The Systems Enabling Objectives

THE SYSTEMS ENABLING OBJECTIVES From CNATRA P-402 (Rev. 7-98)

CHAPTER 1 INTRODUCTION TO THE TH-57

1.1. Described the basic design characteristics and capabilities of the TH-57 helicopter.

Land-based, skid-configured, utility type helicopter, designed to land and takeoff from reasonably level terrain. An Allison 250C-20J turboshaft engine powers the main rotor, which is used for lift and thrust, and the tail rotor, which is used to counteract torque and provide yaw control. The maximum takeoff Gross weight is 3,200 lbs. The maximum forward air speed at sea level standard day is 130 kts, maximum sideward airspeed is 25 kts and maximum reward airspeed is 15 knots.

Dimensions Main rotor diameter: 33’-4” Main rotor to ground clearance: 11’-7.5” / 6’-5” Main rotor to tail boom clearance: 1’-1.5”

1.2. Recognize the five major sections of the airframe.

1. Forward / cabin 2. Landing gear 3. Cowling 4. Tail boom 5. Vertical fin

1.2.1. Recognize the items associated with the forward section.

Construction Primarily constructed of aluminum honeycomb covered with either fiberglass or aluminum skin. Provides excellent weight-to-strength ratio and also helps sound proof that cabin area.

Battery compartment, windshield, control column, cockpit, passenger compartment, aft cabin deck, baggage compartment, aft electrical compartment, aft cabin space.

1.2.2. Identify the components of the landing gear.

1. 2 aluminum alloy curved cross tubes 2. 2 aluminum skid tubes 3. formed steel tail skid 4. tow rings 5. replaceable skid shoes 6. ground handling wheels

1.2.3. State the main purpose of the cowling section.

Streamlines air flow to reduce drag

1 1.2.4. State the purpose of the tail boom, and identify its associated parts and their functions.

The tail boom provides an extended support to mount the tail rotor and vertical fin (By increasing the moment arm of the tail rotor it allows the size of the tail rotor to be smaller than the main rotor). The tail boom is mounted to the forward section via a mounting pad and four bolts.. Horizontal Stabilizer: Negative camber pushes the tail down to allow near level forward flight.

1.2.5. Described the purpose of the vertical fin and state how it provides directional stability and reduces tail rotor loads.

Construction Semi-monocoque made of aluminum honeycomb and aluminum skin mounted at a 5.5o offset from the longitudinal axis.

Purpose Provides directional stability and reduced tail rotor loads at cruise airspeed. This is accomplished by a 5.5o offset from the longitudinal axis. This offset provides a horizontal lift component that assists in countering torque. (The greater the airspeed, the greater the lift.)

1.3. Identify the components of the cargo hook assembly with its associated capabilities and limitations. Components 1. Frame 2. Hook assembly 3. 1 Electrical and 2 manual (one in the cockpit and one on the hook assembly) release

Capabilities and Limitations Structural weight capacity of 1500 lbs

1.4. Described aircraft lighting system and identify the associated lights, switches, and circuit protection.

Exterior lights 1. Position Lights: Controlled by 3 switches on the overhead panel: On / Off, Bright / Dim, and Steady / Flashing. On the ground at night during periods of low rotor RPM use Flashing / Bright to signal other aircraft not to taxi near you. 2. Anti-Collision lights: 2 lights controlled by a single overhead switch (On / Off): Top of vertical fin and below fuselage aft of the baggage compartment. Circuit protection is provided by the anti-collision circuit breaker type switch located on the overhead console. 3. Landing Light: controlled by a single overhead switch. Light is located just aft of forward cross tube. CAUTION: 10 minute operating limitation due to fire hazard. 4. Search Light: Controlled by 2 switches on the pilot’s collective. On / Off / Stow Switch: CAUTION Once the automatic Stow cycle is complete the switch should be turned OFF. If left in Stow a faulty heat switch may cause the extend / retract motor to burn out. Directional Control Switch: The light is adjustable from 0o to 120 o of extension. From 0 o to 60 o the light will rotate 90 o left or right. From 60 o to 120 o the light will rotate 360 o. Circuitry is protected by two circuit breaker, labeled SEARCHLIGHT POWER and SEARCHLIGHT CONT.

2 Interior Lights 1. Instrument Lights – Controlled by overheat On / Off Rheostat. This does not control all Cockpit lights: Caution Panel has it’s own Bright – Dim switch. UHF Control has it’s own Dim switch as well. 2. Cockpit Lights – 2 Grimes lights mounted on the control column. 3. Cabin Lights – 2 overhead lights

1.5. Identify the seat restraints, doors, and caution panel within its associated capabilities and limitations.

Seat Restraints Pilot and copilot seats are equipped with a lap safety belt and inertia reel shoulder harness with a manual lock-unlock handle. The reel will automatically lock when the helo encounters a longitudinal impact and a deceleration force of 2 to 3 G’s.

Doors Both cockpits are accessed via forward hinged doors that are jettisonable in the event of an emergency landing or are removable during the execution of external lifts.

Caution Panel The caution panel is located on the instrument panel. Illumination of any of the lights on the caution panel alerts the pilot to a system fault or condition. The caution panel is powered by ESS #2 BUS and is protected by the CAUTION LT circuit breaker.

3 CHAPTER 2 ALLISON 250C–20J TURBOSHAFT ENGINE

2.1. Identify the type of gas turbine engine use in the TH – 57.

Allison 250C-20J series turboshaft engine. This engine is an internal combustion gas turbine engine featuring a “free” power turbine consisting of a combination axial- centrifugal compressor; a single “can” type combustor; a turbine assembly which incorporates a two-stage power turbine and exhaust collector; and an accessory gear box which incorporates a gas producer gear train and a power turbine gear grain.

