FLIGHT LINE

SUPPLEMENT UH-60M
BLACKHAWK

F Co. 1-212th Aviation Regiment

Fort Rucker, AL 36362-5000
Version 1.0

October 2010

CVS
 UH-60M Flight Line Supplement
Table of Contents

HYDRAULIC SYSTEM 3
ELECTRICAL SYSTEM 10
FUEL SYSTEM 13
T701D ENGINE 15
DIGITAL ELECTRONIC CONTROL (DEC) 16
ENGINE ALTERNATOR 19
HYDROMECHANICAL UNIT (HMU) 19
MECHANICAL MIXING UNIT (MMU) 21
ELECTRONIC COUPLING 21
AUTOMATIC FLIGHT CONTROL SYSTEM (AFCS) 22
STABILITY AUGMENTATION SYSTEM (SAS) 25
TRIM SYSTEM 25
FLIGHT PATH STABILIZATION (FPS) 27
COUPLED FLIGHT DIRECTOR 30
STABILATOR 31
PNEUMATIC SYSTEM 34
AIRCRAFT QUICK REFERENCE MATERIAL 35
AIRCRAFT ANTENNA LAYOUT 37

The purpose of the flight line supplement is to provide basic supplemental information about the
UH-60M to aviation students. The references used for this supplement include:

TM 1-1520-280-10 (Operators Manual)

Sikorsky Training Information for the UH-60M (IKTB)
TM 1-2840-248-23 (Aviation Unit and Intermediate maintenance Manual)
UH-60M Transition Student Handout
Sikorsky Operations and Screens Document for the Flight Management System

Any questions or comments about the information contained in this supplement should be
addressed to:

Commander
F Co. 1-212th Aviation Regiment
ATTN: Standardization Officer
Fort Rucker, AL 36362

2
 HYDRAULIC SYSTEM
The #1 and #2 hydraulic pumps are driven by separate transmission accessory gearbox modules
and supply 3000 PSI hydraulic pressure when the main rotors are turning. The backup hydraulic
pump is powered by a 115 VAC electric motor. If AC power is available, the backup pump is
able to provide hydraulic pressure should one or both hydraulic pumps fail.

Each hydraulic pump is a constant pressure, variable volume pump. They maintain the required
3000 PSI over any rate of control movement for their own system. However, if the #1 and #2
hydraulic pumps are inoperative at the same time, there will be a slight restriction in the
maximum rate of flight control movement. This is due to the backup pump's inability to provide
the required volume because it exceeds the flow rate capability of the pumps.

Each pump has a pressure and return filter. The filters have indicator buttons that extend when
the filter is clogged. Additionally, the return filter has a bypass capability, allowing hydraulic
fluid to return from the system in the event of a filter clog. If a return filter were to become
completely clogged and have no bypass capability, fluid would be blocked in the return lines
causing a hydrostatic flight control lock up.

The #1 hydraulic pump supplies hydraulic pressure to the #1(first stage) fore, aft, and lateral
primary servos and the #1 tail rotor servo.

The #2 hydraulic pump supplies hydraulic pressure to the #2(second stage) fore, aft, and lateral
primary servos and the components of the pilot assist area; pitch, roll, and yaw SAS actuators,
pitch boost servo, collective and yaw boost servos, and the pitch trim assembly. A pressure
reducer in the pilot assist module reduces the 3000 PSI pump pressure to 1000 PSI for use by the
pitch trim servo.

The backup hydraulic pump can supply hydraulic pressure to the #1 and/or #2 hydraulic systems
independently or simultaneously. It is also the only pump that can supply hydraulic pressure to
the 2nd stage tail rotor servo and the APU accumulator. The backup pump supplies hydraulic
pressure to flight control components during ground checks (main rotors not turning). The
backup pump will come on and provide pressure when any of the following conditions occur:

#1 Reservoir Low (caution light) #2 Hydraulic Pump (caution light)

#1 Hydraulic Pump (caution light) APU Accumulator Low (advisory light)
#1 Tail Rotor Servo (caution light)

An automatic turn off feature is provided through a switch marked SERVO OFF, 1ST STG-
NORM-2ND STG located on the MISC SWITCH panel on the lower console. In the NORM
position both stages of the primary servos are on and operating. When the SERVO OFF switch
is moved to the 1ST STG or 2ND STG position, the primary servos associated with that system are
turned off. When the SERVO OFF switch is moved to 1ST STG, fluid to the #1 forward, aft, and
lateral servos are turned off at the transfer module. A loss of pressure in the #2 primary servos
will cause the first stage, turned off but has pressure, to automatically turn back on in case the
backup system does not take over the function of the failed second stage. If the second stage is
initially turned off, the sequence is reversed.

3
 UH-60M #1 HYDRAULIC SYSTEM

RSVR 1 LOW

HYD PUMP 1 FAIL

T/R SERVO 1 FAIL

PRI SERVO 1 FAIL

4
Note: The SAS/BOOST switch on the AFCS panel is new to the UH-60M. The switch activates or deactivates
hydraulic pressure to the SAS and Collective, Yaw, Pitch boost assemblies. This engineering change allows
one switch operation of the hydraulic features for these systems.

TRIM FAIL
BOOST SERVO

SAS OFF
OFF

FCC 2
FPS FAIL
UH-60M #2 HYDRAULIC SYSTEM

HYD PUMP 2 FAIL

RSVR 2 LOW
SAS/BOOST
PRI SERVO 2 FAIL

5
 UH-60M BACKUP HYDRAULIC SYSTEM
Conditions that activate the back-up
pump
RSVR 1 LOW HYD PUMP 1 FAIL T/R SERVO 1 FAIL HYD PUMP 2 FAIL APU ACCUM LOW

APU ACCUM LOW

T/R SERVO 2 ON BACK UP PUMP ON BACK UP RSVR LOW

6
 HYDRAULIC LEAK DETECTION/ISOLATION SYSTEM.

The leak detection/isolation (LDI) system protects the flight control hydraulic system by
preventing the further loss of hydraulic fluid in case of a leak. The LDI system uses pressure
switches and fluid level sensors for monitoring pump hydraulic fluid level, and pump pressure
for primary and tail rotor servos, and pilot-assist servos. When a pump module reservoir fluid
level switch detects a fluid loss, the logic module follows the sequence detailed in the figures on
pages 8 and 9 to isolate the leak. To accomplish this, the logic module operates the required
shutoff valve(s) to isolate the leak and turns on the backup pump when required. In the cockpit
the RSVR LOW caution for that system appears. Backup pump and shutoff valve(s) operation is
automatic through the logic module. If after the isolation sequence, the leak continues, the
leakage is probably in the stage 1 or 2 primary servos and the SERVO OFF switch must be
moved to the 1ST STG or 2ND STG position by the pilot.

Hydraulic fluid levels may be up to 3/8” into the overfill zone (black or blue depending on type
of reservoir) when fluid is hot.

There are temperature sensitive labels used on all hydraulic pumps. There should be one label
on the pressure side and one label on the return side. When the temperature exceeds 132°C
(270°F) on the label an entry shall be made in the on DA Form 2408-13-1. The aircraft should
not be flown until appropriate maintenance action has been taken.

Rotor Brake starts requires a minimum of 450 psi. If

450 psi is not attained do not attempt a Rotor Brake
start. Do not service the Rotor Brake with the
servicing hand pump, instead contact maintenance for
assistance.

To Rotor Brake

The hydraulic pump modules on the aircraft consist of a combination hydraulic pump and
reservoir that is sealed from the atmosphere. A spring in the reservoir maintains a positive
pressure in the system at all times. Because the system is sealed, a servicing hand pump and
reservoir are incorporated to allow for replenishment of hydraulic fluid. A selector valve allows
the servicing hand pump to be connected to the #1 hydraulic pump, the #2 hydraulic pump, or
the backup hydraulic pump (labeled #3 on the selector valve) or to the rotor brake, #4. To fill a
hydraulic reservoir, select the desired position, and while depressing the selector valve, rotate the
servicing hand pump handle in the direction indicated on the reservoir. Use caution not to
overfill the reservoir being serviced.

