Skycrane Operation Manual
Transcript of Skycrane Operation Manual
MILVIZ SIKORSKY S-64 SKYCRANE - OPERATION MANUAL Page 2
TABLE OF CONTENTS
General Information Page 3
Landing Gear System Page 3
Fire Warning System Page 4
Engines Page 5
Fuel System Page 11
Hydraulic and Pneumatic Systems Page 14
Power Train System Page 15
Main and tail Rotors Page 17
Pitot Heat Page 18
Electrical Power Supply and Distribution Systems
AC Supply System Page 19
DC Supply System Page 22
Auxiliary Power Plant Page 23
Panel Layouts
Overview Page 25
Main Panel Page 26
Middle Console Page 28
Overhead Panel Page 29
Engine Control Unit Page 30
Aft Pilot's Compartment Page 31
Normal Procedures Page 32
MILVIZ SIKORSKY S-64 SKYCRANE - OPERATION MANUAL Page 3
GENERAL
The CH-54 helicopter, is a twin-turbine, all metal, flying crane. It has a maximum gross weight of
47,000 pounds. The helicopter is designed to carry detachable pods for transporting personnel and
cargo, to carry hoist loads, hard-point attached loads, and/or to tow surface type vehicles.
The pilot’s compartment seats five air crewmen, the pilot and copilot, aft pilot, flight engineer, and
one additional crewman. The flight engineer and crewman are seated in removable seats installed
in the well between the pilot and copilot, and a small jump type seat behind the pilot. The aft pilot is
seated behind the copilot and faces aft and uses a remote stick (cyclic and yaw), a collective pitch
control, and has cargo handling instruments. The pilot and copilot have standard instruments and
controls. Pilot’s compartment windows provide visibility in all directions. The pilot and copilot have
entrance doors on the left and right sides of the helicopter and an all door is on the right rear side of
the pilot's compartment. Armor plating may be installed in vital areas such as night controls
hydraulic lines and controls. etc.
The helicopter has a fixed tricycle landing gear, a full swiveling nose wheel and dual main wheels
supported by struts. Only the main wheels are equipped with brakes. The helicopter can carry
single-point loads from a cargo hook at its center of gravity. There are 32 structural hard points
from which loads may be slung. Four structural hard points which symmetrically bracket the center
of gravity are attachment points for the load leveler system. The load leveler system uses four load
suspension units that are raised and lowered by hydraulic cylinders to level the load. Each of the
four suspension units has a cargo lashing reel for attaching the loads The single-point hoist is a
hydraulically powered winch with a cargo hook, below the center of gravity of the helicopter.
The helicopter is powered by two axial flow gas turbine engines mounted side-by-side on top of
the fuselage above and aft of the pilot’s compartment. Engine torque is transmitted through a
system of gear boxes and drive shafts to the main and tail rotors. The main rotor consists of a fully
articulated main rotor head and six main rotor blades. The tail rotor has a tail rotor head and four
tail rotor blades. An electrically-actuated retractable tail skid on the lower pylon protects the tail
rotor blades from striking the ground if the nose is too high.
Electrical power is supplied by a 115/200 volts alternating current (vac) system, and is rectified
to provide a 28 volts direct current (vdc) system. The auxiliary power plant (APP), aft of the main
transmission, enables ground starting of the engines and ground operation of the hydraulic and
electrical systems. Antennas for radio communication, navigation and IFF are on the exterior of the
helicopter.
LANDING GEAR SYSTEM
The listed tricycle gear consists of two fixed main landing gear assemblies and a full-swiveling
nose wheel. The main landing gear wheels have power boost hydraulic brakes. The two fixed main
landing gears are aft of the center of gravity and the main gear box. The wide tread gives excellent
lateral stability for landing ground handling and cargo loading. The main landing gear wheels and
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the nose wheel have tubeless tires. The nose wheel is self-centering full-swiveling and has a
shimmy damper. Each main landing gear support has four steps for climbing to the top of the
fuselage. On helicopters with cold weather kit, installed the size of the steps have been increased by
adding step extensions. The added step extension will permit crewmembers wearing arctic vapor
barrier boots to safely climb to the top of the fuselage. The landing gear system is protected against
hard impact landings by metering pins within the oleos which aid in energy absorption.
A tail skid is used to protect the tail rotor blades if a nose high landing occurs. A retractable tail
skid is on the rear fuselage beneath the pylori. The extension and retraction of the tail skid is
controlled through a switch, marked TAIL SKID UP and DOWN, on the center console. The tail skid
should be extended for landing but can be retracted when loading and unloading cargo, The tail
skid is operated by 28 vdc power from the primary bus. A light on the advisory panel, marked TAIL
SKID UP, goes on whenever the tail skid is retracted.
The main landing gear wheels have hydraulic boost power brakes connected to the utility
hydraulic system and operated by toe pedals on the tail rotor control pedals. A parking brake
handle operates a hydraulic valve to lock the wheel brakes. When the parking brake is on a green
light on the advisory panel, marked PARKING BRAKE ON, goes on whenever dc primary electrical
power is applied. The parking brake handle marked PARKING BRAKE is on the aft left-side of the
center console. The parking brake handle is applied by first depressing the toe brake pedals and
then pulling the parking brake handle out and releasing toe brakes. Pressing the brake pedals will
release the parking brake, causing the parking brake handle to return to OFF. Boost hydraulic
pressure for the wheel brakes will not be available because of an interrupted circuit from the
landing gear scissor switch to the hoist isolation valve. This is noted by the ISOLATION VALVE
OPEN advisory light on the caution/advisory panel being on when the weight of the helicopter is on
the landing gear.
ENGINE FIRE WARNING SYSTREM
The engine tire warning system consists of two fire detector systems and a fire surveillance
facility. Two fire detector systems, one for each engine, warn the pilot of an engine fire. Two
continuous element fire detector cables for each engine, are adjacent to the engines. They are wired
into a closed series loop connected to a control unit that lights warning lights in the pilot’s
compartment in case of a fire. The fire detector systems operate on 28 vdc from the primary bus.
The engine fire warning system consists of two red master FIRE warning lights on the instrument
panel, two red engine FIRE warning lights, installed on the engine control quadrant, and a test
switch, marked SHORT and TEST, on the FIRE WARNING TEST panel on the overhead control panel.
Both master FIRE warning lights on the instrument panel and one of the engine FIRE warning lights
on the overhead quadrant will go on in case of a fire in the corresponding engine area. The test
switch is spring-loaded to the center position. For normal operation of the system, when the switch
is placed to SHORT, the FIRE warning lights should not go on. When the switch is placed to TEST,
the circuit continuity is checked and all four FIRE warning lights should go on.
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The system operates on 28 DC from the DC primary bus through two circuit breakers marked
FIRE DET and two marked FIRE DET CONT on the dc circuit breaker panel.
A fire surveillance facility warns the pilot and copilot of a possible engine or auxiliary power plant
(APP). The facility consists of five flame detectors in each engine area and two at the APP unit, three
control amplifiers in the attic behind the cockpit and two warning lights on each side of the
overhead quadrant. In addition to the pilot’s and copilot’s master FIRE warning lights the engine
fire warning lights and the FIRE WARNING TEST control panel an APP control panel FIRE warning
light on the overhead control panel are installed to complete the facility. The facility operates on 28
DC from the dc primary.
The flame detectors are photoconductive cells that react to hydrocarbon radiation of a fire.
Flames in the area of a detector will create a signal to a control amplifier which in turn provides at
dc signal to light the respective fire warning lights. The facility resets itself when the source of
radiation ceases. Flames in the area of a detector will create at signal to a control amplifier which in
turn provides a dc signal to light the respective warning lights. The facility resets itself when the
source of radiation ceases. The detectors are shielded with aluminum and are located in areas that
prevent the filtering of sunlight from activating them and causing the fire warning lights to go on.
An infrared surveillance rotary test switch on the FIRE WARNING TEST panel when activated,
sends a test signal throughout the facility to light the warning devices. When a fire is detected in
either engine area or the APP area the two master FIRE warning lights, marked FIRE on the
instrument panel, go on. If the fire is in the No. 1 or No. 2 engine area the respective light on the
overhead quadrant will also go on. If the fire is in the APP area the FIRE warning light on the APP
control panel will also go on.