2.1.1. Stated the power output of the 250C–20J engine.

Allison’s 250C-20J engine is a 420 SHP engine derated by Bell Helicopter to 317 SHP due to power train limitations.

2.2. Identify the four sub-assemblies of the engine.

1. Compressor Section 2. Accessory Gearbox Section 3. Turbine section 4. Combustion Section

2.2.1. Identify the four sub-assemblies of the compressor.

1. Front Support 2. Compressor rotor wheels and blades 3. Case Assemblies 4. Diffuser Scroll

2.2.1.1. State the three functions served by the front support of the compressor and identify how each function is achieved.

1. Hollow struts serve a part of the anti-ice system- 500oF air from the diffuser scroll is channeled into the struts via the anti-valve. This heated air prevents the inlet vanes from cooling to the point of freezing. 2. Struts serve as inlet guide vanes, ensuring ambient air strikes the first stage compressor at the proper angle. 3. Support the forward end of the compressor shaft via bearings and struts.

2.2.2. Described the compressor used in the 250C–20J engine by identifying the components, the types of compression utilized, the number of stages of compression used, and where the components are mounted.

Compressor: A 6 stage axial compressor (rotor followed by stator) followed by a single stage centrifugal compressor provides a compression ratio of 6.5 to 1 and a temperature of 500oF. (The compressor bleed air valve is located at the 5th stage of axial compression)

Diffuser Scroll: Mounted on output of centrifugal compressor; used to direct air from compressor to the Air Transfer Tubes. Houses 4 bleed air taps: 1. Anti-Ice 2. Cabin Heating

4 3. Fuel Control / Nf Governor control Air 4. Compressor Bleed Air Valve Control Air

Air Transfer Tubes: Transfers air from diffuser scroll to combustion section.

2.2.3. State the main purpose of the compressor bleed air system.

Bleeds air from the 5th stage of the compressor during starting, acceleration, and at low compressor pressure ratio operation to prevent / minimize compressor stalls and surging.

2.2.3.1. Identify the location of the compressor bleed air system.

Compressor bleed air valve is located at the 5th stage of axial compression approx. 2 ‘o clock.

2.2.4. Trace the path of air flow from the air intake through the engine by identifying the components in the proper sequence with respect to flow.

Ambient air enters the front support and is guided by the vanes to strike the first stage compressor rotor at the proper angle. The rotor accelerates the air and directs it to the stator which slows the air, increasing static pressure. The stator then directs the air to the second stage rotor. This process continues through the six stages of axial compression. At the 5th stage the compressor bleed air valve bleeds air as necessary to prevent a compressor stall. After the 6 stages of axial compression, air enters the centrifugal compressor (1 stage). After the single stage of centrifugal compression air is collected, in the diffuser scroll (6.5 times compression, 500oF) and a portion is bled off (Bleed Air Valve Control Air, Cabin Heating, Anti-Icing, Fuel Control / Nf Governor Control Air).

The remaining air is directed from the diffuser scroll to the combustion section via the air transfer tubes. In the combustion section 75% of the air is used for combustion cooling and the remaining 25% is mixed with fuel for combustion.

The hot expanding gasses from the combustion section move forward to the 4 stage turbine section. First the gasses reach the 2 stage Gas Producer turbine st assembly (Ng). The gasses are accelerated by the 1 stage nozzle and then directed to the 1st stage turbine wheel which converts heat energy to mechanical energy. The 2nd stage of the gas producer turbine identical to the 1st (Nozzle followed by the turbine wheel). The 2 stage Gas Producer turbine assembly extracts 2/3’s of the available energy from the hot expanding gases and drives g the compressor on the Ng gear train. The hot gases exit the N turbine and pass over 4 TOT thermocouples which sense the temperature and send an averaged signal to the 28v DC TOT gauge on the instrument panel. After the 4 TOT Thermocouples the gases continue forward to the 2 stage Free Power Turbine (Nf) which consists of two sets of Nozzle then wheel pairs. The Nf extracts the remaining 1/3 of energy from the gases to drive the Nf gear train.

The remaining exhaust gases are collected in the exhaust collector and then out the twin exhaust.

2.2.5. Identify the type of combustor used in the 250C – 20J engine.

Single Can Type

5 2.2.5.1. Identify the location of the components in the combustion section and state their purpose.

Outer Case and Inner Liner: Provides thermal separation from the combustion chamber and the exterior of the combustor section. Also provides a flow path for cooling air from the Air Transfer tubes.

2.2.5.2. State the purpose of air flow through the compressor and amount used for each purpose.

75% for cooling, 25% for combustion

2.2.6. State the purpose of the turbine section in the 250C – 20J engine.

Converts heat energy to mechanical energy

2.2.6.1. Identify the components of the turbine section and state their function.

1. 2 stage Gas Producer Turbine: Drives the compressor and Ng gear train. 2. 2 stage Free / Power Turbine: Drives the Nf gear train. 3. Thermocouple Assembly: 4 thermocouples provide signal to TOT gauge

2.2.6.2. State what portion of the available energy is used to drive the gas producer and free power turbine sections.

2/3’s Compressor and Ng drive train (51,989 rpm at 100%) 1/3 Nf drive train (33,956 rpm at 100%)

2.2.6.3. Identify the location and name of the components used for measuring turbine outlet temperature (TOT).

System Requires 28v DC 4 TOT thermocouples between the Ng and Nf turbines TOT indicator circuit breaker TOT gauge TOT Light (Comes on when START limits are exceeded. Maintenance must reset with a key in the battery compartment)

2.2.7. Identify the functions of accessory gear box.

Primary structural member for the engine. Provides mount for compressor and turbine sections and 4 engine mounts (3 used to mount the engine, 4th used to mount the anti-ice motor)

2.2.7.1. Identify the accessories driven by the gas producer.

Starter Generator Tach generator (Ng) Oil Pump (Shares the same shaft as the Ng Tach) Fuel Pump Fuel Control Unit Standby Generator

6 2.2.7.2. Identify the units driven by the power turbine gear train.