7
 Logic Modules

Two logic modules, one in the left relay panel and the other in the right relay panel, are used to
control the operation of the hydraulic systems. The logic modules continually monitor the
operation of the hydraulic systems by inputs received from pressure switches, fluid level
switches on the pump modules, and inputs received from control switches in the hydraulic
system. The outputs of the logic modules will either activate cautions/advisories notifying the
pilot of a failure and/or turn off one or more valves due to a system malfunction. All switching
functions of the hydraulic logic modules are automatic, except as shown by a dagger (†) in the
figures which indicates a required crewmember action to complete the step. (Figure 2-20).
8
9
 ELECTRICAL SYSTEM
Alternating current (ac) is the primary source of power. The primary electrical system consists
of two independent systems, each capable of supplying the total helicopter power requirements.
The prime source of each system is a 115/200 vac generator. A subsystem feeds two
independent ac primary buses and a 26 vac ref bus. A portion of each ac primary bus load is
converted to 28 volts direct current (vdc) by two 400 ampere ac/dc converters.

The main AC electrical system remains largely unchanged from the UH-60A/L system
configuration. However, due to the electrical requirements of modern electronics, UH-60A/L
AC Essential Bus has been eliminated and a 26 VAC Reference Bus is incorporated to provide
stable reference for use by stabilator system. Additionally, an autotransformer is added for
system redundancy.
The DC and battery electrical systems have been impacted the greatest. The 200 amp DC
converters have been replaced by 400 amp converters. A new battery system has been designed
and implemented. Some contactors have been replaced with newer generation components and
circuit breaker panels and various relays have been modified or revised to reflect changes in
aircraft systems.
The two new 24 VDC 5 amp-hour SLAB batteries provide power to independent Battery,
Battery Utility, and DC Essential buses and also provide for uninterrupted emergency power to
essential systems for approximately 15 minutes. (There is approximately 12 minutes of battery
power available at night with all dc electrical equipment on including the search light.) Under
batteries the following items are operational;
• LIGHTS: Standby Compass Light, Cockpit Utility Lights, Controllable
Searchlight (use during landing only).
• AVIONICS: No. 3 Multi-Function Display (MFD)(pilot inboard), Pilot Flight
Management System (FMS), Pilot Flight Director/Display Control Panel
(FD/DCP), ICS system, COM 3 (AN/ARC-231), Transponder, Number 2
Embedded GPS/Inertial Navigation System (EGI) (pilot), Number 2 Data
Concentrator Unit (DCU) (pilot), VOR/ILS, COM 1 VHF-FM (without power
amp), Radar Altimeter, Number 2 ADC, Audible Warning System, battery
system/DC essential relay contactors, digital clock.
• APU: APU system control, APU/engine fire detection systems, APU/engine fire
detection extinguisher system.
• MISC Systems: Manual Stabilator Slew Control (#1 Stab. Actuator), SAS turn-
off valve control, Fuel shut-off valve control, Crashworthy External Fuel System
transfer control valve, Backup Hydraulic System/pump control, ESSS emergency
jettison system, Cockpit Voice Recorder/Flight Data Recorder (CVR/FDR),
Cargo Hook emergency release system, Rescue Hoist cable shear.
• The battery system now provides power directly to all buses to hold up bus power
during power switching transient events.
• The system is designed so that either battery can power both DC Essential Buses
and Battery Buses.

10
A Battery System Cold Start Status Indicator/Caution Light, has been implemented with the
removal of the Caution/Advisory panel. The panel provides indications for APU status, battery
condition, and indications for emergency release tests. One of the four cold start status indicator
capsules is dedicated to providing battery status indications as either BATT GOOD (green) or
BATT LOW (yellow). Battery condition may be checked at any time by placing the battery
power switch(s) to the TEST position.
When the switch is placed to the TEST position a load of approximately 4 amperes is applied on
the battery. The BATT GOOD (green) or BATT LOW (yellow) status light on the Cold Start
Status Indicator will illuminate indicating the status of the battery. If the batteries are turned ON
and voltage drops below 20.5 VDC, BATT LOW CHARGE caution light will illuminate.
BATT LOW CHARGE caution will disappear and batteries begin to recharge when AC power
is applied. The caution light will not appear in flight unless both converters drop offline and the
batteries are the sole source of power.
An auxiliary AC power system is provided through the use of an auxiliary power unit driven air
cooled generator. The generator is 3-phase 115/200 VAC, providing 20/23.8 KVA of power.
This is the same generator used in the UH-60A/L. The output specifications have been redefined
as 20/23.8 KVA continuous, with a peak 5 minute transient output of 30 KVA. A dedicated
APU Generator GCU, like the main generator GCUs, monitors generator output through the use
of current transformers.

The APU generator is only capable of providing power to the aircraft electrical system when the
main generators are not online. When the GENERATORS APU TEST-OFF/RESET-ON
switch is placed to ON, power is supplied to the No. 2 AC Primary Bus and routed through a set
of current limiters (CL) to power the No. 1 AC Primary Bus. The EICAS APU ON advisory
will indicate that the APU is on and operating. Operating limits have not changed. If the APU
Generator is the only source of AC power, the windshield anti-ice system cannot be run while
the Back-up Pump is operating. Blade de-ice operation still requires that at least one main
generator be online and operating.

11
12
 FUEL SYSTEM
The engine fuel supply system consists primarily of the low pressure suction engine driven boost
pump, fuel filter, fuel filter bypass valve, fuel pressure sensor switch, hydro mechanical unit
(HMU), and overspeed and drain value (ODV).

Fuel is drawn from a main tank up through the fuel selector valve and proceeds through the low
pressure suction engine driven boost pump, fuel filter, HMU, liquid to liquid cooler, then to the
ODV, out to the main fuel nozzles for starting and engine operation.

Fuel from the main tanks is by design provided by the low-pressure suction engine driven boost
pumps, resulting in a negative pressure in the fuel lines. Two electrically operated submerged
boost pumps provide positive pressure from the main tanks.

An automatic engine fuel prime feature is activated during engine start and then 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.

FUEL LOW caution light(s) and the MASTER CAUTION light on the Master Warning Panel
flashes when the fuel in the respective cell reaches approximately 172 pounds. Additionally,
when a main tank level falls below 200 pounds, the MAIN FUEL tape and readout turn yellow.

Fuel tanks are crashworthy and ballistic-resistant. The fuel line network includes self-sealing,
breakaway valves, and all main engine fuel lines are self-sealing at the fittings. The APU fuel
lines are not self-sealing.

The Fuel System Selector Levers allow:

_ DIR - Fuel is drawn to the respective engine (#1 fuel tank to #1 engine, and #2
fuel tank to #2 engine).

_ XFD - Allows the engines to draw fuel from the opposite tank (#1 engine draws
fuel from the #2 fuel tank, and #2 engine draws fuel from the #1 fuel tank).

_ OFF - Fuel to the main engines is shut off (valves closed). Fuel for the APU is
drawn from the #1 fuel cell ONLY.
14
 T701D Engine
The T700-GE-701D 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. Through upgrades in materials and design the T701D provides better
performance in hot and high altitude conditions.

The T700-GE-701D engine is derived from the T700-GE-700 engine (85% common parts) with
aero-thermo-dynamic improvements necessary to achieve an increase in maximum power. The
engine maintains the same basic structure and configuration. Improvements were made
primarily in the Compressor Rotor Assembly, the Hot Section Module, and the Power Turbine
Module. Improvements include a Digital Electronic Control (DEC) which replaces the Electrical
Control Unit (ECU). Other improvements include:
An Overspeed Drain Valve (ODV) replaces the Pressurizing Overspeed Unit (POU).
A History Counter replaces the History Recorder.
An additional Nr sensor (located on the left transmission accessory module) and Collective
Position Transducer (located on the aircraft mixing unit) have been added to work with Transient
Droop Improvement. (The engines are more responsive to aircraft maneuvers by providing a
collective and Nr rate compensation.)

