on the CIS mode selector. The CISP provides cyclic roll bar commands which, when
followed, result in a straight line, wind corrected flight over distances greater than 0.2 kilometer from
the destination.
• The initial course is the course the Doppler/GPS computes from the helicopter’s position to
the destination at the time the fly to destination was entered.
• The VSI and HSI course sensitivity is ±1000 meters when farther than 12 km from the fly-to
destination. Course sensitivity gradually scales down from ±1000 meters at 12 km to ±200 meters
at 2 km and less from the fly-to destination.
• To achieve a pictorially correct view of the course, rotate the course knob to the head of the No. 1
needle when the fly-to destination is entered.
• The DPLR, DPLR/GPS NAV logic detects the condition of station passover and automatically
switches to the heading mode. This mode does not automatically reengage, but will require manual
reengagement of the NAV switch.
FM HOME mode
The mode is engaged by selecting the FM HOME switch on the pilot’s VSI/HSI mode selector and the
NAV switch on the pilot’s CIS mode selector. The FM homing mode directs FM homing signals only
to the VSI. Other NAV modes will be retained on the HSI, if previously selected.
• The CISP provides cyclic roll commands to aid the pilot in homing on a radio station selected on
the No. 1 VHF-FM radio when properly followed, results in no more than two overshoot
O
heading changes before maintaining a tracking error not to go over 3 .
O
• The CISP will revert to heading mode whenever lateral deviation rate is over 1.5 /sec for a
period of over 1 second.
• Concurrent VOR and FM or concurrent DPLR and FM mode inputs will be considered an FM mode
input to the CISP.
38
INTEGRATED VEHICLE HEALTH MANAGEMENT SYSTEM (IVHMS)
IVHMS will perform the following functions: built-in test, mechanical diagnostics, structural usage
monitoring, exceedance monitoring, engine monitoring, engine output shaft balancing, absorber tuning,
and rotor track and balance.
Built-in test: IVHMS performs BIT checks of all system LRUs and accelerometers. The built-in test
screens display the status of the STARTUP (SBIT), PERIODIC (PBIT), and INITIATED (IBIT) BIT’s
and allow the operator to initiate and reset the IBIT.
Mechanical diagnostics: This function identifies degraded components and types of faults for all
monitored components. It is designed to perform in an automatic mode to collect data without aircrew
input during all flights and it is also designed to include a manual mode that will collect data on aircrew
demand.
Exceedance monitoring: The goal of the exceedance monitoring function is to observe and record all
aircraft exceedances and their duration for more accurate maintenance troubleshooting and repair.
Additionally, a specified time prior to and after the exceedance will be recorded. The total time for
this recording will not exceed 75 seconds.
NOTE
Absence or failure of one or more system function, or segment, will not affect the
performance of any other function (e.g. inoperative mechanical diagnostics functions will
not affect the rotor track and balance function).
The cockpit voice and flight data recorder simultaneously records up to four channels of audio. Each
channel retains the most recent 120 minutes of all recording audio, digital, and timing data, and a
minimum of 13 hours of flight data information.
Cockpit Display Unit (CDU)
The CDU provides the interface between the IVHMS and the aircrew.
• EVT CAP (Event Capture): Allows the operator to mark an event. This option commands IVHMS to
immediately record a predefined set of data (15 seconds before and after the button was pressed.)
When activated, the EVENT CAPTURED COMMANDED screen will be displayed for approximately
10 seconds. A data captured message will appear when the recording is complete.
39
• ADMIN (Administration Menu): The operator is offered the following administrative functions: Data
Transfer Unit (DTU) card change, crew chg/end of op, aircraft configuration change*, date change*,
time change*, part numbers, gross weight, change password*, and reprogram*. An * indicates a
password protected function.
When the CREW CHG/END OF OP procedure is selected, the operator will be prompted to
“PRESS ENTER TO INITIATE A CREW CHANGE OR END OF OPERATION.” After pressing the
ENTER button a prompt “WAIT WHILE THE IVHMU SAVES LATEST DATA” will be displayed on
the center of the screen and “DTU Status: DOWNLOADING DATA” will be displayed at the bottom
of the screen. The DTU Status will display download complete when finished.
• MAIN (Main Menu): The operator is offered the following functions: power assurance,
administration, vibration, active exceedances, exceedance history, built-in test, data menu.