Governor, Nf Output Shaft Tach generator (Nf) Torque meter

2.3. Identify the type of engine oil system utilized in the 250C – 20J.

Pressurized, circulating, dry-sump system (Dry sump, meaning oil reservoir is located separately from the oil pump.)

2.3.1. State the function and location of each component of the engine oil system.

5.5 quart oil reservoir: Holds engine oil and provides mount for the oil temp bulb. Oil Pump: Located inside the accessory gear box, driven by same shaft which drives the Ng tach generator. Spur gear oil pump with 1 pressure and 4 scavenge elements. Delivers oil under pressure the Internal Oil Filter Assembly. Internal Oil Filter Assembly: Contains Pressure Regulating Valve, Filter Element, and Differential Pressure bypass valve. Gearbox Housing and Oil Pressure Sensing Port: Delivers signal via wet line to the Oil Pressure Gauge. Airframe External Oil Filter: 10 micron pleated paper, bypass type filter. Temperature Control Valve: Located near the Airframe External Oil Filter. Cool oil is sent back to the reservoir. At 71oC 160oF valve begins to open and allow hot oil to flow through the radiator type cooler. Radiator Type Cooler: Cooled by air from the squirrel cage fan, which is driven off of the tail rotor shaft. This same air is ducted and sent forward to cool the transmission oil and hydraulic fluid.

2.3.2. Identify how to determine the engine oil level and why it is important to know the exact level.

Sight gauge: Located on the Stbd. side of the Oil Tank. Dipstick / Filler Cap: Located on Port side of the Oil Tank. This is the only way to know the exact amount. This is important because if freewheeling unit oil seal fails, transmission oil under pressure will enter the accessory gear box (which is not under pressure) and cause and overfill condition.

2.3.3. State how engine oil temperature and pressure information is transmitted to the cockpit.

Temperature: 28 v DC oil temp bulb located in the oil tank. Pressure: Wet line from pressure sensing port (downstream of internal filter)

2.4. State why fuel metering is important for a turbo shaft engine by identifying the problems which can occur if fuel is not metered correctly.

Excessive temperature, Compressor Stall, Rich or Lean Blowout

2.4.1. Identify the type of fuel control used in the 250C–20J engine and state how it controls engine power.

Ng Fuel Control Unit: Ng gear train drives flyweights which meters fuel. During Start compressor discharge pressure is used to meter fuel.

7 Nf Governor: Varies bleed air control line air pressure which goes to the Ng Fuel Control which in turn varies fuel to the engine.

2.4.2. Described the function of the power turbine governor by stating the engine performance parameters sensed and its connection to fuel control.

Twist Grip Must be full open and Nf gear train up to speed. Nf gear train drives fly weights which varies compressor discharge air pressure which is set to the FCU and used to meter fuel.

2.4.3. State what components control the power turbine governor and gas producer fuel control.

Twist Grip Position, compressor Discharge Pressure, Ng speed, Nf speed, Nf Governor RPM Beep switch which controls the linear actuator which controls the Nf governor.

2.4.4. State how the twist grip functions.

3 positions: Fuel Off, Flight Idle and Fuel Off

2.4.5. Described the droop compensation system by identifying the cockpit controls, engine controls, and its effect on power output. Collective position controls Nf governor throttle lever position which changes flyweight tension which changes compressor air pressure signal which controls the fuel control unit. This allows the fuel control unit to anticipate power changes based on collective position.

8 CHAPTER 3 FUEL SYSTEM

3.1. State the major components of the fuel system and their functions.

1. Fuel Cell – (91 gal new B & all C; 76 gal Old B) 2. 2 fuel boost pumps – Prevents vaporization in the fuel lines above 6000 ft PA by pressurizing the line at 4-30 psi. 3. fuel quantity measuring system – 2 floats, 1 gauge calibrated in Gal. Fuel Low Warning System: 20 Gal fuel low switch and Caution light 4. Pressure Indicating system – Each fuel boost pump has a 3.5 psi pressure switch and they both are connected to Caution light. A fuel pressure transducers located between fuel cell and fuel shutoff valve and sends a signal to the fuel pressure gauge. 5. Shutoff Valve – 28 v dc shutoff valve with a thermal relief feature located just above fuel cap in main fuel supply line. If DC power is lost it will stay in the last energize position. 6. Airframe fuel filter – Impending Bypass function, trips caution light. 7. Engine Fuel pump / Filter – Low pressure (4-30 psi) bypassable filter element filters fuel and delivers it to the pump which pressurizes the fuel (650-750 psi) and since it to the Ng fuel control unit.

3.1.1. Identify the characteristics of the fuel cell.

Single bladder-type fuel cell, located below and aft of the passenger seat, is crash resistant, but not self-sealing.

Capacity 91 Gal (late model B’s and all C’s) 76 Gal (early B’s BUNO 161XXX)

3.1.2. Identify the methods used to re-fuel the fuel cell.

Early B’s – Gravity fueling only Late C B’s and all C’s – Gravity and Pressure fueling. (125 psi max press)

3.1.3. State how the fuel boost pumps and their components work to supply fuel to the engine and information to the pilot.

Two pumps (4-30 psi), connected in parallel, energized anytime the 28 vdc boss is energized. Each plant has a 3.5 psi pressure switch to indicate failure to a common FUEL PUMP Caution Light.