A Hot Start Prevention (HSP) feature has been added to the DEC.
Compressor Rotor Assembly: Changes to the Impeller and Diffuser enhance performance and
provides additional ventilation via Compressor Rear Shaft/Gas Generator Shaft to the #4 bearing.
Hot Section Module: Improved materials and increased part cooling efficiency allow an increase
in operating temperature.
Power Turbine Module: Improved one piece casting design, air flow improvement and additional
C-sump oil jets provides increased performance.
 DIGITAL ELECTRONIC CONTROL - DEC
a. The DEC replaces the analogue ECU. The DEC contains a micro-computer processor in a
conductive composite molded case. It is mounted below the compressor casing in the same
location as the ECU. The DEC does not have an insulation blanket and is cooled entirely by
airflow through the scroll case. It provides the same basic functions as the ECU.
b. Two (2) Power Sources:
1. Primary: Engine Alternator.
2. Secondary: 400 Hz, 120 Vac Airframe Power.
Complete loss of engine alternator will result in a loss of Ng signal (with corresponding engine
out audio and light). Airframe power will sustain the DEC and there should be NO failure of the
engine to the high side, or loss of TQ and Np indications.
c. Accepts inputs from other sources for electronic control of the engine:
1. Engine Alternator.
2. Thermocouple Harness.
3. Np Sensor.
4. Torque and Np Overspeed Sensor.
5. Torque signal from the other engine.
6. Collective Position Transducer (New – mounted in the mixing unit).
7. Nr Sensor – (New – mounted on left Accessory Module).
8. Demand Speed from Engine Speed Trim Switch.
9. Feedback signal from the Hydromechanical Unit (HMU) for system stabilization –
Linear Variable Differential Transformer (LVDT).
d. Provides the following signals to the cockpit:
Np – Initial indication is at 4000 RPM vs. 6000 RPM with the ECU, permitting earlier
signal during engine run-up.
Torque – Torque signal locked to zero until Np reaches 35%. This logic eliminates
torque spike signal during engine start and shutdown. This prevents yaw excursions
caused by load sharing response to torque spike during in-flight restart of an engine.

See single source document for DEC functions

Additional/Modified Operational Features:
16
1. Hot Start Preventer (HSP):
Automatically prevents overtemperature during engine starts by shutting down the
engine.
Requires 400 Hz, 120 Vac power for operation (APU generator).
Requires signals from the Np sensor, Ng Alternator Winding, and the Thermocouple
Harness (TGT).
HSP activates a solenoid in the ODV which shuts off fuel flow to the engine when Np
and Ng are below their Hot Start Reference Values and the TGT exceeds its reference
value.
1. Ng < 60%. 2. Np < 50%. 3. TGT > 900C.
Pilot then places Engine Power Control Lever OFF and turns ignition off. Before
attempting a restart the TGT must decrease below 300°C or wait 25 seconds.
In an emergency HSP can be disabled by pressing and holding either of the Np overspeed
test buttons during the start. To ensure the system is disabled throughout the full start
sequence, hold button in until Ng > 60%.

2. Transient Droop Improvement (TDI)

Engines more responsive to aircraft maneuvers by providing a collective and Nr rate
compensation.
Signals required:
1. New Nr Sensor mounted on the left Accessory Module.
2. New Collective Position Transducer mounted with two previous transducers
on a modified bracket assembly in the mixing unit.
3. Engine torque signal.
The Nr and collective signals are routed to rate circuits in the DEC and compared with
Np and TRQ signal values. A priority signal is then processed to quickly trim the Torque
Motor in the HMU to compensate for the demand.
Transient Droop Improvement is 4:1 over the T700-GE-700 engine. If a sensor fails or a
malfunction occurs then the system reverts back to the basic load demand system similar
to that provided by the ECU.

3. TGT Limiting:
The DEC commands the HMU to limit fuel flow to the engine when TGT reaches the
limiting value. TGT Limiting does not prevent overtemp during compressor stalls,

17
 engine starts, or when the engine power control lever is in LOCKOUT. When maximum
power is demanded transient increases above the limiting values can be expected to
occur. TGT limiting results in the following:
Dual Engine Limiting Value: 879 +/- 9°C. (UH-60M) source TM 1-2840-240-23 & P
Contingency Power Limiting Value: 903 +/- 10°C. (UH-60M) source UH60M MTF
1. Automatic contingency power is provided by the DEC by resetting the TGT
Limiting to a higher value.
2. When other engine is inoperative (engine failure or torque split).
3. When torque on other engine is below 50%.
4. More power for single engine operation (approximately 5%).

4. Np Overspeed Protection System:

Primary electrical source is the engine alternator and the secondary source is 400 Hz, 120
Vac airframe power. System is tripped when Np reaches 120% +/- 1% (25,000 RPM).
When an overspeed is sensed by both Overspeed Sensing Circuits A and B, the
Overspeed Protection Relay will activate the Overspeed Solenoid in the ODV shutting off
fuel, and sends a signal to the Delay On Drop Out (DODO) Circuit to turn on the
ignition. When Np drops below 120% +/- 1%, the Overspeed Solenoid Valve
automatically closes allowing fuel flow to return to the engine.
When the Overspeed Solenoid Valve closes, resupplying fuel to the engine, ignition will
come on (channeled through the Overspeed Relay) and remain on for 5 seconds to allow
for restart.
Np Overspeed Protection is still functional during DEC lockout operations. DEC
Lockout Procedures are the same as with the ECU.

OVERSPEED DRAIN VALVE (ODV) replaces the Pressurizing

Overspeed Unit (POU).
1. The ODV sends fuel through main fuel manifold to the 12 main injectors.
There is no start fuel manifold or primer nozzles on the 701C/D.
2. The ODV purges fuel from the main fuel manifold on shutdown to prevent
coking.
3. The ODV provides Np Overspeed Protection/Hot Start Prevention. This is
accomplished by cycling all fuel back to the HMU inlet and shutting off fuel flow
to the engine.

18
 ENGINE ALTERNATOR
The engine alternator supplies ac power to the ignition exciter and DEC. It also supplies the NG
signal to the cockpit. All essential engine electrical functions are powered by the alternator.
When the alternator power supply to the DEC is interrupted, 400 Hz 120 VAC helicopter power
is utilized to prevent engine (high side) failure. There will be a loss of the associated cockpit NG
indication and activation of the ENG OUT warning and audio will occur. Overspeed protection
is still available. When the alternator power supply providing the NG signal is interrupted, a loss
of the associated engine NG indication and activation of the ENG OUT warning and audio will
occur.

HYDROMECHANICAL UNIT - HMU

Control of fuel to the combustion system is provided by the HMU. The HMU, mounted on the
rear center of the accessory gear box, contains a high pressure pump that delivers fuel to the
ODV. Various parameters are sensed by the HMU and influence fuel flow, variable geometry
position, and engine anti-ice start bleed valve operation. There are two mechanical inputs from
the cockpit to the HMU. The load demand system provides mechanical input directly from the
collective output of the flight control mixer to the engine HMU to reduce transient engine and
NR droop. This system acts as a load anticipator that immediately inputs a load change to the
HMU, which resets NG as required to maintain a nearly constant NP. This system makes gross
corrections which are fine tuned by the load sharing and power turbine governing functions of
the DEC. During emergency operations, when the ENG POWER CONT lever is moved to
LOCKOUT and then to an intermediate position, the engine will still vary power in response to
collective signals. Collective movement is transmitted directly to the mechanical flight control
mixer. A bellcrank on the mixer provides input to the push-pull load demand spindle (LDS)
cables, and then to the engine HMUs via a rotary actuator. The rotary actuator is spring loaded
to increase the collective direction to minimize system freeplay, and will also drive the LDS to
maximum demand (corresponding to full up collective), in case of a separation in the system,
resulting in full engine power available. At the mixer end of the cable, a shear pin/roll pin is
provided to prevent a flight control jam in the event of a load demand system malfunction. If the
pin shears, the LDS cable to the HMU for that engine will drive full aft and produce the
maximum LDS demand. This may under certain flight conditions produce a torque split. If the
pin shears and the load requirement of the engine is reduced (e.g., an autorotation) the DEC will
not have enough down trimming authority to reduce torque to zero. Under this condition the
excess engine torque produced may be sufficient to drive the engine or NR to NP1 or NP2
overspeed.