ENGINES
The helicopter is powered by two axial flow gas turbine engines model T73-P-700. The free
turbine engine consists of a gas generator and an independent constant speed two-stage free
turbine mounted at the rear of the gas generator. Major sections of the engine are the compressor,
combustion turbine, free turbine and accessory gear box. The gas generator has a nine-stage single-
rotor compressor driven by a two-stage turbine. The combustion chamber of the gas generator
contains eight annular burner cans into which fuel is sprayed through nozzles mounted at the inlet
of each can. A fuel drain valve at the bottom of the combustion chamber case automatically drains
the chamber alter engine operation.
Two high energy ignition units and igniter plugs are used to start combustion. The engine has
three separate air-bleed systems and external thigh pressure internal (cooling) and interstage
(overboard) system. The external system is used to anti-ice the engine inlet case bellmouth and for
fuel de-icing or to operate EAPS (engine air particle separator) blowers when installed. The internal
system is used to pressurize the main bearing seals and to cool hot sections. The interstage system
countersets engine compressor instability. Two separate anti-icing air systems prevent icing of the
compressor inlet surfaces and bellmouth.
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The gas turbine section exhaust gases drive a two-stage free turbine assembly. A component drive
gear box is on the compressor section at the bottom of the engine. Power is supplied to the gear box
from the compressor rotor shaft. The fuel control and engine driven fuel pump are mounted on the
right front and the hydraulic starter on the left front of the engine. The fuel pump the hydraulic
starter and the engine tachometer-generator are gear driven. The main oil pump assembly is in the
component drive gear box.
Most of the engine cooling air system’s cooling air passes through the inside of the engine. The
hottest area is at the entrance to the first stage of the free turbine. Internal air is bled from the rear
of the compressor section, passes through the inside of the diffuser section to the combustion
section then through the duct to the free turbine.
Two engine air particle separator assemblies may be installed to protect the engines from erosion
by airborne particles. Neither the separator nor the bellmouth is anti-iced when the separator is
installed. With the separator installed the blower is controlled by the same switch used for the
engine ANTI-ICE system.
The helicopter is equipped with engine anti-icing system if EAPS not installed. The engine
bellmouth and the inlet guide vanes are susceptible to icing, Air-inlet icing may occur even at
temperatures well above freezing when there is no evidence of other icing on the helicopter. Once
inlet ice begins to form accumulation can build up with starting rapidity at the engine air-inlet
results in a loss of power which may be noted by a drop in engine pressure ratio if the condition is
severe enough. One of the first indication of ice may very likely be at compressor stall which could
he slight. or very pronounced (surging). Ice at the inlet reduces airflow and internal pressure.
Since fuel flow is controlled by a pressure signal from the engine burner section, fuel flow will be
reduced as icing at the inlet causes burner pressure to fall off. The reduced burner pressure and fuel
flow will allow a buildup of engine-inlet ice without the T5 increasing significantly, until the ice
accumulates. Ice may begin to break off before at rise in T5 is observed. A rise in T5 is often
associated with inlet-icing and frequently results from ice passing through the engine, not from ice
buildup at the inlet. Therefore, T5 is not considered a reliable indication that ice is forming at the
air-inlet. A large buildup of ice on the guide vanes can result in compressor damage or possibly an
engine flameout.
The engine anti-icing system, which uses hot compressor bleed air to warm the compressor inlet
guide vanes, is not a de-icing device to avoid the possibility of ice breaking off and passing through
the engine, the anti-icing system should be turned on when icing conditions are suspected, to
preheat the parts of the engine air-inlet that are susceptible to icing before it is actually
encountered. The power loss will be slight when compared with that which might occur once inlet
ice starts to form.
Anti-icing heat should be used when required for all ground operation, during take-off and in
flight. At the relatively high power settings used during climb or cruise the anti-icing system will
supply excess heat for protection against the accumulation of ice in the inlet section of the engine.
An anti-icing valve regulates the flow of anti-icing air to the bellmouth automatically, with changing
compressor discharge temperature. The anti-icing valve reduces the flow of anti-icing air with
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increasing compressor inlet temperature. The air for inlet guide vane anti-icing is bled from the
compressor section and is carried forward to the inlet case by an external tube location on the left
side of the engine. Air How through this tube is controlled by an electrically-actuated valve.
Two switches marked EAPS/ANTI-ICE No. I and No. 2 are on the center console. The switches
have marked positions ON and OFF. When the switches are placed ON, heated air from the diffuser
section provides anti-icing for the engine air inlet and bellmouth. The valves operate on 28 vdc
power from the dc primary bus. The engine anti-icing system should be activated one engine at a
time due to a slight engine power loss. The anti-ice systems may be activated at any power setting.
When the engine air panicle separator is installed, air for the bellmouth anti-icing system is used to
power the EAPS blowers and inlet guide vane air is shut off by closing the inlet guide vane anti-icing
valves. The valves are spring-loaded to open position and will close when DC power is applied to
the solenoid.
During icing conditions, the anti-icing switches should remain on through all ground and flight
operations. Two caution lights marked #1 ANTI-ICE and #2 ANTI-ICE are on the caution panel.
When electrical power from the 28 vdc primary bus is applied the lights will go on if temperature in
the bellmouth is below +4.4°C. Below this temperature engine icing conditions are possible and the
anti-ice switches should be turned on. When air temperature in the bellmouth increases to above
+4.4°C the caution lights will go of The caution light will go on with or without EAPS installed.
Whenever the anti-ice system is operating, two lights on the advisory panel marked #1 ENG ANTI-
ICE ON and #2 ENG ANTI-ICE ON will go on.
A hydro-mechanical fuel control with a free turbine speed-sensing unit governs the free turbine
rotor speed and schedules fuel flow. A collective bias control feeds collective inputs through a direct
mechanical linkage to the fuel control. The collective bias input compensates for droop when
collective is increased, and overspeeds when collective is lowered. The fuel control unit operates an
airbleed valve (bleed strap ), that bleeds interstage compressor air overboard to aid engine starting
and engine operation at low power.
The control has a fuel metering system and a computing system. The metering system selects the
rate of fuel flow to be supplied to the engine to meet power demands. It is subject, however to
engine operating limitations scheduled by the computing system in accordance with data received
from engine sensors. Incorporated within the main rotor collective pitch control is a mechanical
push-pull system (collective bias) that gives collective bias cam inputs to the fuel control that are
approximately proportional to the rotor load. The system is automatic since variations in collective
pitch stick position are mechanically fed into the fuel control. This, in turn, provides limited control
to prevent free turbine and rotor droop or overspeed.
An electrical protective device senses N2 overspeed at approximately 113 percent N2. It permits
oscillation of the engine between 111 and 113 percent N2 to contain the engine N2 and will
automatically contain the engine so as not to exceed engine. rotor, or transmission limits, the
protective device shuts down the gas generator, the N1 speed control lever must be retarded to
SHUT-OFF to release a lockout in the fuel control before a restart can be made. The N2 speed engine
fuel control system sense cable signals the engine fuel control overspeed protection system and
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automatically shuts off the engine if free turbine speed reaches 115 percent. If the N1 speed sense
cable connecting the free turbine tachometer and the line fuel control breaks it can cause engine
failure or free turbine overspeed. Should the N2 speed sense cable intermittently bind then release
a false signal will be sent to the fuel control and the overspeed latching mechanism of the fuel
control will shut oil' the engine. Complete breakage of the N2 speed sense cable will allow the
engine to accelerate to about twice the SHP of that engine at the time of the failure. Should the
tower shaft fail both overspeed systems are inoperative.
An electrical secondary protective device has been added should the N2 speed sense cable fail. It
permits oscillation of the engine below primary overspeed protective system shutdown speed and
will automatically contain the engine so as not to exceed engine, rotor or gear box limitations.
Two N1 speed control levers one for each engine marked ENG N1 CONT permit scheduling of the
gas generator speed (N1) in the starting and idle range. The levers also provide topping governor
rpm control in the power range and permit the manual resetting of the maximum engine speed
limit. When the levers are moved below IDLE the fuel shutoff valve in the fuel control closes. The
levers are connected by cables to the fuel control. The levers have marked positions SHUT-OFF,
GRD IDLE and FULL OPEN.