3.1.4. Identify the various fuel pressure limitations and the associated operational restrictions.

Allowable fuel boost pump pressure limits are 4-30 psi. Fuel pump light comes on when either pump pressure falls below 3.5 psi. If one fuel boost pump fails the fuel pressure gauge will still read 4 – 30 psi, if both pumps fail the gauge will read 0 psi. If either (or both) fuel pump(s) fail you must descend below 6000’ PA. Minimum usable fuel becomes 20 gal.

3.1.4.1. Identify the various pages, switches, etc., that relay pressure information, along with their limitations.

9 FUEL PRESSURE GAUGE: Scale is 0-30 psi. Normal range is 4-30 psi. A fuel pressure transducer sends a signal to the gauge. NOTE: VHF transmission may cause pressure gauge to fluctuate if helo has a solid state fuel pressure transducer due to EMI.

FUEL BOOST PUMP PRESSURE SWITCH: 1 on each boost pump; activates fuel pump caution light if pump pressure falls below 3.5 psi.

AIR FRAME FULTER CARTUIN LIGHT: When differential pressure is sensed across the airframe fuel filter the caution light comes on to warn the pilot of an impending bypass.

LOW FUEL CAUTION LIGHT: Comes on when fuel is approximately 20 gal.

3.1.5. Identify the components and characteristics of the fuel quantity measuring system.

2 float type sensors provide inputs to a single fuel quantity indicator (calibrated in gallons, 0-100) 1 float typw switch activates fuel low caution light when fuel drops to approximately 20 gllons.

3.1.6. State how the pressure indicating system works to relay fuel pressure information to the pilot.

Fuel pressure transducer, located between boost pumps and fuel shutoff valve, sends a signal to cockpit to indicate fuel boost pressure. (28 vdc)

3.1.7. State how shut-off power works to permit or restrict fuel flow.

28 vdc system. Shutoff valve is located in main fuel supply line in fuel compartment just above fuel filler cap. Controlled by cockpit switch. When DC power is lost the valve stays in the last energized position.

3.1.8. State the location of the airframe fuel filter and identify the functions of its components.

Located on STBD side, aft of the forward firewall and consists of: Replaceable filter element, drain valve, bypass valve, impending bypass switch, manual test button.

3.1.9. State the operational characteristics of the engine fuel pump/ filter.

Delivers filtered fuel at 650-750 psi to the fuel control unit. NOTE: Fuel enters pump housing and passes through a low pressure filter (4-30 psi) and then to the gear type pump which pressurizes the fuel to 650-750 psi and delivers it to the FCU. The internal filter will bypass if clogged but there is no indication until maintenance inspects the filter. The FCU meters the proper amount of fuel and sends it to the fuel nozzle. Any excess fuel hot needed by the fuel control unit is returned to the engine driven fuel pump.

10 CHAPTER 4 TH-57B/C POWER TRAIN

4.1. Identify the components of the TH–57 power train.

1. Transmission 2. Engine drive shaft (Barbell Shaft) 3. Freewheeling Unit 4. Forward short shaft 5. Oil cooler fan driveshaft 6. Aft short shaft 7. 5 tail rotor drive shafts 8. Tail rotor gear box

4.1.1. State the type of lubrication system utilized in the main transmission.

Wet sump, pressure lubrication system

4.1.2. Identify the components of the main transmission oil system.

1. Wet sump: Lower portion of transmission case (5 quart capacity) 2. Oil Pump: Driven by transmission accessory drive (also drives Hydraulic power pack: pump and Nr tach Gen) mounted internally 3. Oil filter Head Assembly: a) High Temp Sensor: 110oC turns on caution Light b) Temp Bulb: Sends signal to temp gauge on instrument panel c) Oil Bypass Valve: Bypasses filter when clogged d) Oil Monitor: Magnetic screen type chip detector (not part of Chip detector Caution Light system.) 4. Drain Valve: Self explanatory hopefully 5. Oil Cooler: Radiator type cooler with temperature bypass feature when oil is cold. Cooling air supplied by the squirrel cage fan. 6. 2 pressure jets: Sprays oil on transmission gears and bearings. A tee fitting between the two jets supplies oil to the freewheeling unit and sends oil to the wet line pressure gauge on the instrument panel and the low pressure switch. 7. Oil pressure regulator: Allows maintenance ot adjust oil pressure

4.1.2.1. State the function of each component of the transmission oil system.

See 4.1.2

4.1.3. State the location of the barbell shaft.

Between transmission and freewheeling unit

4.1.3.1. State function of the barbell shaft.

Transmits power from freewheeling unit to transmission and has a flexible splined coupling on each end which gives it its barbell shape. The flexibility if these couplings allows for momentary misalignment of the shaft, cause by movement of the transmission durring flight.

4.1.4. State function and location of the freewheeling unit. (AND) 4.1.4.1. State the source of lubrication for the freewheeling unit.

11 Mounted to the front of the accessory gear box. The freewheeling unit housing serves a an oil sump and holds the magnetic chip detector (not electrical and not part of the chip light system). The freewheeling unit is composed of an outer race which is splined to Nf turbine output gear and an Inner race which connects to the Barbell shaft and Tail Rotor Drive Shaft. Between the tow races is the sprag clutch which disconnects the power train from a failed Engine.

4.1.5. Identify the components of the tail rotor drive shaft.

1. Steel Forward Short shaft 2. Steel Oil Cooler Fan Driveshaft 3. Aluminum Aft Short Shaft 4. 5 Aluminum Tail Rotor Drive shafts 5. Tail rotor Gear Box (Magnesium which changes direction of drive by 90 degrees and provides gear reduction from 6000 rpm to 2554 RPM (2.35:1)

4.1.5.1. State the function of the tail rotor drive shaft.

Serves as part of the tail rotor drive train and as it turns it also turns the squirrel cage fan which in turn cools engine oil, transmission oil and hydraulic oil.