The ENG POWER CONT lever operates the Power Available Spindle (PAS) in the HMU.
When the engine is operating in FLY, the LDS and DEC control have maximum power available
to be utilized. If the ENG POWER CONT lever is placed in LOCKOUT, the NP1 or NP2
servo normally controlled by the DEC is locked out to its maximum power position. Engine
power must now be set manually by adjusting the ENG POWER CONT lever. The engine will
still increase/decrease power in response to (LDS) collective changes. DEC functions which
19
utilize the NP1 or NP2 servo (TGT limiting, NP1 or NP2 governing, LOAD sharing) are
inoperative in LOCKOUT.

Fuel from the HMU flows to the ODV. The ODV sends fuel through the main fuel manifold to
the injectors or starting, acceleration and engine operations. It purges fuel from the main fuel
manifold and allows back flow of high pressure air for purging. It shuts off fuel flow to prevent
an engine overspeed when the overspeed system is tripped as sensed by the DEC. It also shuts
off fuel to prevent hot starts when activated by the Hot Start Preventor.

See single source document for HMU functions

NOTES concerning engine indications

The change in the engine oil temperature lower limit of (-20°C) is not an actual temperature /
functional limit change; it is based on the physical reporting limits of the Engine Instrument
Caution Advisory System (EICAS). The indication is not reportable below (-20°C).
The UH-60M DEC bias has been changed to 112°C, this provides for similar displayed limit
indicating numbers when compared to the UH-60L. The UH-60L was biased to show lower than
actual indicating numbers on the central display unit due to the display’s inability to show four
digit numbers. On the UH-60M TGT display an indication of 903°C is in reality 1015°C actual
temperature.

20
 MECHANICAL MIXING UNIT - MMU
The mechanical mixing unit is designed to reduce inherent control coupling. (Inherent control
coupling might be referred to as a side effect of a control input. You increase collective to bring
the aircraft up to a hover. However, a side effect of this is the increased torque causing the nose
to turn to the right). This is an average mixing for all flight conditions and is not based on any
particular weight. The pilot, with assistance from AFCS, will make the adjustments needed if
these mixings are too much or too little. The mechanical mixing unit helps reduce pilot
workload through the following inputs and outputs:

1. Collective to Pitch mixing Compensates for downwash on the aft fuselage and
stabilator. As collective is increased, the main rotor disk is tilted forward slightly, and as
collective is decreased the main rotor disk is tilted aft slightly.

2. Collective to Roll mixing Compensates for translating tendency. As the collective is

increased the main rotor disk is tilted slightly left and as the collective is decreased the main
rotor disk is tilted slightly back toward the right.

3. Collective to Yaw mixing Compensates for the torque effect of the main rotor. As
collective is increased, tail rotor pitch is increased and as collective is decreased, tail rotor pitch
is decreased.

4. Yaw to Pitch mixing Compensates for tail rotor lift vectors. As tail rotor pitch is
increased the main rotor disk is tilted slightly aft, and as tail rotor pitch is decreased the main
rotor disk is tilted slightly forward. The tail-rotor on the UH-60 is mounted to provide a vertical
lift component to help offset the aft CG component associated with airframe design.

In the four mechanical mixings the first word in the title is the pilot input to the flight controls.
The second word in the title is the mixing unit output.

Electronic Coupling
Collective/Airspeed to Yaw This mixing is in addition to collective to yaw mechanical mixing.
It helps compensate for the torque effect caused by changes in collective position. 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 Flight Control
Computer (#2 FCC) 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.
 Automatic Flight Control System - AFCS

AFCS enhances the stability and handling qualities of the helicopter and provides autopilot
functions.
The AFCS is comprised of five basic subsystems:
• Stability Augmentation System (SAS) – provides short-term rate dampening in
pitch, roll, and yaw axes.
• Trim – provides control positioning and force gradient functions.
• FPS – basic autopilot functions when coupled with trim.
• Coupled Flight Director – provides a means for the pilot to select defined
autopilot functions, automatic flight control positioning, and provides steering
cues for display on the MFD when coupled.
• Stabilator – improves flying qualities with the use of electromechanical actuators
in response to collective, airspeed, pitch rate, and lateral acceleration inputs.

Flight Control Computers - FCCs

The central component of the AFCS is two redundant Flight Control Computers. Each computer
receives signals from the pilots’ controls, motion sensors (accelerometers), control panels,
avionics systems, embedded GPS/Inertial navigation, and air data computers to compute
commands which are sent to the cockpit, trim actuators, SAS actuators, and stabilator actuators.
The No. 2 FCC is the only FCC that sends signals to the trim actuators for dynamic flight control
inputs.
The No. 1 FCC and No. 2 FCC are used in the dual stabilator units. The FCCs are responsible
for processing signals for airspeed, lateral acceleration, collective stick position, and pitch rate
data.
The FCCs contain power supply, actuator drive circuits, processing and feedback circuits, and
fault monitor circuits.
By communication over the cross channel data link (CCDL) of automatic mode engagement,
each FCC will know the status of the other. When one fails from loss of power or fault
detection, automatic mode is disengaged.
Lateral, Longitudinal, and Vertical Accelerometers provide information to the FCCs concerning
aircraft acceleration and deceleration. This information is processed for use by the AFCS to
improve aircraft stability.

22
Collective Stick Position Sensors are operated by movement of the collective linkage and send a
signal to the FCCs indicating collective stick position. The No. 1 and No. 2 collective stick
position sensors generate an output signal proportional to the collective stick position. This
information is used to help determine stabilator position. Additionally, the same signals are used
for other aircraft systems in the UH60M.
Trim Servos are used in the pitch, roll, yaw, and collective to allow the FPS and Coupled Flight
Director to make the necessary inputs to maintain the desired pilot selections.
Air Data Computers (ADCs) provide airspeed, barometric altitude, and barometric rate
information to the FCCs. The ADCs connect to the Pitot-Static system. Both provide data that
contains information concerning the helicopter airspeed, altitude, and altitude rate. The
stabilator uses the airspeed data. The airspeed data for the No. 1 FCC is generated by the No. 1
ADC. The airspeed data for the No. 2 FCC is generated by the No. 2 ADC. An ADC is located
on each side of the forward cabin in the flight control broom closet.
The Embedded Global Positioning/Inertial Navigation Unit (EGI) replaced the Directional Gyro,
pilot and copilot Vertical Gyros, compass control panel, and compass flux valve to include the
standard Doppler System of the UH-60 A/L. The primary function of the EGI is to compute
attitude, heading, present position, and turn rate data. This information is provided to the FCCs
from two self-contained EGI systems consisting of an inertial navigation system and an
embedded global positioning receiver. The EGIs are used in the stabilator system. Each EGI
provides data that contains information concerning the pitching rate of the helicopter. Each EGI
pitch rate data is fed to its respective FCC where it is conditioned. The filtered pitch rate data is
used by the stabilator to enhance the AFCS system’s ability to correct short term pitch
disturbances.
Both the inertial and GPS determine their position independently. The GPS determines its
position from the GPS satellite constellation. The GPS receiver determines three-dimensional
worldwide position coordinates. The GPS provides for all-weather continuous operation based
on the principle of satellite ranging. Satellite ranging is the process of measuring the time
satellite signals travel from the satellite to the helicopter GPS antennas. By ranging to four
satellites, three-dimensional position can be determined.

The inertial position is determined by speed/heading integration over time applied to a known
starting position. For the inertial system to work correctly it must have a known starting
location, normally provided by the GPS. Both GPS and inertial systems locations are compared
by the EGI for relative errors to determine the rate of correction required for the inertial position.
These corrections are applied to provide the blended position. Inertial systems are generally self
contained and do not emit an external signal. However, their accuracy can drift if not
maintained. In the event of GPS failure or loss of satellite constellation, navigation could
continue with the inertial system alone. However, updates for the inertial system would be
required.