The fuel de-icing heater in the fuel inlet line to the fuel control unit uses heat from the external air
bleed system of the engine to eliminate fuel system icing. Two switches, marked HEATER and OFF,
are on the fuel management panel. The HEATER position opens the valve in the external air bleed
line, providing heat to de-ice the fuel before entering the fuel control. The valves operate on 28 vdc
from the primary bus. The fuel low temperature light on the caution advisory panel is marked FUEL
TEMP LOW. Whenever the fuel temperature is +0.6 C° ± 0.6° or below, the light goes on indicating
that the fuel heater should be turned on. The light will go off only when the temperature of the fuel
rises above +0.6 C° ± 0.6°. The light operates on 28 vdc power from the primary bus.
The fuel filter bypass caution lights on the caution advisory panel are marked #1 ENG FUEL
BYPASS and #2 ENG FUEL BYPASS. The lights will remain on when there is an accumulation of solid
contaminant in the fuel tiller. The lights are operated by 28 vdc power from the primary bus. The
fuel de-icing heater is designed to be used intermittently when icing is detected.
Each engine has an independent oil tank pressure system, scavenge system, breather system, and
oil/fuel heat exchanger fuel oil cooler system. It is a completely automatic system requiring no
action by the pilot and serves both lubricating and cooling purposes. Oil is gravity-fed from the tank
to the inlet of the engine driven main oil pump. The pump forces oil through the main oil filter and
check valve. If the main oil filter and check valve become clogged, the filter bypass valve opens. A
pressure relief valve maintains a constant oil pressure regardless of altitude and engine speed. The
main pump discharge normally flows through the oil cooler. If the oil is flowing at a low
temperature, the bypass valve opens and the oil then flows directly to the engine bearings and gear
box. The scavenge pumps return oil from the bearings and gear box to the oil
tank. The oil pressure system on the left side of the component gear-bolt, consists of the main oil
pump main oil filter, tilter bypass valve, pressure relief valve, bypass valve incorporating
temperature regulation and pressure relief, jet filters, and low pressure warning light.
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The ignition system is a high energy, capacitor discharge type system that provides both high
voltage and a hot spark. The system consists of two identical, independent units, one for each spark
igniter at the numbers three and six burner cans, respectively. The electrical source of power is the
28 vdc primary bus. The spark rate is one to two sparks per second. Filters eliminate radio
interference. The igniters operate only during the time the starter is engaged. The ignition system
should never be operated continuously for more than 3 minutes at any one time. The duty cycle for
cooling purposes during engine starting is 2 minutes on, 3 minutes off, 2 minutes on and then 23
minutes off. Do not exceed two start attempts in a 30-minute period of time.
Two ignition switches marked IGN SW NO. l. NO. 2. NORM, OFF and TEST are on the center
console. The switches operate on 28 vdc power from the primary bus.
The main engine starting system consists of hydraulic starters, flow regulator valves, relief valves,
starter valves, tilters, plumbing, and necessary electrical controls and wiring. Hydraulic power for
engine starting is supplied by the utility hydraulic pump on the aft end of the main gear box. The
pump output is set automatically during an engine start cycle and develops a maximum 3500 psi.
The system is protected by a relief valve set at 3500 ± 50. The flow regulator is provided to control
maximum speed of the starter motor during an engine start cycle. Hydraulic fluid from the
hydraulic pump is transmitted to the now regulator through the engine start valve (shutoff), and
then through the hydraulic starters. Return fluid from the hydraulic starters is routed to the inlet
side of the reservoir.
Two starter switches are on the control quadrant behind the appropriate engine N1 lever. The
starter circuit is energized by pushing in on the starter switch on the N1 lever and holding the lever
at SHUT-OFF. When the start switch is depressed, the utility pump control valve energizes, directing
l500 psi hydraulic pressure back to the utility pump, changing the pump to the high pressure phase
(3500 psi) for engine starting. The engine hydraulic pressure switch will energize, holding in the
start circuit. A time delay relay will then open the engine start valve allowing the 3500 psi to how to
the start motor. The starter switch does not have to be continually depressed throughout the
starting cycle and the dropout switch will automatically deactivate the starter circuit and the
ignition circuit as soon as the engine reaches 29 ± 3 percent N1. Both the No. I and No. 2 engine
starting circuits receive power from the No. I and No. 2 DC primary bus respectively. The abort
switch, when activated by pulling the appropriate engine N1 lever out and to SHUT-OFF, will abort
an engine start any time during the starting cycle.
The turbine case and turbine wheel remain at about the same temperature when the engine is
operating. However, the turbine wheel is relatively massive, compared with the ease, and is not so
readily cooled. The turbine case is exposed to cooling air from both within and without the engine.
Consequently, the case and the wheel lose residual heat at different rates after the engine has been
shut down. The case, cooling faster, tends to shrink down upon the wheel which is still rotating.
Under extreme conditions, the turbine blades may rub and seize, hence the requirement for a
cooling period if the engine has been operating at prolonged high speed. Should the turbine wheel
seize, no harm will normally result, provided no attempt is made to turn the engine over until it has
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cooled enough to free the wheel. In spite of this, it is obvious that every effort should he made to
avoid seizure. The fuel boost pump must not be turned OFF before the N1 lever is placed to SHUT-
OFF to be sure that fuel remains in lines and that engine-driven fuel pumps do not lose their prime.
The same situation would apply to the fuel shutoff lever, however this lever should never be used
as a means of shutting off fuel to the engine except in an emergency. Since it is upstream of the fuel
control, shutting of the fuel system shutoff lever before shutting off the fuel at the fuel control might
drain or even collapse the fuel lines, and will cause the engine driven fuel pumps to lose their prime.
Under such conditions, the fuel boost pumps usually are unable to re-prime the engine-driven fuel
pumps without fuel and/or air being bled from the fuel control.
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FUEL SYSTEM
Each engine is supplied fuel from a separate tank and fuel line containing a manually-operated
gate valve. The forward tank supplies fuel to the No. I engine and the aft tank supplies the No. 2
engine. Two main tanks containing two self-sealing cells are installed in the fuselage, one forward
and one aft of the cargo hoist well. The self-sealing fuel tanks are protected by plastic anti-flowering
panels. Two booster pumps (No. 1 and No. 2) are installed in the aft cell of the forward tank and the
forward cell of the aft tank. The pumps draw fuel from both cells of its tank and maintain a constant
pressure to the engine-driven pump.
A fuel crossfeed system between the two separate feed lines from the main tanks allows fuel from
both forward and aft tanks to be directed to one engine, during single-engine operation, or fuel
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from either forward or aft tanks to be directed to both engines if one or the other fuel tank system
fails. Check valves within the tanks prevent fuel transfer between tanks. A crossfeed switch, marked
X-FEED and CLOSE, is on the fuel management panel. The switch actuates a valve that is powered
by 28 vdc from the primary bus to open or close fuel crossfeed valve.
Two in-line filters are on the left side of the fuselage for No. 1 engine and two for No. 2 engine.
The filters prevent foreign particles from entering the engine fuel control. When a filter is clogged,
the pressure differential within the filter unit will open a fuel bypass line. Fuel then bypasses the
filter element to prevent fuel starvation. Caution lights marked #1 AFR FUEL FILTER or #2 AFR
FUEL FILTER will go on when the filter element is being bypassed. A drain valve is incorporated in
each filter assembly and a manual shutoff valve is in the fuel line at each filter inlet. When a filter is
clogged, the pressure differential closes a switch which lights the #1 AFR FUEL FILTER or #2 AFR
FUEL FILTER capsule on the advisory panel. The caution lights operate from the dc primary bus.
Two manually-operated fuel shutoff levers, one for each engine are on the engine control
quadrant on the overhead control panel. When the levers are placed full aft they close the main fuel
shutoff valves. When they are placed full forward they open the valves.