4.1.6. State the two functions of the tail rotor gear box.

Changes direction of drive by 90 degrees and provides gear reduction from 6000 rpm to 2554 RPM (2.35:1)

4.1.7. Identify the type of lubrication system utilized in the tail rotor gear box.

Self contained splash type. 3/8 pint capacity, contains an electrical ship detector wired to a T/R CHIP caution light.

12 CHAPTER 5 MAIN ROTOR, TAIL ROTOR, AND FLIGHT CONTROL SYSTEM

5.1. State the major components of the main rotor system.

1. Splined trunnion 2. Yoke 3. Tension torsion-straps 4. Blade grips 5. Pitch change horns 6. Flap restraint assembly (kit) 7. Main rotor blades

5.1.1. Described the design of the main rotor system.

Two bladed, semi-rigid, flapping type with an underslung hub

5.1.2. State the purpose of the splined trunnion.

Provides the flapping axis for the main rotor and a mounting point for the yoke assembly.

5.1.3. State the type of mounting used between the trunnion and mast of the main rotor.

Splined mounted to the mast

5.1.4. State the purpose of preconing.

2.25o of preconing, helps relieve bending stress of he yoke, blade grips, and root of main Rotor Blades.

5.1.5. State the purpose of the tension torsion strap.

Connects blade grips to the yoke assembly and absorb centrifugal forces while allowing for twist to permit pitch change action.

5.1.6. Described the purpose of the latch bolt.

The horizontal latch bolt connects the blade grips to the tension torsion strap.

5.1.7. Described the method used to ensure high rotational inertia for autorotations.

During construction weights are added to the tip and mid-span of each main rotor blade.

5.1.8. State the purpose of the flap restraint kit. Limits the amount of flapping at low rotor RPM (below 25-32% Nr)

5.1.8.1. Described the operation of the flap restraint kept.

At low RPM, springs hold the restraint arms in place and prevent excessive flapping. As RPM increases, flyweights overcome spring tension (at 25-32% Nr) and restraint arms move allowing full flapping motion.

5.1.9. State the need for a rotor break.

13 Provides a means of rapidly decelerating the rotor after engine shutdown for personnel and aircraft safety.

5.1.9.1. Differentiate between the TH-57 hydraulic system and the rotor break hydraulic system.

Totally independent of each other. Rotor brake uses a hand operated master cylinder (100-120 psi).

5.1.9.2. State type and location of the gauge associated with the TH-57C rotor break.

Direct reading gauge located overhead between the pilots.

5.2. State the function of tail rotor system.

To compensate for the torque effect of the main rotor and to allow control of the helo about he yaw axis.

5.2.1. Describe the tail rotor system.

2 stainless steel blades, semi-rigid, flapping type system (5’-5” in Dia)

5.2.2. Identify the major components of the tail rotor system.

1. Stainless Steel Rotor Blades 2. Pitch Change Horns 3. Cross Head 4. Control Tube 5. Balance Wheel 6. Static Stop 7. Yoke Assembly

5.2.2.1. Describe the construction of the tail rotor blades.

Aluminum honeycomb covered by stainless steel skin. Stainless steel doublers added at the foot for added strength. Stainless steel leading edge abrasion strips. Root end has an aluminum alloy retention block which houses two spherical bearings to mount the blade to the yoke.

5.2.3. State the purpose of the blade doublers.

Added Strength

5.2.4. State the purpose of the balance wheel.

Weights are added to the holes in the wheel to Dynamically Balance the tail rotor.

5.2.5. Identify the component which limits tail rotor flapping.

Static Stop mounted to the gearbox output shaft

5.3. Identify the type of flight control systems used on the TH-57.

Conventional type mechanical flight control system

14 5.3.1. Identify the three flight control systems.

Cyclic, collective, anti-torque pedals

5.3.2. State the purpose of the cyclic control.

A cyclic control input will result in the rotor disc tilting and the aircraft moving in the direction of the control input.

5.3.2.1. Describe the sequence of events with a cyclic input.

As the cyclic moves it causes the pivot support to move. Pilot and copilot pivot supports are inter-connected by a torque tube. A yoke assembly connects to the top of each pivot support and transmits the cyclic input to the mixing lever located at the base of the control column. The mixing lever transmits the mixed (fore / aft with left/ right) cyclic input to the cyclic hydraulic servos ant the stationary swashplate. The hydraulic servos boost the mechanical input. The stationary swashplate tilts and there fore transmits the cyclic input to the rotating swashplate which tilts and provides input to the pitch change horns via the pitch change rods. NOTE: The stationary swashplate is mounted on the uniball which his what allows the stationary swashplate to tilt in any direction.

5.3.3. Identify the type of boost used with the flight controls.

Hydraulic Boost

5.3.4. State the purpose of the collective control.

The collective changes the pitch on both main rotor blades equally and in the same direction to give vertical control of the helo.

5.3.4.1. Described the sequence of events with a collective input.

The collective is connected to a jackstaff which is where the friction adjuster is mounted. Collective control inputs are transmitted through a lever assembly and control tube up to the hydraulic servo and the collective lever (item with “Top” on it in preflight). As the pilot pulls up on the collective, the collective lever is puled down. This will raise the pivot sleeve and uniball assembly and therefore raise the stationary and rotating swashplates which will increase the pitch equally on both main rotor blades via the pitch change rods and pitch change horns.

5.3.5. State the purpose of the anti-torque tail rotor.

To compensate for the torque effect of the main rotor and to allow control of the helo about the vertical axis.

5.3.6. Describe the composition of the tail rotor blades.

Aluminum honeycomb covered by stainless steel skin. Stainless steel doublers added at the foot for added strength. Stainless steel leading edge abrasion strips. Root end has an aluminum alloy retention block which houses two spherical bearings to mount the blade to the yoke.