The EGI provides information for other aircraft systems to support navigation, time distribution,
flight control, and primary flight displays. Navigation signals provide navigational information
23
to the pilots through the MFDs and provide reference to control attitude and heading of the
aircraft through the Automatic Flight Control System (AFCS).
The Flight Control Computers commands inner-loop SAS actuators and the outer-loop trim
actuators in pitch, roll, and yaw control channels. The computers also provide self-monitoring,
fault isolation, and failure advisory. The AFCS provides two types of control, identified as
inner-loop and outer-loop. The inner-loop (SAS) employs rate damping to improve helicopter
stability. This system is fast in response, limited in authority, and operates without causing
movement of the flight controls. The outer-loop (Trim) provides long-term inputs by trimming
the flight controls to the position required to maintain the selected flight regime. It is capable of
driving the flight controls throughout their full range of travel (100% authority) at a limited rate
of 10% per second. Both inner and outer loops allow for complete pilot override through the
normal use of the flight controls. The FCCs process incoming information from various sensors
aboard the aircraft and stores this information in its memory. The sensor information is used by
the computer Central Processing Unit (CPU) to compute required correction signals. Inner-loop
correction signals are routed to the SAS actuators and outer-loop signals are routed to trim servos
and actuators.

24
 STABILITY AUGMENTATION SYSTEM - SAS
The UH-60M incorporates two SAS systems to help maintain a stable platform in flight. SAS
responds to short-term aerodynamic disturbances such as wind gust or up and down drafts by
effectively dampening 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. SAS enhances dynamic stability in the pitch,
roll, and yaw axes. (Prevents pitching in the longitudinal axis, rocking in the roll axis, and fish
tailing in the yaw axis.) SAS also enhances turn coordination at airspeeds greater than 50 knots.
(SAS 2 enhance turn coordination by deriving commands from lateral accelerometers which
together with roll rate signals are sent to their respective yaw channels automatically at airspeeds
greater than 50 knots)
The FCCs receive their respective on-side EGI signals for SAS actuator movement. A
malfunction may be detected as an erratic motion in the helicopter. If the malfunction is of an
intermittent nature, the indication can be cleared by pressing FAILURE RESET CPTR 1 and/or
CPTR 2 switches. 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. With SAS 1 or
SAS 2 off, the gain of the remaining SAS is doubled. (Doubling the gain simply means that the
sensitivity of the remaining SAS has been doubled.)
SAS Actuator hydraulic pressure is monitored. In case of loss of SAS actuator hydraulic
pressure, or if both SAS 1 and SAS 2 are turned off, the SAS OFF caution indication will appear
on the MFD. In the latter case, FPS will automatically turn off and the FPS ON segment light
will extinguish. The SAS DEGRADED advisory will appear when the pitch, roll, and/or yaw
SAS channels from SAS 1 or SAS 2 has disengaged due to a fault.

Trim System
The Trim units provides cyclic (pitch and roll), pedal (yaw), and collective flight control position
reference and control gradient. It maintains the cyclic stick, tail rotor pedals, and collective stick
at a desired position. Each trim actuator and servo contains a force gradient spring to permit
pilot override of trim reference. The Trim units may be 100% overridden through use of the
flight controls. TRIM is engaged by pressing the TRIM switch on the stabilator control/auto
flight control panel.
The Set consists of the:
• Pitch trim servo, on the top deck.
• Roll trim actuator, on the top deck.
• Yaw trim actuator, on the top deck.

25
• Collective trim actuator, on the top deck.

Trim, when coupled with FPS provides the following:

Below 50 KIAS Above

Att. Hold Pitch Att. Hold

Att. Hold Roll Att. Hold

Hdg. Hold Yaw Heading hold/Tn. Coordination

Cyclic Trim Beeper

The cyclic trim beeper is new to the UH-60M. The cyclic trim beeper is a five-way trim switch
on each cyclic stick that 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
cyclic trim beeper switch in one direction at a time. When FPS is engaged, the cyclic trim beeper
switch changes the pitch and roll attitude reference. Cyclic trim beeper activation in any of four
directions trims the cyclic stick toward the desired location. Release of the cyclic trim beeper
stops cyclic movement. The fifth movement of the cyclic trim beeper is a Z axis, or plunge
feature. The plunge feature is activated by simply pushing in on the trim beeper. If the aircraft is
hovering when the plunge feature is selected automatic radar altitude engagement, on the
FD/DCP, will be maintained over the selected location. The HVR mode has three states:
velocity hold (HVR VHLD), position hold (HVR PSN), and deceleration mode (HVR DECL).
The HVR mode will always attempt to engage coupled on a coupled FD. If neither FD is coupled
the FD will attempt to engage HVR on the FD on which the mode was selected from the
FD/DCP or cyclic trim beeper switch. If the plunge command is selected above 60 knots
barometric altitude engagement, on the FD/DCP, instead of the radar altitude engagement will
occur. In this mode the plunge command will engage IAS mode at 80 knots if between 60 and
80 knots, 120 knots if above 120 knots, and current airspeed if between 80 and 120 knots.

26
 Collective Trim Beeper
A heading change can be commanded through the use of a 4-way trim switch
on each collective stick. Actuating this switch L (left) or R (right) below 50
knots will cause the tail rotor pedals to drive until the helicopter attains a yaw
rate of approximately 3° per second (Standard rate turn). When trim is
released, heading hold is re-engaged to hold the new heading. Actuating the
trim switch DN (down) or UP will adjust the reference altitude for collective
flight director functions. When trim is released, altitude hold is re-engaged to
hold the new altitude.

The collective trim beeper is deactivated any time the heading select mode is engaged.

Flight Path Stabilization - FPS

Fight Path Stabilization provides basic autopilot functions in the pitch, roll, and yaw axes.
Proper operation of FPS requires that the SAS/BOOST, TRIM, and SAS 1 and/or SAS 2
functions be selected on the control panel. Performance of the system will be improved by the
proper operation of the stabilator in the automatic mode. The design of the system is such that
the FCCs, through FCC 2, provide FPS command signals to the pitch, roll, and yaw trim
actuators which in turn reposition the flight controls. The AFCS provides command signals to
the trim actuators to reposition the flight controls using the trim system. (Trim is the muscle in
the system. The AFCS moves the flight controls through the use of trim and allows the aircraft
the ability to fly and maintain a predetermined flight profile.) FPS may be disengaged by
pressing the TRIM switch on the AUTO FLIGHT CONTROL panel. However FPS must be
reselected individually. (In other words, the FPS will not work without the trim system and if
trim is turned on again the FPS must also be reselected.)
The No. 2 FCC has sole input to the H-60M Trim actuators. All commands to move the trim
actuators are provided by the No. 2 FCC. The No. 1 and No. 2 FCC are identical. However, the
No.1 FCC does not have the ability to direct movement of the flight controls.

Flight Path Stabilization (ROLL)

At airspeeds greater than 50 knots heading hold will be automatically disengaged and turn
coordinated engaged under these conditions:

27
 TRIM beeper is actuated in the lateral direction.
TRIM REL switch is pressed and roll attitude is greater than 2.5 degrees.
About 1⁄2 inch cyclic displacement and a roll attitude of about 1.0°.
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.
Alternately, pressing and holding the collective TRIM beeper left or right will initiate a
standard rate turn (roll axis) at airspeeds above 50 knots.
Note: It is possible to have turn coordination below 50 knots. To achieve this, the
aircraft must go into a coordinated turn above 50 knots and then decelerate to below 50
knots. If the aircraft is placed into a turn below 50 knots it will skid and not achieve turn
coordination as a result of heading hold.

Flight Path Stabilization (YAW)

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. The WOW switch disengages the heading hold when the helicopter is on the
ground. (with FPS ON)
Below 50 knots, the 4-way collective TRIM beeper may be used to slew the reference (FPS)
heading to the desired heading.
If a malfunction is detected, the FPS FAIL caution light on the EICAS will illuminate. The 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 FPS off or evaluate performance to determine the degree and type of
degradation and continue flight with the remaining FPS features.

Coupled Flight Director - FD

The UH60M allows the pilot the ability to utilize the aircraft autopilot system.
The coupled flight director set consists of two independent Flight Control Computers (FCCs) and
Flight Director/Display Control Panels (FD/DCPs).