The fuel boost pumps are mounted in the aft cell of the forward tank and the forward cell of the
aft tank. Each pump is provided with fuel inlets at the pump housing and in the opposite cell of the
tank. With either or both inlets submerged each pump will provide maximum fuel flow
requirements for one engine. Check valves in each of the fuel boost pump lines prevent transfer of
fuel from either tank when the crossfeed valve is open or if a fuel boost pump weakens or fails. The
No. 1 fuel boost pumps are powered by 115 vac from the No. 1 primary bus and are protected by ac
circuit breakers marked NO. I FUEL PUMP FWD TANK and AFT TANK. The No. 2 fuel boost NO. 2
FUEL PUMP FWD TANK and AFT TANK. Four fuel boost pump switches on the fuel management
panel control the boost pumps. The switches are in sets of two marked FWD TANK and AFT TANK.
Above each switch is it number 1 or 2 to designate the pump controlled by that switch. Each switch
has two marked positions, PUMP (ON) and OFF. The switches control 28 vdc from the primary bus
and.
Two fuel quantity indicators, marked FWD and AFT, are on the instrument panel and indicate fuel
quantity in pounds of fuel. The indicators are powered by 115 vac from the No. l and No. 2 ac
primary buses and the ground inverter and. The fuel low-level caution lights are on the upper
section of the caution panel. The lights, marked FWD FUEL LOW and AFT FUEL LOW, will go on
when about 500 ± 50 pounds of fuel per tank remain in the tank. The caution lights are powered by
28 vdc power from the primary bus. Four fuel boost pump failure lights, marked 1 PUMP 2. FWD-
TANK and I PUMP 2. AFT-TANK, under the heading FAILURE, are on the fuel management panel.
The warning lights will go on if any one of the respective pumps fails. The warning lights are
powered by 28 vdc power from the primary bus. The fuel pressure caution lights marked #I and #2
FUEL PRESS, go on whenever fuel pressure for that engine is below normal. The lights are powered
by the 28 vdc primary bus and are protected by circuit breakers on the dc circuit breaker panel
marked No. 1 and No. 2 FUEL press.
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An auxiliary fuel tank consisting of a single self-sealing cell is aft of the main tank. Fuel from the
auxiliary fuel tank is transferred to the main tanks by transfer pumps mounted in the sump filler
necks are on each tank. Only one transfer pump should be used at any given time since the AUX
FUEL PRESS caution light will not indicate the failure of an individual transfer pump if both are
operating. During normal operation of the auxiliary fuel system, the transfer pumps will supply fuel
to both main tanks. Two switches, marked NO. l and NO. 2. ON and OFF, under the heading AUX
FUEL PUMPS control the auxiliary tanks fuel pumps. A light on the caution panel, marked AUX FUEL
PRESS, will go on when a transfer pump switch is on and no fuel transfer is taking place. When the
light goes on, either the auxiliary fuel tank is empty or a pump has failed.
The fuel quantity gage marked AUX on the instrument panel indicates fuel quantity in pounds of
fuel. The fuel quantity gage indicator has a fuel quantity selector knob that limits the fuel quantity
entering the auxiliary fuel tank. This permits selective refueling via the pressure refueling system,
to any quantity desired. As the selected fuel quantity is attained, the primary high-level shutoff
solenoid, using a signal supplied from the gaging system, shuts off the pressure refueling system.
The quantity selected will be indicated by the refueling "bug" on the outer scale of the fuel gage. The
gage is powered by 115 vac power from the No. 1 ac primary bus and through the ground inverter.
A fuel quantity gage toggle test switch for the auxiliary tank is on the instrument panel. The AUX
FUEL PRESS caution light on the caution panel will go on only when a fuel transfer pump is ON and
no fuel pressure is present. When the light goes on, fuel transfer has stopped, either because the
auxiliary fuel tank is empty or the transfer pump in use has failed.
The range extension system is installed to increase the range of endurance of the helicopter. The
system does not provide fuel directly to the engines, but functions to replenish fuel of the forward,
aft main, or auxiliary tanks. The range extension system consists of two external fuel tanks, No. 1
and No. 2. attached to the tank support box beams suspended from the main landing gear supports
inboard of the main landing gear, a tank pressurizing system and two control panels. Fuel transfer
is made by placing No. 1 and/or No. 2 transfer switches to TRANSFER.
U.S. GALLONS AND POUNDS
Forward tank 2 cells: Each tank 431 U.S. Gallons or about 2800 pounds.
Aft tank 2 cells: Each tank 431 U.S. Gallons or about 2800 pounds.
Auxiliary tank: 431 U.S. Gallons or about 2800 pounds.
Range extension tanks: 300 U.S. Gallons each tank or about 1950 pounds each. (3900
pounds total.)
MILVIZ SIKORSKY S-64 SKYCRANE - OPERATION MANUAL Page 14
HYDRAULIC AND PNEUMATIC SYSTEMS
Flight control serve hydraulic consist of a two-stage tandem servo system. The tandem servo
hydraulic system supplies hydraulic pressure to operate the main and tail rotor tandem servo-
cylinders to relieve control forces. The system consists of two independent stages. The first stage
supplies 3000 psi to the main rotor tandem servo-cylinders pressure for the first stage tail rotor
tandem servo-cylinder is supplied by the utility hydraulic system. The second stage supplies 3000
psi to the main rotor tandem servo-cylinders and 1500 psi to the tail rotor tandem servo-cylinders.
Each stage consists of a pump, manifold, filter, reservoir, three main rotor tandem servo-cylinders
which are common to both stages, a tail rotor tandem servo-cylinder common to both stages, check
valves, pressure switch, quick disconnects, pressure transmitter, and a pressure gage.
The reservoir supplies the pump which in turn supplies the system with the fluid how and
pressure required. Fluid is directed to the manifold that controls operation of the entire stage and
on to the servo-cylinders that actuate the control surfaces. The first and second stage systems are
controlled by the flight control servo switches on the pilot’s and copilot`s collective sticks The
stages are electrically interlocked, which prevents both stages from being shut off at the same time.
The interlock automatically turns on the system that was off if pressure drops or fails in the
operating system regardless of the flight control servo switch position. Electrical power to operate
the first stage three-way valve comes from the second stage pressure switch. Electrical power to
operate the second stage three-way valve comes from the first stage pressure switch. The first stage
cannot be shut off if pressure in the second stage is below 2000 psi. The second stage cannot be
shut off if pressure in the first stage is below 2000 psi. Low pressure in either stage is indicated by
either the 1ST STAGE SERVO PRESS or 2ND STAGE SERVO PRESS light capsule on the caution panel.
A pilot valve bind in one of the servo-cylinders will be indicated by the lighting of the IST STAGE
SERVO or 2ND STAGE SERVO caution light on the caution panel. Electrical power for the system is
supplied by 28 vdc primary bus. Under normal operating conditions a hydraulic pressure switch
directs electrical power to the flight control servo switches. When hydraulic pressure drops below
operating pressure for a particular stage the pressure switch directs electrical power to light the
low pressure caution light for that particular stage and to the VW system.
The first stage servo hydraulic pump is driven by the accessory section of the main transmission
and is in operation whenever the APP is running or rotor is turning. The second stage servo
hydraulic pump is driven by the No. 2 input of the main transmission.
Hydraulic pressure to either the first stage or the second stage servo system is shut off in the
hydraulic manifold assembly with a three-position first and second stage servo shut off switch, The
switch marked 1ST STAGE OFF 2ND STAGE OFF is on the collective pitch control grip. Both servo
hydraulic systems normally operate with the switch in (on) center position. The 1ST and second
stage servo hydraulic low pressure warning lights, on the caution panel, are marked 1ST STAGE
SERVO PRESS and 2ND STAGE SERVO PRESS and will go on if the pressure is below its respective
switch setting. Two servo system hydraulic pressure indicators are on the center of the instrument
panel, they are marked 1ST STAGE HYD PRESS and 2ND STAGE HYD PRESS and indicate the
MILVIZ SIKORSKY S-64 SKYCRANE - OPERATION MANUAL Page 15
pressure in the first stage servo system and second stage servo system. Malfunction of a pilot valve
in one of the main rotor servo units will be indicated by the lighting of the 1ST STAGE SERVO or
2ND STAGE SERVO caution light on the caution panel.