15 5.3.7. Recognize components of the tail rotor system and their function.

Stainless steel rotor blades, Pitch Change Horns transmit pitch changes to the Blades from the pitch change links, Cross head transmits pitch changes from control tube to the pitch change links, Control tube transmits pitch changes through the gear box to the cross head, Balance Wheel Static Stop, Yoke Assembly.

5.3.8. Described the sequence of events with a tail rotor input.

Right pedal input is transmitted via a push-pull tube to the pitch change mechanism mounted on the tail rotor gearbox. The pitch-change mechanism consists of a Lever, Control Tube, Crosshead, and Pitch Change Links. Right pedal causes the control tube to extend and pushes the crosshead away from the yoke assembly and right pedal decreases pitch……. Left pedal increases pitch.

16 CHAPTER 6 HYDRAULIC SYSTEM

6.1. State the basic purpose of the TH-57 hydraulic system and identify which control systems receive hydraulic assistance.

Reduces pilot work loads by reducing cyclic and collective control pressures and feedback generated by the main rotor system.

6.2. State the major components of the hydraulic system and their functions.

1. Power Pack: Consists of a 1 pint cooling reservoir, pump, regulating valve (600 ± 50 psi), and a mount for the Nr tach generator. It is driven by the transmission accessory drive-shaft. 2. Filter: Micronic metal filter with No bypass function. If clogged the red pip pin indicator will pop up. 3. Pressure Switch: Closes and completes the “HYDRAULIC PRESSURE” caution lithe circuit when the pressure falls bellow 300 psi. The switch opens, turning off the caution light when pressure rises above 400 psi. 4. Solenoid Valve: (Fail Safe Valve): Spring loaded to the on / open position allowing hydraulic fluid to go to the three servos. Requires 28 vdc to move the valve to the bypass position which direct fluid back to the reservoir instead of going to the servos. 5. Servos: 2 outboard servos are the cyclic servos, middle is the collective. Used to boost the control inputs.

6.2.1. Identify the components of the power pack and their functions.

1. Reservoir: Finned for cooling, 1 pint capacity 2. Pump Section: Pressurizes hydraulic fluid 3. Pressure Regulator: Regulates hydraulic fluid pressure (600 ± 50 psi) 4. Nr mounting pad

6.2.1.1. State the capacity and pressure requirements of the system.

Reservoir Capacity 1 pint (MIL-H-83282 or MIL-H-5606) System Capacity 2.25 pints System Pressure 600 ± 50 psi

6.2.2. Identify the components of the filter and their functions.

Head, Micronic Metal Filter Element, Body: A Red warning button on top of the filter pops when filter is clogged. O cockpit indication other than loss of hydraulic system and “Hydraulic Pressure” caution light. There is No bypass Function so when filter clogs the hydraulic system becomes inoperative.

6.2.2.1. State the primary indication of filter stoppage and corrective action.

In flight Indications 1. HYDRAUIC PRESSURE light 2. Increased force required for control movement 3. Feedback in control Procedures

17 1. Airspeed Adjust (to obtain most comfortable control movement level) 2. HYDRAULIC BOOST switch Check ON 3. HYD BOOST circuit breaker OUT If system is restored 4. Land as soon as possible If system is not restored: 5. HYD BOOST circuit breaker IN 6. HYDRAULIC BOOST switch OFF 7. © FORCE TRIM (FT) ON 8. © AFCS STAB ON 9. © AFCS ALT OFF 10. © Land as soon as practicable

During Preflight Notify Maintenance

6.2.3. Identify the function of the pressure switch.

Continuously monitors hydraulic system pressure “downstream” from the filter.

6.2.3.1. State how the hydraulic pressure light operates with the pressure switch.

Closes and completes the “HYDRAULIC PRESSURE” caution lithe circuit when the pressure falls bellow 300 psi. The switch opens, turning off the caution light when pressure rises above 400 psi.

6.2.4. Identify the functions of the solenoid.

Spring loaded to the OPEN/on position and allows the hydraulic servos to be bypassed by applying 28 vdc to the solenoid.

6.2.4.1. Describe the operation of the solenoid valve relative to the hydraulic switch.

Switch Position Solenoid Valve ON No power to solenoid, spring loaded OPEN position, direct hydraulic fluid to the servos

OFF 28 vdc applied to solenoid which overcomes the spring pressure and moves the valve to the BYPASS position which directs hydraulic fluid to the reservoir instead of going to the servos.

6.2.4.2. State how the hydraulic switch circuit is protected.

5 amp circuit breaker (outboard of igniter CB). If hydraulic boost switch shorts out the hydraulic solenoid will go to the bypass position. To remove the 28 vdc from the solenoid pull the circuit breaker which will restore the hydraulic system.

6.2.5. Name in the three major servo valves and state their functions.

18 Sequence Valve: Traps fluid in servo in case system pressure is lost. This allows control movement and absorbs rotor system feedback. (Maintains irreversibility) Pilot Valve: Heart of the servo, it takes control inputs (from pilot) and ports fluid to proper side of servo to allow for boosted control inputs. When the flight control is at the proper position, pilot valve centers and fluid flow stops in the servo. When system pressure is lost the sequence valve traps fluid in the servo and the pilot valve allows fluid to travel from one side of the servo to the other when flight control inputs are made. Although the flight control inputs are not boosted because there is no hydraulic pressure, the trapped fluid doe not provide rotor system feedback reduction. Differential Relief Valve: If heavy main rotor loads cause the back pressure to exceed system pressure the differential relief valve will open and allow pressure fluid to be directed to the return line and back to the reservoir.