28
When either FCC is operating in the coupled mode, it provides commands to Pitch, Roll, Yaw,
and Collective by way of the Automatic Flight Control System (AFCS) to drive trim servos.
The flight director also provides command cues, for the engaged mode, on the pilot and copilot
Multifunction Display (MFD) Primary Flight Displays (PFDs) when operating in the coupled or
decoupled mode.
The Flight Director/Display Control Panel FD/DCPs provide independent control of both flight
director and MFD display control selections and settings. The control knobs have momentary
switch action for use in push-to- sync functions. The FD/DCPs communicate with the FCCs and
MFDs.
The flight director processes signals from navigation and flight instruments to produce steering
commands that are performed through the pitch, roll, yaw or collective axis. The steering
commands allow the flight path stabilization (FPS) to:
• maintain a desired heading
• acquire a selected course
• acquire and track localizer and glide slope for ILS approach
• maintain fixed altitude and/or airspeed
• make a programmed acceleration/deceleration and climb/descent to a predetermined
airspeed and altitude
• acquire and track an FMS route
The flight director also provides radar altitude hold mode which consists of; Hover Position,
Hover Velocity, and Hover Deceleration modes

29
COUPLED FLIGHT DIRECTOR MODES
IAS Airspeed Hold 50 - 150 KIAS
HDG Heading Hold Above 50 KIAS
ALT Barometric Altitude Hold Above 50 KIAS
RALT Radar Altitude Hold 0 - 1500 ft AGL
0 - 2000 fpm. Above 50
VS Vertical Speed Hold
KIAS
GA Go Around Goes to 70 KIAS, 750 fpm
climb, engaged at any
airspeed / hvr
Can be armed at any time in
ALTP Altitude Pre-select
flight or on the ground

NAVIGATION MODES

FMS FMS Flight Plan Navigation Roll command to follow

FMS flight plan. Can be
coupled above 50 KIAS.
VNAV FMS Vertical Navigation NOT IMPLEMENTED
VOR VOR Navigation Can be coupled above 50
KIAS.
LOC ILS Localizer Roll command. Can be
coupled above 50 KIAS.
GS ILS Glideslope Collective command. Can
be coupled above 50 KIAS.
Pitch command to slow to
70 KIAS by 200 ft AGL
DECL ILS Deceleration
using the decal mode-0.5
kts. per sec deceleration
HOVER MODES

From 60 KIAS or less at 1.5

HVR DECL Decelerate to Hover – 3.5 kts/sec selectable by
pilot
Maintains GPS position over
HVR POS Hover Position Hold
the ground
Holds groundspeed up to 60
HVR VEL Velocity Hold Mode
kts fwd, 20 aft, 20 lateral

30
 STABILATOR
The stabilator is a variable angle of incidence airfoil that enhances the handling qualities and
longitudinal control of the aircraft. The automatic mode of operation positions the stabilator to
the best angle of attack for the existing flight conditions.

Unlike the UH60-A/L there are no stabilator amplifiers in the UH60-M. Instead the UH60M has
Flight Control Computers to process signals from various sources to adjust the stabilator to the
proper flight position. The functions of the stabilator remain the same as in the UH60A/L with
the primary concern being longitudinal control of the aircraft. To achieve this control the
stabilator is programmed to its optimum angle to provide the following functions:

Streamline or Align stabilator to main rotor downwash in low speed flight to minimize nose-up
attitude resulting from downwash. (30 KIAS and below to minimize nose-up attitudes from the
main rotor downwash on the stabilator)

Collective coupling to minimize pitch attitude excursions due to collective inputs from the pilot.
Collective position sensors detect pilot collective displacement and program the stabilator a
corresponding amount to counteract the pitch changes. This coupling begins at 30 KIAS.

Angle of incidence decrease with increased airspeed to improve static stability. (programs up
with increased airspeed)

Lateral Sideslip to Pitch Coupling to reduce susceptibility to gusts. This feature compensates
for downwash on the stabilator and tail rotor efficiency. In forward flight, the downwash on the
retreating side is weaker than the downwash on the advancing side. If a right sideslip is entered
(left pedal applied), the stabilator encounters increased downwash and the nose tends to pitch up,
therefore the stabilator programs down to prevent the nose up attitude. A right sideslip condition
also results in increased induced flow through the tail rotor and a corresponding decrease in the
amount of lift provided by the tail rotor. In a left sideslip (right pedal applied), the stabilator is
positioned in a reduced downwash area, causing the nose tends to pitch down. The stabilator
programs trailing edge up to prevent the nose from pitching down. A left sideslip condition also
results in decreased induced flow through the tail rotor and a corresponding increase in the
amount of lift provided by the tail rotor.

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 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 reaction.

See single source document for input signals.

31
 Flight Mission Display System
The stabilator angular position is displayed in the upper left corner on the MFD primary flight
display (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).
The stabilator symbology is:
• White when the stabilator is operating in AUTO mode
• Yellow when operating in MAN mode
• Red when failed STB (stabilator) fail flag is displayed

Stabilator center section

Stabilator Lock pins

The stabilator lock pins cable contains switches that close when the stabilator is locked in the
flight position. Closure of the switches completes the automatic power on reset signal path that
initiates automatic mode. When the stabilator is not in the locked condition a STAB
UNLOCKED caution will be displayed on the Engine Instrument Caution Advisory System
(EICAS) display.

32
COMPONENT FUNCTION BRAINS MUSCLE CONTROL AXIX REMARKS
AUTHORITY
Stabilator Helps maintain 2 Flight 2 Stabilator N/A N/A Programmed for 5 basic functions.
level flight Control Actuators
attitudes. Computers
Enhances
handling qualities. (FCC 1 & 2)
Improves CG.

SAS 1 Provides short 2 Flight P/R/Y SAS 5% each for a total Pitch SAS feedback is eliminated by the pitch boost
term P/R/Y Control Actuators of 10% servo. The Pitch Boost is on when the
correction and Computers Roll SAS/BOOST switch is turned on. SAS will not
SAS 2 dampening. Yaw move the flight controls.
Dynamic stability. (FCC 1 & 2)

Trim Force gradient Monitored by Pitch, Roll, 100% when coupled Pitch Pitch Trim is Hydro-Electro-Mechanical. Roll,
the No.2 FCC Yaw, with FPS Roll Yaw, Collective trim servos are
Collective, Yaw Electromechanical.
Trim Coll.
Actuators

FPS Provides long FCC 1 & 2 Trim 100% when coupled Pitch Booth FCCs, through FCC 2, provide FPS
term P/R/Y Actuators with Trim Roll command signals to the pitch, roll, and yaw trim
correction and Yaw actuators which in turn reposition the flight
dampening. (Static controls. For FPS to work you must have SAS 1
stability) or 2, Trim, and Boost on.

Below 50 KIAS Above

Att. Hold Pitch Att. Hold

Att. Hold Roll Att. Hold
Hdg. Hold Yaw Heading hold/Tn. Coordination

33
 PNEUMATIC SYSTEM
The pneumatic system of the UH-60 operates from bleed air furnished by the main engines,
APU, or an external air source. The pneumatic system consists of the engine start system, the
anti-ice system, the heating system, and the extended range fuel system.

Bleed air supplied by a main engine, the APU, or an external air source can be used for starting
purposes. The normal means of starting is through use of the APU as an air source. The APU
can provide sufficient air pressure and volume to accomplish single engine starts throughout a
wide range of ambient conditions and to accomplish a dual engine start when conditions permit.
Chapter 5 of the operator’s manual lists single and dual engine start envelopes for given ambient
conditions.

Engine starts may also be accomplished by using an operating engine as an air

source, and are known as cross bleed starts. Cross bleed starts can be accomplished providing the
Anti-ice light on the source engine is off, NG for the source engine is above 90%, and RPMR is
100%. If an engine is to be used as an air source, the Air Source Heat Start switch must be
placed in the engine position.

The heater system is pressurized by bleed air from any of the sources listed above. The heater
system mixes ambient air with heated bleed air to provide warm air to the crew stations and to
help prevent the overhead windows, and gunner’s windows from fogging. The heater will
automatically disengage when the starter is engaged.