The utility hydraulic pump is on the accessory section of the main transmission and is driven by
the APP or the main gear box when the rotor system is turning. The pump provides 3,500 psi
hydraulic pressure to operate the engine start system, hoist de-coupler charging system, APP
accumulator charging system, load leveler system and load leveler cable release system. During
flight, when the hoist control switch is placed to PILOT, COPILOT or AFT PILOT, the utility pump
provides control power to the hoist pump.
The AFCS servo system introduces AFCS inputs to the flight control system, supplies pressure to
the tail rotor pedal damper and supplies pressure for the operation of the cyclic stick trim system.
The AFCS servo unit is powered from the 1st stage hydraulic pump system pump, which is driven
by the main gear box or APP. A light on the caution panel marked AFCS SERVO PRESS goes on when
first stage pressure is lost or the system is offline
POWER TRAIN SYSTEM
The transmission system consists of one main transmission and two gear boxes that transmit
power to the main rotor and tail rotor. The main transmission reduces engine rpm and connects the
two engines to the main rotor. A freewheeling unit, at each engine input to the main transmission,
permits the main rotor to auto-rotate without engine drag. An auxiliary power plant (APP) is used
to drive a portion of the accessory section of the main transmission on the ground without engaging
the main rotor head. A freewheeling unit in the accessory gear train of the main transmission,
engages the accessory section when the main rotor is turning.
The APP is disengaged from driving the accessory section through a freewheel unit, forward of
the centrifugal clutch on the APP. Engine torque at reduced rpm is transmitted to the main rotor
and through a tail rotor drive shaft to the intermediate gear box at the base of the pylon. From the
intermediate gear box a pylon drive shaft extends upward to the tail gear box to drive the tail rotor.
The main transmission and the two gear boxes have electric magnetic and non magnetic chip
detectors to warn of foreign particles in the gear box oil. A transmission chip detector caution light
for each gear box is on the caution panel. The chip detector warning light are marked CHIP DET
under the heading INDICATOR LTS on the overhead circuit breaker panel.
The main transmission mounted above the fuselage aft of the engines, contains a tour-stage
reduction gear system to drive the main rotor at a ratio of 48.609 to l. A tail takeoff gear in the main
transmission is coupled to the tail takeoff flange to drive the tail rotor. The intermediate gear box at
the base of the tail rotor pylon contains a bevel gear reduction drive system which reverses the
rotation of the input shafting and transmits torque to the tail gear box. The tail rotor gear box at the
upper end of the tail rotor pylon consists of a bevel gear reduction drive system to transmit torque
to the tail rotor.
MILVIZ SIKORSKY S-64 SKYCRANE - OPERATION MANUAL Page 16
Each of the three transmission system gear boxes has an individual lubricating nil system. Cooling
air is forced through the oil cooler by a blower driven by belts from the tail drive shaft. After
passing through the oil cooler, the oil passes through a filter bypass valve, and relief valve, and
returns to the main transmission where it is sprayed onto the gears and bearings through jets. The
oil filter with dipstick is on the left side of the main transmission. A pressure indicator is on the
instrument panel.
The indicator is marked in pounds per square inch and is actuated by a pressure transmitter at
the main transmission oil inlet port. The pressure indicator operates on 26 vac from the No. l
generator or No. 2 generator or front the ground inverter through the autotransformer. The
transmission oil low pressure caution light, marked TRANS OIL PRESS is on the upper section of the
caution panel. The light will go on when the main transmission oil pressure drops below 13 to I7 psi
at the last oil pressure jet in the box.
The temperature indicator marked XMSN OIL TEMP on the instrument panel is marked in
degrees Celsius. The transmission oil temperature caution light marked TRANS OIL HOT, is on the
upper section of the caution panel. The light will go on when the transmission oil reaches +121 C°.
The intermediate gear box is pressure-lubricated by a vane-type mechanically-driven pump in
she sump of the gear box. A pressure switch in the pump output passage closes when the pressure
decrease to 7 ± 1 psi. This lights the INT TRANS OIL PRESS light on the caution panel and warns the
pilot of a low oil pressure condition. The tail gear box is splash-lubricated from its sump system.
A hydraulic system power package for the rotor brake contains an electric mono-driven pump,
and supplies hydraulic power to the disc-type brake mounted on the left side of the main
transmission. The rotor brake stops rotation of the rotor system and prevents its rotation when she
helicopter is parked. The package consists of a solenoid-operated dump valve, pressure switches,
relief valve, pump, electric motor reservoir accumulator and relay. The rotor brake provides
stoppage of the rotor system with both engines shut off (stop cocked) and a 45-second braking
interval. The 45-second braking interval consists of a natural rotor decay from 100 percent Nr (185
rpm) to 30 percent Nr (55.5 rpm) with the rotor brake applied at a speed not to exceed 30 percent
Nr (55.5 rpm). When the rotor brake switch is applied a time delay relay and a 350 psi pressure
switch installed in the hydraulic line between the rotor brake package and the rotor brake assembly
provides a low energy ( soft) stop for the rotor system. Alter 20 seconds the brake lining pressure
will automatically increase to 600 psi. The rotor brake application can be made only when both N1
engine control levers are in a shutoff position.
An advisory light marked ROTOR BRAKE ON and a caution light marked ROTOR BRAKE PRESS
are controlled by the high and low hydraulic pressure switches to indicate when the solenoid-
operated dump valve is activated. The low pressure switch also acts as a monitor and will light the
caution light, indicating that 10 ± 2 psi pressure exists in the rotor brake.
MILVIZ SIKORSKY S-64 SKYCRANE - OPERATION MANUAL Page 17
The switch marked ROTOR BRAKE, ON, OFF, REL is on the center console. It has a detent in the
OFF position. The switch is momentary at REL and is spring-loaded to return to OFF. When the
switch is placed OFF, the switch is de-energized, but not the rotor brake. To actuate the rotor brake
place the switch ON. This allows 28 vdc from the primary bus to complete the electrical circuit
through the high pressure switch, closing a relay. The ROTOR BRAKE PRESS caution light will go on
when the low pressure switch is closed, as a result of the hydraulic pressure developed by the
pump.
When the pump develops 600 psi to open the high pressure switch, the ROTOR BRAKE ON
advisory light will go on, indicating hydraulic pressure is being directed to the brake. Pump
pressure is controlled by the high pressure switch. Whenever hydraulic pressure drops below
minimum and the rotor brake switch is ON, the advisory light will go off and the pressure switch
will activate a relay to energize the pump. Hydraulic pressure is built up again and the advisory
light goes on. To release the brake, hold the rotor brake switch to REL. This will energize the dump
valve, discharging the pressure to the reservoir. The advisory light and the caution light will go off
indicating the brake is offline. The rotor brake is powered from the 28 vdc primary bus.
The advisory light, on the advisory panel of the instrument panel, is marked ROTOR BRAKE ON.
The light goes on when the high pressure switch opens to pressurize the rotor brake. The light will
remain on until the pressure is below 600 ± 50 psi. The light is powered by the 28 vdc primary bus
and is protected by a circuit breaker on the panel marked ROT OR BRAKE ON. The caution light, on
the caution panel, is marked ROTOR BRAKE PRESS, The light goes on when the rotor brake switch
is placed ON, indicating that the pump is building up pressure in the system. The light will remain
on until the system pressure drops below the minimum setting of the switch.
MAIN AND TAIL ROTOR GROUP
The main rotor system consists of six rotor blades, the rotor head and the linkage necessary to
transmit main rotor flight control movement to the blades. The blades are hinged at the rotor head
so that each blade is free to flap (move vertically), hunt (move horizontally) and turn about its axis
to change the angle of incidence (pitch). Droop restrainers, attached to the hub, limit the downward
movement. Anti-napping restrainers limit the upward movement when the blades are stopped or
turning at low speed.
At about 26 percent Nr, centrifugal force releases the anti-flapping restrainers. When rotor speed
is increased to about 65-70 percent Nr, centrifugal force releases the droop restrainers (stops) and
the blades are supported by the combination of centrifugal force and lift. Hydraulic dampers,
connected between each blade and the rotor hub, lessen hunting movement about the vertical hinge
as the blades rotate. Prevent shock when the rotor is started or stopped, and help to prevent
ground resonance.