6.2.5.1. State how a fluid return point is established.

600 ±50 psi hydraulic fluid enters the servo pressure port and is directed to the sequence valve. This pressure pushes down on the popet valve and spring (part of the sequence valve) which opens the return port for fluid to return to the reservoir. If pressure is lost, the spring pops up and closes the return port.

6.2.5.2. Described how the system dampens main rotor feedback.

When system pressure is lost the sequence valve traps fluid in the servo and the pilot valve allows fluid to travel from one side of the servo to the other when flight control inputs are made. Although the flight control inputs are not boosted because there is no hydraulic pressure, the trapped fluid doe not provide rotor system feedback reduction.

6.2.5.3. State the result of excessive system back pressure.

If heavy main rotor loads cause the back pressure to exceed system pressure the differential relief valve will open and allow pressure fluid to be directed to the return line and back to the reservoir.

6.2.5.4. State how flight control receive hydraulic boost.

The pilot valve directs 600 ±50 psi hydraulic fluid to one side of the servo. When the servo moves to the proper position the pilot valve centers and fluid flow is stopped.

6.2.5.5. State how the actuator works when hydraulic pressure is lost.

When system pressure is lost the sequence valve traps fluid in the servo and the pilot valve allows fluid to travel from one side of the servo to the other when flight control inputs are made. Although the flight control inputs are not boosted because there is no hydraulic pressure, the trapped fluid doe not provide rotor system feedback reduction.

19 CHAPTER 7 ENVIRONMENTAL CONTROL SYSTEM

7.1. Identify the sections of the ventilation and defog systems and state how their associated components work to provide ram air ventilation and defogging.

1. Grill: Keeps FOD out of system 2. Flapper Valve: Allows outside airflow to enter cockpit, controlled by bent control knob. 3. Axial Blower: Controlled by DEFOG Blower CB type switch located on overhead console. 4. Defroster Nozzle: Directs air from blower to windscreen. 5. Bent Control Knobs: Controls flapper valve, 2 on either side of center console.

Defog works: 1. With both Bent Control Knobs Pulled out all the way. 2. (BEST) Close vent Control Knobs and turn on A/C or heat with defog blower on.

7.2. Identify the major components of the vapor cycle air conditioner and state how they work independently and together to provide cabin air cooling.

1. Switch panel: Contains 2 switches (Air conditioner On / OFF and fan HI / LOW) and a temperature rheostat. 2. Compressor: Increases pressure of FREON gas. 3. Cools freon gas and changes it from a gas to a liquid. 4. Condenser Blower: Blows air through condenser to cool the freon. 5. Evaporator: Changes freon liquid to a gas thereby absorbing all the heat from the surrounding area. 6. Evaporator Blower: Blows cabin air through evaporator to cool it.

NOTE: System requires about 5 HP to operate.

7.3. Identify the major components of the bleed air heater system in the new B and C, and state how these components warm the cabin air.

1. Heater Silencer: Mixes bleed air and cabin air. 2. Control Valve: Valve overhead pilot’s seat governs the control air. 3. Regulator Valve: Controlled by control air. Allows bleed air to enter heating system. 4. Evaporator Fan: Circulates the air 5. Duct Temp Switch: Activates Caution light when Temp in duct becomes excessive.

7.3.1. Identify the operational procedures associated with monitoring system temperatures.

When Caution light Illuminates: (Heater Malfunction – Duct too Hot) Procedures: 1. CABIN HEAT VALVE OFF 2. AIR COND/FAN switch FAN 3. HI/FAN/LO switch HI If light extinguishes: 4. Continue flight. If light does not extinguish: 5. Land as soon as possible

20 CHAPTER 8 TH-57B ELECTRICAL SYSTEM

8.1. Identify the three sources of electrical power for the TH-57B.

Battery Generator External Power Unit

8.1.1. State the power rating of the battery.

24 volt,17 ampere hour

8.1.2. State the generator output.

Rated at 30 Volts, 150 amp Regulated to 28 volts, 105 amp

8.1.2.1. State the purpose of the voltage regulator used in the generator system.

Maintains constant generator voltage under varying loads.

8.1.2.2. State the purpose of the overvoltage sensing relay.

Trips the generator reset relay, disconnecting the generator from the circuit, when line voltage reaches 31±1 volts.

8.1.3. State the purpose of the reverse current relay.

1. Connects generator to common BUS only when proper voltage is obtained 2. Prevents current flow from the battery to the generator 3. Disconnects the generator from the common BUS when voltage drops below a safe level

8.1.4. Identify the component that allows the generator to function as a starter.

Starter relay

8.1.4.1. State when the igniters receive power.

When the field control relay is actuated

8.1.5. State the power requirements for an engine start when using external power.

28 vdc 400 amp

8.1.6. Recall the circumstances necessary to illuminate the battery caution lights.

54oC ±3o BATTERY TEMP caution light illuminates 60oC ±3o BATTERY HOT caution light illuminates

8.1.7. State the loadmeter and voltmeter indications that would signify a generator failure.

Load Meter 0 Voltmeter 22 to 24 volts

21 CHAPTER 9 TH-57C ELECTRICAL SYSTEM

9.1. Identify five sources of dc electrical power for the TH-57C.

1. Main Battery 2. Standby Battery 3. Main Generator 4. Standby Generator 5. External Power Unit

9.2. State the power rating of the standby battery.

22.5 volt. 1.8 ampere hour

9.3. State the purpose of the standby battery.

To supply to 22.5 v power to the pilot’s attitude gyro in the event of a complete loss of power to essential BUS #1.

9.4. State the power rating of the main battery.

24 v, 17 amp hour

9.5. State the purpose of the RCB.

Provides over current protection to the battery (250 amp for 10-20 sec, but will trip with a constant current of 125 amps.)