The engine anti-ice systems consist of the engine anti-ice and engine inlet anti-ice systems. Both
anti-ice systems utilize bleed air from their respective engine only. The APU and external
sources are not used for the engine anti-ice systems. The engine anti-ice system utilizes bleed air
extracted from the compressor of the engine. This bleed air prevents ice formation on the vanes
in the inlet. Anti-ice functions are controlled by the pilot through a switch labeled #1 Eng Anti-
ice or #2 Eng Anti-ice. Placing the switch on results in the anti-ice start bleed valve remaining
open to direct heated air to inlet vanes. An advisory light labeled #1 Eng Anti-ice On or #2 Eng
Antiice On will illuminate when the anti-ice start bleed valve opens.
These lights will also illuminate during low power conditions (88-92% NG), because the anti-ice
start bleed valves open to dump excess compressor discharge pressure preventing the possibility
of an engine flameout.

The engine inlet anti-ice is controlled by the same switch listed above, but utilizes air from a
separate bleed air tap. Bleed air is routed directly to the engine inlet anti-ice modulating valve.
When the engine anti-ice switch is placed on, the anti-ice modulating valve samples ambient
temperature through the insulated ambient sense port. If the ambient temperature is above 13º C
(Celsius), the modulating valves may not open. If the ambient temperature is between 4º and 13º
C, the modulating valves may open. If the ambient temperature is below 4º C, the modulating
valves must open. When the modulating valve opens, heated bleed air is forced through the
airframe engine inlet and out the slits on the inlet. When the inlet temperature reaches 93º C, an
advisory light labeled #1 Eng Inlet Anti-ice On or #2 Eng Inlet Anti-ice On will illuminate.
The illumination of the #1 and #2 Eng Inlet Anti-ice On light indicates proper system operation.