The angle of incidence or pitch is controlled by the flight control system which connects to the
blades through the swashplate assembly at the lower section of the rotor head. The swashplate
assembly consists of an upper swashplate and a lower swashplate. Both swashplates are mounted
on a ballring and socket which keeps them parallel, but allows them to be tilted, raised, or lowered
MILVIZ SIKORSKY S-64 SKYCRANE - OPERATION MANUAL Page 18
simultaneously by the rotor night control system. Linkage on the rotating swashplate transmits the
control motion to the blades.
Except for steel cuffs the blades are constructed of aluminum alloy. The cuffs attach the root ends
of the blades to the sleeve-spindles on the main rotor hub. The blade is pressurized with nitrogen to
approximately 10 psi at an ambient temperature of + 18 to +24 C°.
The tail rotor system consists of four blades, the tail rotor hub, and the pitch change mechanism.
The tail rotor hub is splined to the tail gear box drive shaft which transmits engine torque to the
blades. The blades are attached to a semi-rigid, self-lubricated rotor head. The blade pitch changing
mechanism transmits tail rotor flight control movement through the hollow tail gear box drive shaft
to the blades.
PITOT HEAT
To prevent icing, both pitot tube heads may be healed electrically by turning on a single switch
marked PITOT HEAT and OFF on the MASTER SWITCH panel on the center console. The switch
controls the operation of both pitot heaters. Caution lights marked PILOT PITOT HEAT – COPILOT
PITOT HEAT will light to indicate dc power failure to the associated pilot tube.
MILVIZ SIKORSKY S-64 SKYCRANE - OPERATION MANUAL Page 19
ELECTRICAL POWER SUPPLY AND DISTRIBUTION SYSTEM
The primary electrical system is a 115/200-volt alternating current (ac) system. Alternating
current is rectified to provide a 28-volt direct current (dc) system. Power for the ac system is
supplied by two generators. Secondary sources of ac power are the dc-operated inverter and an ac
external power receptacle.
Alternating current (AC) power supply system
MILVIZ SIKORSKY S-64 SKYCRANE - OPERATION MANUAL Page 20
The two 115/200 volt, three-phase, 400 Hz generators are mounted on, and driven by, the
accessory section of the main gear box. The accessory section is driven by the APP when the rotor
system is static and by the main gear box after rotor engagement. Each generator output is routed
through a supervisory panel. The supervisory panels provide control and feeder fault protection,
and protect the system from under-frequency, under-voltage, and over-voltage.
The under-frequency protection is not available when the weight of the helicopter is removed
from the landing gear wheels. Bach generator furnishes power to its respective ac primary and
monitored bus when both generators are operating, the generator switches are in the l and 2 (ON)
position, and the supervisory panel is satisfied that voltage and frequency output is within
prescribed limits failure of either generator is indicated by lights on the caution panel. Each
generator control circuit is protected by circuit breakers, marked NO. 1 and NO. 2 ENGINE GEN, on
the overhead circuit breaker panel.
The generator switches, on the center console, are marked GEN NO. l, ON, NO. 2, OFF-RESET, and
TEST. Placing the switches to ON connects generator power to appropriate buses. Placing either
generator switch to OFF RESET will take the respective generator on' the line. If a temporary
overvoltage condition occurs, placing either switch to OFF RESET, then ON, will bring the generator
back on the line. TEST is used by personnel during maintenance checks. Failure of either generator
is indicated by lights on the caution panel marked NO. 1 GENERATOR and NO. 2 GENERATOR. The
lights are powered by the dc primary bus.
Three autotransformers step ac voltage down from 115 volts to 26 volts for operation of certain
instruments and radio equipment. Each autotransformer receives power from an ac primary bus
when energized, or from the ground inverter when the ac primary buses are not energized. Each
autotransformer is protected by a circuit breaker marked AUTO-XFMR, both of which are on the
overhead circuit breaker panel. Each individual circuit breaker is labeled for the system or
component it protects. If a circuit overloads, the circuit breaker will pop out, removing electrical
power from the protected system or component. The circuit is reactivated by pushing in the circuit
breaker button.
The 250 VA inverter in the attic provides 115 vac to essential ac-operated equipment and to the
autotransformers which provide 26 vac for prestart for some of the instruments. The inverter is
powered from the dc primary bus until the ac primary buses are energized. When the ac primary
buses are energized, by one or both generators or ac external power, the inverter control relay is
energized and the inverter is automatically shut off. The inverter control relay is protected by an ac
circuit breaker marked INV CONT. The inverter switch on the center console is marked INV, ON and
OFF. The switch should be left ON to allow the inverter to automatically turn on if the ac primary
buses are de-energized. Placing the switch OFF will disconnect the dc power source to the inverter.
The inverter switch should be ON during pressure refueling operations.
The 115/200 vac external power receptacle is on the right side of the fuselage below the pilot’s
window. External power is controlled by a switch in the pilot's compartment. The external power
switch is on the master switch panel portion of the center console. The switch is marked EXT PWR,
ON, and OFF, The ON position connects 28 vdc power from the battery to the external power
MILVIZ SIKORSKY S-64 SKYCRANE - OPERATION MANUAL Page 21
monitor panel when ac power is plugged in. This circuit monitors ac input and, if satisfactory,
permits ac external power to energize the ac electrical system. The OFF position prevents ac
external power from energizing the ac buses.
Alternating Current Distribution.
Power for the operation of AC electrical equipment is distributed through appropriate
supervisory panels and bus relays to a primary and monitored bus for each generator. Since both
generator bus systems operate the same, they will be discussed under the general headings of
primary buses and monitored buses. The AC primary buses are energized by their respective
generators, or by external power when connected and the generators are not operating. The
primary buses are energized through their primary bus relays that are energized when the
generator switches are placed ON. Power is then distributed to AC operated equipment, respective
transformer rectifiers, and autotransformers. Failure of either generator will cause the operating
generator to energize both primary buses and automatically drop both monitored buses from the
system.
The monitor buses are energized by the respective generators when both are operating, or by
external power when connected and the generators are not operating. The monitor buses are
energized through respective double monitor bus relays that are energized by power from the
generator switches through the primary bus relays. When both generators are operating, each
double monitor bus relay is energized to connect generator power to the monitor buses. The power
normally received by the monitor buses does not pass through the primary bus relays, but is
received from the supervisory panel output before being distributed to the primary buses. This
makes power available to both primary buses if either generator or transformer rectifier should
fail. Then, the double monitor bus relay of the failed circuit will be completely de-energized while
the double monitor bus relay of the operating circuit will be energized to the connecting but to
provide a path for the operating generator to energize both primary buses. Both monitor
buses will be automatically dropped from the system if either generator fails.
MILVIZ SIKORSKY S-64 SKYCRANE - OPERATION MANUAL Page 22
Direct Current (DC) Power Supply System
The DC power supply system is powered by two converters (transformer-rectifiers). A battery
and DC external power receptacle provide a secondary source of DC power. The two 200-arnpere
converters (transformer-rectifiers), each normally powered by a respective generator and AC
primary bus, convert AC power to DC power. Each converter is controlled by a switch on the center
console. The converters are connected directly to the DC system whenever they are energized by
AC power and the converter switches are ON. Failure of either converter will be indicated by lights
on the caution panel marked # 1 RECTIFIER and # 2 RECTIFIER respectively.
MILVIZ SIKORSKY S-64 SKYCRANE - OPERATION MANUAL Page 23
The converter switches marked RECT 1 and 2 and OFF, are on the center console. When both
switches are placed ON. DC power is connected through the respective reverse current cutout relay
to the primary and monitored buses. With both switches OFF, and the battery switch ON, battery
power only will be available to the DC primary bus. When only one switch is placed OFF, the
monitored bus will be dropped from the system and only the primary bus will be energized.
The 24-volt, 22 ampere hour, nickel-cadmium battery, in the nose section forward of the pilot’s
compartment, is reached from outside the helicopter for maintenance. Battery power is used for
limited ground operations and as an emergency source of power to the DC primary bus, if both
generators or both converters should fail. If either converter is on, it will recharge the battery when
the battery switch is ON. When DC external power is used for more than 45 minutes, the battery
switch should be set to OFF, to avoid overcharging the battery. The battery switch on the center
console, marked BAT, has marked positions ON and OFF. When the switch is placed to ON, battery
power is supplied to the DC primary bus. Placing the switch to OFF will disconnect battery power
from the primary bus.