Active exceedances
Ø
Ø
Ø
The ACTIVE EXCEEDANCES screen displays IVHMS BIT failures during aircraft operations.
The screen will automatically display when the Data Transfer Memory Unit (DTMU) is low to notify the
operator.
When the BIT FAILURE ACTIVE message is displayed, press the ACK button then access the Built-In Test
screens to determine the failed component.
Built-in test
Ø
Ø
Ø
The SBIT/PBIT status displays the overall status of the IVHMS subsystems. When selected, the SBIT/PBIT
STATUS will consist of two screens.
Once entering the IBIT status screen the IBIT will automatically be initiated. An IBIT IN
PROGRESS message will appear and remain on until the IBIT is complete (approximately 2.5
minutes).
The status of components will be indicated by: PASS (system component OK), FAIL (system component
inoperative), ----- (invalid data), WAIT (test in progress), ? (BIT not initiated).
• HIT
The operator can complete the operational HIT check, by following the prompts provides on the
CDU.
Ø
If a parameter is out of limits, a message will be displayed on the bottom of the screen.
40
ATTITUDE HEADING REFERENCE UNIT (AHRU)
The AHRU is a self-contained strap-down AHRU that provides vehicle pitch, roll, magnetic heading, and
turn rate for integration within an avionics subsystem. It is capable of interfacing with flux valve (magnetic
detector unit), Doppler navigation set, HSI, VSI, and AFCS.
The AHRU contains an implemented Built-in test circuitry allowing the detection of 95% of system
failures. In case of a failure, the FAIL lamp on the control unit is illuminated. The BIT History Memory is
capable of storing the last 16 faults.
The BIT results are structured in 2 words: The LRU/SRU and the failure code.
The AHRU is capable of receiving mode selection from the control unit. The selection of operating modes
is done through the MODE and the ENT key on the control unit panel.
SLAVE mode
• When SLAVE mode is selected, the magnetic heading output becomes a gyro stabilized Magentic
Detector Unit (MDU) heading.
• Displays the current magnetic heading error, computed as the difference between the MDU
heading and the AHRU heading output. The error is displayed as a vertical bar on the control unit
display, moving between a minus (-) and a plus (+) sign.
• In SLAVE mode, three functions are enabled: BIT Status, Fast Erect Sync (during flight only), and
ETI Request.
BIT Status – This functionality is used to get from the desired AHRU to the BIT status.
Fast Erect Sync – May be activated airborne to perform a fast heading error reset. The GYRO
ERECT button on the MISCELLANEOUS SWITCH PANEL may also be utilized.
ETI Request – When activated, displays the current system ETI of the desired AHRU.
Free Gyro Mode (DG)
• In the DG mode, the AHRU output rate, stabilized attitude, and free gyro heading are corrected for
earth and aircraft rate.
• The system automatically detects the transition from ground-to-air at the first take off.
• In the DG mode, three functions are enabled: BIT Status, Heading Slew, and ETI Get Request.
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BIT Status – Same as SLAVE mode.
Heading Slew – This functionality allows the user to manually change the system heading.
ETI Get Request – Same as ETI Request.
Compass Calibration (CCAL) Mode
• The purpose is to perform MDU calibration and is capable of compensating the aircraft magnetic
heading as represented by the flux valve.
• CCAL mode can only be selected on the ground and is not possible to inadvertently select while in
flight.
Operation of the AHRU
The LOWER CONSOLE LT must be on the view the Digital Control Panel (DCP).
When AC power is applied to the system:
• The DCP will perform a lamp self-test for 2.5 seconds.
• After 2.5 seconds, the ALN lamp will be illuminated and the ATT fail flag on the VSI will be in view.
• The AHRU will align and self-test within 45 seconds.
• After 45 seconds, the system will default to the SLAVE mode and the ATT fail flag will
disappear. If a failure is detected:
• The FAIL lamp will illuminate and the display will show a message that the AHRU has failed.
• Press ENTER to clear the fail light.
If the fail light clears, continue the mission.
If the fail light does not clear, contact maintenance.
Electrical power interruptions
• A primary power interruption of less than 200 milliseconds does not cause system realignment.
• A primary power interruption of more than 200 milliseconds will cause system in-air realignment.
The ATT and NAV flags will be in view to indicate system realignment.
During the realignment, maintain straight and unaccelerated flight, for a minimum of 45 seconds.
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