34
NP1/NP2 Starter Hot Wx 4. Liquid to liquid cooler
5. ODV
12 Second Trans 105-107% Temp.°C Attempts _________________________________________
Transient 101-105% <15 1-60 sec-1/3 min cool/1-60 sec-1/30 min
Airspeed Limits
Continuous 95-101% >15 -52 1-60 sec-1/30 min cool
193 Vne
Transient 91-95% *if<1hr. since shutdwn, starter 60 sec.
180 Max Landing Light Ext.
Avoid 20-40% and 60-90%
Slope Landing Max Search Light Ext.
NR Nose Up/Right/Left 15° 170 1 Hyd Pump inop
Power On: Minus 2° for every 5 knots of wind 1 SAS inop
Transient 101-107% Nose Down 6° 15k Max Tail Wind Max Gunner Win. Open
Continuous 95-101% 150 Max Auto <16825 GW
Transient 91-95%
Landing Gear 2 Hyd Pump inop (VMC)
GW Level Slope
Power Off: 2 SAS inop (IMC)
<16,825 540 fpm 360 fpm
Maximum 110% 145 Max Cabin Doors Open w/no sound
Transient 105-110% >16,825 300 fpm 180 fpm
proof or secured proper
Normal 90-105% Bank Angle 140 2 Hyd Pump inop (IMC)
NG PRI SERVO FAIL Caution 30° 2 SAS inop (IMC)
12 Second Trans 105-106% Windshield Anti-Ice Sling Load Wt up to 8,000
Continuous 0-105% 130 Max One Engine Inop.
Do not test FAT >27°C
Max Auto >16,825 GW
Q High Speed Yaw 120 Auto w/ Stab Failure
DUAL ENG Avoid full pedal inputs >80 knots W/>8,000 lbs Sling Load
< 80 > _ 110 Max glide dist Clean Config.
APU Operating Hot Wx
Cont 0-120% 0-100% 100 Search Light Ext By
Temp/Cond Time
10 sec. 120-144% 100-144% Max glide dist High Drag Con.
≥43°C/Eng&Rotor Oper 30 min
SINGLE ENG Max A/S Cabin Doors open w/sound
≤51°C/No Eng&Rotor continuous
Cont 0-135% proofing not secured
10 sec. 135-144% Prohibited Maneuvers 80 Recommended Autorotation
Hover turns >30°/second 60 Max G.S. for Roll-On landing
Engine P Pitch >± 30°
5 Minute Limit 100-120 psi 45 Max Hover Side/Rear
Roll >±60° Max Wind Vel. Start/Stop
Continuous 26-100 psi
2 PCL’s idle/off in flight 40 < Stab full down, manual
Idle 22-26 psi
Rearward ground taxi > Stab 0°, manual
Engine T Eng against gust if rotor brake equip 35 Side/Rear ERFS/Sling
30 Minute Limit 135-150 Stab starts program
30 Minute Limits 30-50
Continuous -20-135
TGT 793-846°C Lights, Switches, & MISC
TGT Eng Oil Temp. 135-150°C #1 and #2 FUEL LOW≤ 172 pounds
12 Second Tran. 903-949 Landing/Search Light Ext. #1 and #2 FUEL PRESS≤ 9 psi
2 ½ Minute Cont. 879-903 Fuel filter bypass pop = 8-10 psid
Light Extend by Don’t exceed
Start Abort 851 Fuel Boost pump = 25 psi
Landing 130 knots 180 knots #1 and #2 PRI SVO PRESS ≤ 2000 psi
10 Minute Limit 846-879
Search 100 knots 180 knots BOOST SERVO OFF ≤ 2000 psi
30 Minute Limit 793-846
Normal 0-793 Hyd/SAS inop A/S Limits #1 and #2 RSVR LOW ≤ 60% cap.
1 Hyd/SAS inop 170 knots #1/ #2 Hyd return bypass = 100±10 psid
Transmission P 2 Hyd/SAS inop 150 knots #1/ #2 Hyd Impd byp. = 70±10 psid
Precautionary 65-130 psi Hyd. System pres. = 3000 psi
“ “ IMC 140 knots
Steady state 45-60 psi Pilot Assist pres reducer btn pops = 1375 psi
Normal 30-65 psi Utility mod. velocity fuse > 1.5 gpm
Idle 20-30 psi
Rotor Brake XMSN OIL PRESS < 14±2 psi
Do not apply w/ Eng/NR excpt Emer #1 and #2 ENG INLET ANTI-ICE 93°C
>10 degree up 30-35 psi
Max NR for emer use 76% Hyd pump temp labels > 132°C (270F)
>15 degree up 10-15 psi
Nose up att. (+6°) may flux up to 30 min. Routine stop NR<40%, 150-180psi #1/#2 ENG OUT light/audio <55% NG
not less than 12 sec. LOW ROTOR RPM audio/lt <96%NR
Transmission T Min press for start 450, 690 Max Lose Main GENs on gnd 93-95% NR
Precautionary 105-140 deg Single/Dual Eng IDLE only, no time limit Lose Main GENs in air 85-89% NRPMR
Continuous -20-105 deg Gravity Refuel: 360gal
Autorotation Airspeed Pressure Refuel: 359gal, 300gpm, 55psi
Backup Pump Hot Wx Weight Airspeed Closed Circuit: 356gal, 110gpm, 15psi
(Rotor Static) <16,825 150 knots Eng Anti-ice on: 18% Tq +100lbs/hr
Temp°C Oper. Cooldown >16,825 130 knots Cockpit Heater on: 4% Tq +20 lbs/hr
-54-32 Unlim. N/A Fuel Flow in the Engine Engine Idle fuel @ sea lvl 322lbs/hr
33-38 24 min. 72 min. 1. Pump (engine driven) Engine Idle fuel @ 5000ft 277lbs/hr
39-52 16 min. 48 min. 2. Fuel filter Engine Fly fuel @ sea lvl 552lbs/hr
____________________________________ 3. HMU Engine Fly fuel @ 4000ft 479lbs/hr
35
APU fuel flow @ sea lvl 120lbs/hr 1. Pitch,Roll, and Yaw signals from resp. EGIs 3. Fault Indication. DEC code sent to
APU fuel flow @ 4000ft 101lbs/hr (Nose) corresponding ENG Q indicator and the FMS
2. Airspeed – Air Data Computer DEC status page.
Mechanical Mixing Malf indicated by erratic movmnt, Reset CPTR1/
Coll. to Pitch – Compensates for rotor downwash 4. Load Sharing/Torque Matching. Increase Q
2, or turn off that SAS of low side to match up to 3% NP above
on the stabilator.
Coll. to Roll – Translating Tendency.
_____________________________ reference.
Coll. to Yaw – Torque Effect. Stabilator 5. Transient Droop Compensation. 4-1
Yaw to Pitch – Lift component of Tail Rotor. Variable angle of incidence airfoil that enhances improvement to compensate for transient
the handling qualities in the pitch axis in forward rotor droop. Uses coll. Pos. transducer
flight. signal.
Electronic Coupling SCALP
Collective to Airspeed to Yaw – Compensates for 1. Streamlines w/ rotor downwash at low speed 6. History Counter. Sends signals to History
torque effect in addition to Collective to Yaw (below 30 kts) to min. nose up attitude from Counter.
mixing based on collective position and airspeed. downwash. (ADC, FCC) 7. Hot Start Preventer. Tells ODV to shut off fuel
FCC#2 commands yaw trim actuator 100% below 2. Collective Coupling to min. pitch attitude when 900°C is reached, Ng<60, Np<50.
40 kts decreases until100kts none above 100 kts. excursions due to collective inputs.
8. Overspeed Protection. Triggered when Np
(Collective Pos. Transducers) begins@30kts
AFCS reaches 120±1%.
3. Angle of Incidence decreases above 30 kts. to
Enhances static stability and handling qualities of 9. TGT Limiting. Limits fuel flow
improve static stability. (ADC, FCC)
the helicopter; autopilot when reaching appx. 879 Dual Eng,
4. Lateral Sideslip to pitch coupling to reduce
Five subsystems: 903 (2½ min) when opp Q is < 50%.
susceptibility to gusts. Stab up for right
1. SAS pedal, stab down for left pedal. (Lat. Engine Alternator
2. Trim accelerometers) Three windings for three functions:
3. FPS 5. Pitch Rate Feedback to improve dynamic 1. Ng signal to cockpit tachometer. (green)
4. Stabilator stability. Dampen pitch excursions due to 2. Ignition power to the exciters during start.
5. Coupled Flight Director turbulence or “G” forces. (EGI, FCC) (yellow)
ODV HOPS 3. DEC power. (yellow)
FPS Loss of alt. does not result in high side failure
1. Hot Start Prevention shuts fuel flow at
Basic autopilot – Enhances static
900°C. Weight-On Wheels Switch
stability through long term rate 2. Over-speed protection (NP) from DEC at CHAFF AIM TUBE
dampening in pitch, roll, and yaw. It 120+1% 1. CVR/FDR: IBIT, erase on gnd
is the “Brain”. When coupled with 3. Purges main fuel manifold and allows back 2. Hydraulic leak test disabled in fight.
Trim FPS has 100% control author. flow of high-pressure air for purging. 3. Audio for low NR disabled on the ground.
Below 50 Kts Above 50 Kts 4. Sends fuel through manifold to 4. Flight Dir. Modes enabled in-flight
Att. Hold Pitch Att. Hold injectors for starting & eng 5. FMS-Model acft, EGI GC or GPS alignment
Att. Hold Roll Att. Hold IBIT on ground/ EGI air alignment in flight
Hdg Hold Yaw Head Hold/Tn coor
HMU
Basic fuel control; incl. high press. fuel pump & 6. AFCS Pitch, Roll Att hold/ Hdg hold/
variable geometry servo-actuator. Maneuvering Stab/ Rad Alt hold/ Airspeed
These systems must be on and PM CAN VDTO hold – Enabled in-flight.
operational for FPS to func. at 100%: 7. IFF Mode 4-Auto Zeroize enabled in flight.
1. SAS 1 and/or SAS 2 Pumps fuel at high pressure 8. MFD status page activation on grnd
2. SAS/Boost 1. Meters fuel to ODV in response to PAS, 9. Thermal protection for B/U pump enabled on
3. Trim LDS, trq. motor from DEC, & eng. variables. the ground.
4. Stabilator (helps, but not req.) 2. Collective Pitch Compensation through the 10. Underfreq. protection disabled in flt. (right
LDS. When the collective is moved, Ng is switch)
Trim 11. B/U Pump Auto OPN regardless of switch
Provides a gradient and detent holding force for reset for immediate Np response.
3. Accel/Decel flow limiting preventing comp. position enabled in flight. Disabled on gnd
pitch, roll, collective, and yaw. Consists of 3 unless Acc Low
electromechanical actuators (collective, roll and stalls, eng. damage, or flame out.
4. Ng limiting – Limits max torque available 12. ESSS jettison disabled on ground. (right
yaw), & 1 electrohydromechanical actuator. (pitch) switch)
SAS/Boost required for operation. under low temp conditions.
Slip clutches req. 80 lbs max in yaw, 13 lbs max 5. Variable Geometry Positioning of the inlet
guide vanes for optimum performance. Equip avail. w/ batt pwr only: CVR/FDR, P
roll, 22 lbs max collective during actuator jam.
6. DEC lockout PAS Override & Control with inboard MFD, P FMS, P FDDCP, #2 ADC, #2
FCC monitors continuously, will turn off driving
ECU Inoperative. Allows pilot to DCU, #2 EGI, #1 stab act (manual), ICS, clock, #1
trim actuator.
mechanically bypass torque motor inputs. FM, VHF, VOR/ILS, IFF, Radar Alt., Searchlight,
AVCS 7. Torque Motor to Trim Ng Output. To fine Stdby Comp. lt, cockpit util lts, APU sys cntl, fire
Reduces cockpit and cabin vibrations by tune engine output. Can be overridden in det/ext, audible warn, SAS ctl, fuel txfr, jettison,
mechanically generating additional vibrations that DEC lockout. backup pump ctl, hook rel.
are out-of-phase with the main rotor (90-105 NR). 8. Opens Vapor Vent for manual HMU priming,
to remove air from HMU.
Coupled Flight Dir Status Modes
Uses 3 Force Generators
DEC
SAS1&2 Controls electrical functions of engines and
Collective Roll Armed:
LOC
Provides short term rate dampening in pitch, roll, transmits operation info to cockpit. Capture: NAV (VOR)
and yaw axes. Enhances turn coord >50 kts. IN FLT HHOT ALT
Hover augment <50 kts. 5% cont. auth. Ea. 1. Isochronous NP Governing. DEC will maintain ALTP Roll Capture:
Controlled by respective FCC comp. Np reference GS BC
Input: 2. NP Reference from the cockpit, (INCR/DECR RALT HDG
from 96-100%) GA LNAV (FMS)
VS LOC
Pitch Capture:
36
NAV (VOR)
DCL
IAS
POSN (Hover)
VHLD (Hover)
 UH-60M/HH-60M JACS Antenna Layout
12-190-160
VHF/UHF LOS GPS #2
COMM 2
VOR / LOC
GPS #1

VU10-826
VHF/UHF LOS AV 405-1
12-190-310 AV 2091-1 Troop Commander VHF-FM #1
VHF/UHF LOS UHF-SATCOM COMM 1
& UPPER IFF COMM 3
COMM 3 FM10-267
S65-8282-301 VHF-FM #2
RADAR L-SATCOM COMM 4
WARNING BFT
(BOTH SIDES)

GPS #1
GPS #2 CMWS EOMS
AFT SENSORS
(BOTH SIDES)

GLIDESCOPE
RADAR WARNING
(BOTH SIDES)

CMWS EOMS
FORWARD SENSORS
(BOTH SIDES) AVR-2B KEY
LDS SENSOR
PLS (MEDEVAC Only)
TACAN VOR / LOC LEGACY
JACS ANTENNA

NEW JACS
ANTENNA
RADAR
ALTIMETER
12-190-160
LF / ADF LOWER IFF VHF/UHF LOS
COMM 2
MARKER
RADAR WARNING BEACON STORMSCOPE
BLADE ANTENNA

37