The 28-volt external power receptacle is on the right side of the fuselage below the pilot's
window. The dc external power system uses the same control switch and advisory light as for the ac
external power system.
Direct Current Distribution.
The power for the operation of DC electrical equipment is distributed through a primary and
monitored bus. The DC primary bus supplies power for operation of all equipment necessary for
mission accomplishment. The primary bus is powered by either or both converters when their
switches are ON or by either converter if the other should fail. The primary bus is also energized by
external power, when connected. And the external power switch is ON, or by the battery when the
battery switch is ON. The monitor bus supplies power for operation at equipment not considered
essential for mission accomplishment. The monitor bus is powered by both converters, when both
are operating and the switches are ON, or by external power, when connected, and the external
power switch is ON. Loss of either converter or either generator will cause the monitor bus relay to
be de-energized and drop the monitor bus from the system. Battery power is not supplied to the
monitor bus.
AUXILIARY POWER PLANT
The auxiliary power plant (APP) aft of the main transmission, is normally used for ground
operation of the hydraulic and electrical systems. The APP system consists of s control pane, a
hydraulic starter motor, a turbine engine, a self-contained oil system, a mechanical drive, an
accumulator, and a hydraulic hand-pump. The hydraulic starter motor is driven by 3000 to 3500
psi hydraulic pressure from the accumulator. The starter motor, in conjunction with 28 DC from the
primary bus, is used to start the APP. After a normal start, the accumulator must be recharged by
motoring either engine starter. If the initial accumulator charge is lost and the APP is not started,
the hydraulic hand-pump should be used to recharge the accumulator.
MILVIZ SIKORSKY S-64 SKYCRANE - OPERATION MANUAL Page 24
During the normal starting sequence, the dc primary bus activates the automatic start and control
circuitry for operation of the APP. A fuel pressure switch opens on an increasing fuel pressure at
110 ± 10 psi, to allow fuel to enter the combustion chamber through both the start fuel nozzle and
the main fuel ejector. The switch also energizes the ignition circuit, starts the hour meter, and lights
the APP ON advisory light. At 90 percent speed, the fuel start valve, ignition and APP starter are
shut off by the speed switch and burning is self-sustaining as long as there is a flow of fuel through
the main fuel valve. A mechanical drive with a centrifugal clutch drives the main transmission
accessories section. The centrifugal clutch contains a freewheel unit that enables shutdown of the
APP when the rotor head is engaged.
Fuel for the APP is supplied by the all fuel tank. Average fuel consumption under lull load is 89
pounds per hour. Under normal operation, the APP provides about 72 shaft horsepower to drive
the accessory gear case. An electric fuel pump is installed in the hoist well to provide a flow of fuel
to the APP before attempting a start. When the APP MASTER switch is placed ON, 28 DC from the
primary bus opens the APP fuel shutoff valve, and operates the APP electric fuel pump. Fuel is then
pumped through the shutoff valve to the APP mechanical fuel pump at about 8.5 psi.
The APP control panel on the overhead control panel, provides controls for operation of the APP.
The control panel consists of a MASTER SWITCH, APP START and STOP switches, LOW OIL PRESS,
HIGH EXH TEMP, and OVSP warning lights, tachometer indicator and an exhaust temperature gage.
A FIRE warning light is also on the control panel. The APP MASTER switch has marked positions ON
and OFF. When the switch is placed ON. Electrical power is supplied to the START switch. the APP
fuel shut off valve is energized to the open position, the APP electrically-operated fuel pump is
activated, the OVSP and HIGH EXH TEMP warning lights go on and current holds the relays
energized when the START switch is released. When the switch is placed OFF, the entire system is
deactivated. The switch operates on 28 DC from the primary bus.
The APP START switch activates the automatic start and control circuitry. When the switch is
depressed, electrical power energizes a series of relays and the OVSP and HIGH EXH TEMP wanting
lights go on. These relays activate the hydraulic start valve, the main and start fuel valves, the
ignition unit. The fuel pressure switch, the low oil pressure switch, an advisory light marked APP
ON and an exhaust temperature gage. When the switch is released holding current from the
MASTER switch keeps the series of relays energized until the operating conditions that will sustain
operation are met.
The APP STOP. when depressed, shuts down the APP system. It does not, however, deactivate the
entire system. The warning lights and the airframe fuel shutoff valve are controlled by the APP
master switch. If the APP STOP switch fails to shut down the APP, placing the MASTER switch OFF
will de-energize (close) both the APP main fuel valve and the airframe APP fuel shutoff valve.
The low oil pressure warning light, marked LOW OIL PRESS, goes on when oil pressure drops to 6
± l psi. This condition will automatically shut down the APP. The high exhaust temperature warning
light, marked HIGH EXH TEMP, will go on and the APP will automatically shut down when the
exhaust temperature reaches about + 1090 F. The overspeed warning light, marked OVSP, will go
on if engine speed reaches 110 percent speed, to indicate the APP has been automatically shut
down because of an overspeed condition.
MILVIZ SIKORSKY S-64 SKYCRANE - OPERATION MANUAL Page 26
MAIN INSTRUMENT PANEL - Pilot Side
1. 1ST stage hydraulic pressure
2. 2nd stage hydraulic pressure
3. Utility hydraulic pressure
4. Forward fuel tank quantity
5. Aft fuel tank quantity
6. Auxiliary tank quantity
7. N1 rpm
8. T5
9. Fuel flow
10. Engine pressure ratio
11. Engine oil temperature
12. Engine oil pressure
13. Hoist cable length
14. Hoist payload
15. Transmission oil temperature
16. Transmission oil pressure
17. Performance indicator
18. VOR 1 indicator
19. Clock
20. Airspeed indicator
21. Engine torque
22. Rotor/N2 rpm
23. FIRE warning light
24. MASTER CAUTION light
25. Attitude indicator
26. ADF indicator
27. Altimeter
28. Vertical speed indicator
29. Turn and slip rate indicator
30. VOR 2 indicator
31. Aux fuel pumps
32. Caution and advisory panel
MILVIZ SIKORSKY S-64 SKYCRANE - OPERATION MANUAL Page 27
MAIN INSTRUMENT PANEL - Copilot Side
1. Airspeed indicator
2. Engine torque
3. Rotor/N2 rpm
4. Fire warning light
5. MASTER CAUTION light
6. Attitude indicator
7. VOR2 indicator
8. Altimeter
9. Vertical speed indicator
10. Turn and slip rate indicator
11. Auxiliary fuel pumps
12. Clock
13. Caution and advisory panel
MILVIZ SIKORSKY S-64 SKYCRANE - OPERATION MANUAL Page 28
MIDDLE CONSOLE
1. Com frequency
2. Com frequency set knob (mouse
wheel=setting, right click=whole/fract)
3. Nav 2 frequency
4. Nav 2 frequency set knob (mouse
wheel=setting, right click=whole/fract)
5. ADF 1 radio
6. ADF 2 frequency
7. ADF 2 frequency set knobs (mouse wheel)
8. FWD TANK PUMP #1 and #2 switches
9. Fuel transfer switch
10. AFT TANK PUMP #1 and #2 switches
11. EAPS/ANTI-ICE switches
12. PITOT HEAT switch
13. TAIL SKID control switch
14. BAT switch
15. NO.1 and NO.2 GEN switches
16. NO.1 and NO.2 RECT switches
17. INV switch
18. IGN NO.1 and NO.2 switches
19. Nav 1/ADF 1 frequency
20. Nav/ADF mode select
21. Nav/ADF frequency set knobs
22. Cargo release mode (SHEAR HOOK=on,
SAFE=off – hoist and water ballast)
23. Hoist master knob (PILOT, or AFT PILOT
to control – hoist and water ballast)
24. Cable shear switch (release hoist ‘n water)
25. ROTOR BRAKE switch
26. Landing light switch
27. 1ST STAGE OFF 2ND STAGE OFF switch
MILVIZ SIKORSKY S-64 SKYCRANE - OPERATION MANUAL Page 29
OVERHEAD PANEL
1. Cockpit lights
2. Position lights
3. Anti collision lights
4. Hoist shear test
5. Fire detector cont. wire test switch
6. Fire surveillance facility test knob
7. Extension tank empty warning lights
8. Extension fuel tank transfer switches
9. APP exhaust gas temperature
10. APP rpm
11. APP MASTER switch
12. APP START switch
13. APP STOP switch
14. APP warning lights
15. APP FIRE warning light
16. Engine overspeed test switches
17. Transponder
18. 4 point jettison shear arm switches
19. point jettison shear switch
MILVIZ SIKORSKY S-64 SKYCRANE - OPERATION MANUAL Page 30
ENGINE CONTROL UNIT
1. Engine FIRE warning lights
2. Engine starter buttons (on outside of N1 levers)
3. N1 levers
4. Fuel shut off levers
MILVIZ SIKORSKY S-64 SKYCRANE - OPERATION MANUAL Page 31
AFT PILOT'S COMPARTMENT
1. Hoist cable length
2. Engine torque
3. Hoist payload
4. 4 point jettison armed light
5. Hoist shear switch
6. 4 point jettison test
MILVIZ SIKORSKY S-64 SKYCRANE - OPERATION MANUAL Page 32
NORMAL PROCEDURES
Note: It is recommended to load the helicopter into the flight simulator's default flight and shutoff all of the
systems manually if you want to perform the startup from a cold and dark cockpit.