9.6. State the purpose of the RCB override circuit.

The RCB override is incorporated into the battery and starter switch circuit to prevent the RCB from taking power away from the starter before a complete engine start is accomplished.

9.7. State that main generator output.

28 v, 105 amps

9.8. Described the functions of the voltage regulator used in the main generator system.

Maintains constant generator voltage under varying loads.

9.9. State the purpose of the standby generator.

Provides a backup power source for essential BUS #1 in case the main generator fails.

9.10. State the standby generator output.

28 v 7.5 amps

9.11. State the power requirements for engine start utilizing external power.

28 volt 400 amps

22 9.12. Describe each function of the voltmeter panel selector switch.

Provides a means to display the voltages at each of the following locations: 1. Main Battery 2. Main Generator 3. STBY Generator 4. STBY Battery 5. ESS 1 BUS 6. ESS 2 BUS 7. NON ESS BUS 8. Flight Control System Inverter

9.13. Describe the alternating current system and identify its power sources.

There are two static inverters (Solid-state) which each take an input of 28 v DC and produce 400 Hz, 115 Volt and 26 v alternating current. The avionics inverter works with the avionics and flight control system yaw axis. The FCS inverter works only with the flight control system. The inverters are controlled by circuit breakers on the overhead console.

23 CHAPTER 10 THE MINISTAB SYSTEM

10.1. Identify the general characteristics and functions of the MINISTAB system.

A basic three axis system (pitch, roll, and yaw) with force trim designed to provide attitude retention and to smooth pilot input to the controls (rate dampening). It also provides altitude hold in cruise flight (above 40 kts).

10.1.1. Identify the controller buttons and their functions.

STAB: Engages and disengages the flight control system FT: Engages and disengages the flight control system ALT: Engages and disengages the altitude hold mode. TEST: Initiates a test of the MINISTAB system.

10.1.2. Identify the functions of the force trim and STAB buttons found on the cyclic.

Force Trim: When pushed ministab goes in standby and force trim gradient is disengaged. When released Resets ministab and force trim systems. STAB: Turns ministab on and off.

10.1.3. Describe the functions of the trim damper units.

The pitch and roll actuators are located at the top of the control tubes and provides input to the hydraulic servo pilot valves. Therefore they have a low output force. The pitch and roll computers provide a signal to the actuators which move about a neutral point (0.5 inch total movement NATOPS 2.12.5 pg 2- 28, 1.0 total movement Systems Workbook (CNATRA P-402 (rev. 7-98))) to maintain attitude. The yaw actuator is connected to the tail rotor control tube. Since the tail rotor flight controls do NOT have hydraulic boost. The tail rotor actuator is larger and has a temperature cutoff switch. The Trim Damper Units (TDUs) provide force trim and smoothes / dampens pilot input. Pitch and Roll FT switch is on the cyclic, Yaw FT is on anti-torque pedals

10.2. Identify the major components of the MINISTAB system and state how they function independently and together in the flight control system.

Junction box Pitch computer Roll computer Yaw computer Air Data computer

The junction box interconnects components of the system. Each axis computer contains a rate gyro, memory circuit (integrated rate), and integration cutoff (ICO) circuit. Within each computer, the rate gyro senses movement about its axis, allowing the computer to detect any deviation from the attitude stored in the memory circuit. The computer then sends a correction signal to the actuator which makes the appropriate control input. This is how aircraft attitude retention is provided. The air data computer contains transducer that sense airspeed and altitude changes and it also incorporates an airspeed trip switch. The airspeed trip switch tells the FCS when the helicopter is above 40 kts indicated airspeed. This switch works in conjunction with altitude hold and the yaw axis computer.

24 System Operation The pitch computer senses changes and makes corrections to maintain the aircraft fuselage in that particular pitch attitude by sending equal signals to both cyclic actuators. Now, to change to a new pitch attitude, all that has to be done is to move the cyclic fore or aft. A micro switch in the pitch TDU senses this longitudinal cyclic movement and activates the pitch computer ICO. Once you stop moving the controls, the computer waits for movement about its axis to stabilize below 1.5o per second for a period of 900 milliseconds. At that time the ICO is secured again. This delay lets the aircraft settle down, and then the computer maintains the new attitude. The roll axis system functions in the same way, except the computer only senses change in the roll axis. For roll inputs the cyclic actuators move simultaneously in opposite directions. To accomplish this, equal signals are sent to both cyclic actuators except one signal is reversed in polarity for roll input, left or right. For the yaw axis, the principal is the same. A rate in the yaw computer senses changes about its axis and makes inputs to the tail rotor controls to hold the last heading set. The yaw system has the micro switches located in the pedal assemblies and not in the TDU. To change heading, move the pedals to the new setting. The micro switch will ICO the computer and once the computer detects that the aircraft has settled down it will maintain the new heading.

10.3. Identify the operational functions and limitations of the MINISTAB system.

10.3.1. State how movement of the flight the controls operate with the various computers in the MINISTAB system to control flight.

The pitch computer senses changes and makes corrections to maintain the aircraft fuselage in that particular pitch attitude by sending equal signals to both cyclic actuators. Now, to change to a new pitch attitude, all that has to be done is to move the cyclic fore or aft. A micro switch in the pitch TDU senses this longitudinal cyclic movement and activates the pitch computer ICO. Once you stop moving the controls, the computer waits for movement about its axis to stabilize below 1.5o per second for a period of 900 milliseconds. At that time the ICO is secured again. This delay lets the aircraft settle down, and then the computer maintains the new attitude. The roll axis system functions in the same way, except the computer only senses change in the roll axis.

10.3.2. Identify the airspeed and attitude limitations associated with the MINISSTAB system.

With the ministab in operation Vne is reduced to 122 knts at density altitudes of 3000 and below. >3000’ DA will depend on gross weight and DA flown.