I. AFT COMPARTMENT
1. Miscellaneous control switches OFF
2. 4 point leveler shear switch OFF, SECURED
3. Hoist control shear switch OFF, SECURED
4. INT switches AS REQ.
II. FORWARD COMPARTMENT
1. Parking brake LOCKED
2. Nose wheel parking lock LOCKED
3. Rotor brake switch ON
4. Cable shear switch SECURED
5. Hoist control switch OFF
6. Cargo release mode SAFE
7. 4 point leveler switch OFF
8. External power switch OFF
9. INV switch OFF
10. IGN switches OFF
11. GEN switches OFF
12. RECT switches OFF
13. POD PWR switch OFF
14. TAIL SKID switch AS REQ.
15. EAPS/ANTI-ICE switches OFF
16. HEATERS switch OFF
17. Avionics OFF, AS REQ.
18. GYRO switch NORM
19. COMPASS SLAVING switch NORM
20. CIPR ICS switch NORM
21. Altimeters set to field elevation
III. PRE START
1. Fuel shut off levers SHUT OFF
2. N1 levers SHUT OFF
3. ANTI-COL switches ON
4. Light switches AS REQ.
5. APP MASTER switch OFF
6. 4 point shear AREMD light OFF
7. WINDSHIELD switch OFF
MILVIZ SIKORSKY S-64 SKYCRANE - OPERATION MANUAL Page 33
8. BAT switch ON
9. Caution and advisory panel TEST
10. APP MASTER switch ON
11. FIRE warning light TEST
12. HIGH EXH TEMP and OVSP light ON
13. LOW OIL PRESS light TEST (press)
14. APP START switch PRESS
15. APP warning lights GO OFF
16. APP rpm ~100%
17. GEN and RECT switches ON
18. GENERATORS and RECTIFIER lights GO OFF
19. Transmission press ~50-65psi
20. Utility hydraulic press ~1300-1800psi
21. 1st hydraulic press ~2600-3300psi
22. TRANS OIL PRESS, AFCS SERVO PRESS,
1ST SERVO PRESS, 1ST SERVO,
1ST STG TL ROTOR SERVO lights GO OFF
25. INV switch ON
26. NO.1 GEN switch OFF
27. #1 GENERATOR light ON
28. #1 RECTIFIER light OFF
29. NO. 2 GEN switch OFF
30. #2 GENERATOR light ON
31. #1 and #2 RECTIFIER lights ON
32. NO. 1 GEN switch ON
33. #1 GENERATOR, #1 and #2 RECTIFIER lights OFF
34. NO. 2 GEN switch ON
35. #2 GENERATOR light OFF
36. Avionics AS REQ.
37. PITOT HEAT switch ON, AS REQ.
38. PILOT and CO-PILOT PITOT HEAT lights GO OFF when on
39. AFCS SERVO button PRESS
40. AFCS SERVO PRESS, AFCS SERVO lights GO OFF
41. AFCS 1 and AFCS 2 button PRESS, on
42. Rotor brake switch REL
43. FWD TANK PUMP #1 and #2 switches ON
44. Engine No.1 shut off lever OPEN (fwd)
45. FUEL PRESS #1 light GO OFF
46. AFT TANK PUMP #1 and # switches ON
47. Engine No.2 shut off lever OPEN (fwd)
48. #2 ENG FUEL PRESS light GO OFF
49. Avionics AS REQ.
50. CONTINOUS WIRE switch SHORT
MILVIZ SIKORSKY S-64 SKYCRANE - OPERATION MANUAL Page 34
51. Fire warning lights OFF
52. CONTINOUS WIRE switch TEST
53. Fire warning lights ON
54. CONTINOUS WIRE switch OFF
55. Fire warning lights OFF
56. IR SURVELLIANCE rotary switch ROTATE
57. Respective fire warning lights ON
58. Master FIRE warning light PRESS, reset
59. IR SURVELLIANCE rotary switch NORM
60. All FIRE warning lights GO OFF
61. Anti collision light ON
62. Position light ON
IV. ENGINE No.1 START
1. N1 lever SHUT OFF
2. IGN SW No.1 switch NORM
2. Starter button PRESS (2-5 sec.)
3. #1 ENGINE STARTER GO ON
4. N1 lever GRD IDLE
5. N1 speed ~40%
6. T5 min 515 C° max +525 C°
7. Eng oil press min 20 psi
8. N1 lever FULL OPEN
9. INTER TRANS OIL PRESS, XMSN OIL PRESS,
XMSN OIL TEMP lights GO OFF
V. ENGINE No.2 START
Use same procedure as used to start No.1
2. System check Normal at GRN IDLE
3. N1 lever FULL OPEN
VI. AFTER ENGINES START
1. EAPS/ANTI-ICE switches AS REQ.
2. ENGINE OVERSPEED ENG2 switch TEST
3. Engine No.2 DECREASE
4. ENGINE OVERSPEED ENG2 switch NORM
5. Engine No.2 NORMAL
6.ENGINE OVERSPEED ENG1 switch TEST
7. Engine No.1 DESCREASE
8. ENGINE OVERSPEED ENG1 switch NORM
9. Engine No.1 NORMAL
10. APP STOP switch PRESS
MILVIZ SIKORSKY S-64 SKYCRANE - OPERATION MANUAL Page 35
11. APP MASTER switch OFF
12. IGN SW No.1 and No.2 OFF
VII. BEFORE TAKE-OFF CHECK
1. N1 levers FULL OPEN
2. Parking brake handle AS REQ.
3. AFCS ON
4. Caution and advisory panel CHECK
5. System instruments CHECK
VIII. BEFORE LANDING CHECK
1. N1 levers FULL OPEN
2. PARKIG BRAKE handle AS REQ.
3. Caution and advisory panel CHECK
4. System instruments CHECK
5. Landing light AS REQ.
IX. ENGINE SHUTDOWN
1. AFCS 1 and 2 OFF
2. Collective FULL DOWN
3. PARKING BRAKE APPLY
4. Heater, VWS switch OFF
5. APP START
6. N1 Levers GRD IDLE
7. Collective RAISE ~2% (to 85% Nr)
8. N1 levers SHUT OFF (one at a time)
9. Cable shear switch COVER
10. ROTOR BRAKE switch ON (below 30% Nr)
11. POSITION LIGHT switches OFF
12. ANTI-COLLISION light OFF
13. Fuel shutoff levers SHUT OFF
14. All switches OFF
Except: AFCS SERVO, STICK TRIM, BAT, INV, ROTOR BRAKE
15. APP STOP switch STOP
16. APP MASTER switch OFF
17. INV switch OFF
18. All light OFF
19. BAT switch OFF