Fam 5 Discuss Items

Syllabus Notes a.The purpose of this block is to continue BAW while introducing additional basic maneuvers, emergency procedures, and autorotation skills. b.Emphasis CRM during all flights, especially during simulated emergency procedures. c.SMA shall fly off wing for at least one, but no more than two flights between C4102-C4203 Special Syllabus Requirements N/A

Discuss Items 1. Engine system (NATOPS, Systems)

A 250-C20J engine, manufactured by Rolls Royce, powers the helicopter. It is an internal combustion, gas turbine engine consisting of a six axial and one centrifugal stage compressor, a single combustion chamber, a two-stage gas producer turbine (Ng), and a two-stage power turbine(Nf). The engine weighs 157 pounds and is capable of developing 420 shaft horsepower (SHP) on a standard day. This has been limited to 317 SHP because of drive train limitations. Additionally, an accessory gearbox powers gear trains from the gas producer and power turbine.

Power is generated in the 250-C20J engine through the use of expanding gases passing over turbine blades. Ambient air enters the compressor inlet section and is compressed approximately 6.5 times its original ambient pressure and 500oF. The high-pressure air is then ducted rearward to the combustion chamber through a diffuser scroll and two air-transfer tubes. The combustion chamber regulates the engine air flow. Fuel is injected through a nozzle in the center rear section of the combustor and is ignited by capacitance-type ignitor plug used only during the starting cycle. Ignition is self-sustained after engine start. The hot combustion gases pass forward through the two-stage gas-producer turbine and the two-stage free-power turbine. The free-power turbine drives the power output shaft. After passing through the power turbine section, the gases are vented through two exhaust ducts. Thrust from the exhausted gases is negligible. The compressor assembly consists of a front support, rotors, case, and diffuser. The compressor front support has seven radial struts that direct the incoming air onto the compressor rotor. An anti-ice system is incorporated to protect both the struts and front support area from inlet icing.

Approximately 75 % of the air going to the combustion chamber is used for cooling. The remainder of the air is mixed with fuel from the fuel nozzle for burning. The combustion assembly consists of a combustion outer case and liner. The fuel nozzle and spark ignitor thread through the aft end of the outer case into the aft end of the combustion liner.

A burner drain valve is located in the lowest point of the combustion chamber to drain any fuel overboard that may collect in the chamber after engine shutdown.

The turbine assembly is made up of a two-stage gas-producer turbine (Ng) and a two-stage free-power turbine (Nf). These two components are not mechanically coupled, but are connected by a gas coupling in which the free-power turbine turns as the exhaust gases impinge on its blades. At 100% power, the gas-producer turbine (Ng) rotates at 51,989 rpm, and the free-power turbine (Nf) rotates at 33,956 rpm. Approximately two-thirds of the energy delivered to the gas-producer turbine (Ng) is used to drive the compressor, while the other one-third is used to power the free-power turbine.

The accessories driven by the gas-producer include:

S - Starter-Generator T - Tachometer-generator O - Oil Pump F - Fuel Control F - Fuel Pump S - Standby Generator (C)

The accessories driven by the free-power turbine include:

G - Governor (Nf) O - Output Shaft (Power) T - Tachometer-Generator (Nf) T – Torquemeter

The engine lubrication system is a circulating dry-sump with an external oil tank and oil cooler. This system is designed to furnish lubrication, scavenging, and cooling as needed for bearings, splines, and gears regardless of helicopter attitude or altitude. Lubrication is provided to all bearings in the compressor, gas-producer turbine, and power turbine, and to bearings and gear meshes in the power turbine gear train with the exception of the power output shaft bearing. The power output shaft bearings are lubricated by an oil mist. A spur gear oil pump assembly, consisting of one pressure element and four scavenge elements, is mounted within the engine accessory gearbox. Oil from the tank is delivered to the pressure pump that pumps oil through the oil filter and to various points of lubrication. The system pressure is adjusted to a maximum of 130 psi by the pressure-regulating valve. The oil system maintains this relatively high pressure in order to balance high axial gear thrust in the torquemeter. This high thrust value is necessary to minimize friction effects and provide accurate measurement of torque. The oil filter, filter bypass valve, and pressure regulating valve assembly are located in the upper right side of the accessory gearbox. A check valve is located in the oil filter outlet passage. Two chip detector elements, the scavenge oil outlet chip detector, and the accessory gearbox chip detector are located on the accessory gearbox section of the engine and illuminate the ENG CHIP caution light on the instrument panel should they detect metallic filings in the sumps. The cooler fan is mounted on the upper structure, aft of the aft firewall, and is driven by the tail rotor driveshaft. The squirrel-cage-type impeller is mounted on a flanged shaft that is mounted in bearing hangers. The fan shaft connects to the forward and aft short tail rotor driveshafts and is part of the tail rotor driveshaft system. The oil-cooling fan provides cooling air for the engine oil system, the transmission oil system, and the hydraulic system.

2. Engine Failures (NATOPS) Under operational conditions, the altitude-airspeed combination for a safe autorotative landing is dependent upon many variables such as pilot capabilities, density altitude, helicopter gross weight, proximity of a suitable landing area, and wind direction and velocity in relation to flightpath. This does not preclude operation in the shaded area of the height velocity diagram under emergency or pressing operation requirements. Immediately upon an engine failure, rotor rpm will decay and the nose of the helicopter will swing to the left. This is because of the loss in power and corresponding reduction in torque. Except in those instances when an engine failure is encountered in close proximity to the surface, it is mandatory that autorotation be established immediately lowering the collective pitch to minimum

Heading can be maintained by depressing the right pedal to decrease the tail rotor thrust. Autorotative rpm will vary with different ambient temperature, pressure altitude, increase in g loading, and gross weight conditions. High gross weights, increase g loads, and high altitudes and temperature will cause increased rpm that can be controlled by increasing collective pitch. Any increase of rotor rpm, other than specified for maximum glide, will result in a greater rate of descent; therefore, if time permits, adjusting the collective pitch lever to produce the desired rotor rpm will result in an extended glide. At an altitude of approximately 75 to 100 feet, a flare should be established by moving the cyclic stick aft with no change in collective pitch. This will decrease both airspeed and rate of descent and cause an increase in rotor rpm. The amount that the rotor rpm will increase is dependent upon gross weight and the rate that the flare is executed. An increase is desirable because more energy will be available to the main rotor when collective pitch is applied. From this condition of airspeed and low altitude, flare capability is limited and caution should be exercised to avoid striking the ground with the tail; the primary objective is to level the skids prior to ground contact. Initial collective reduction varies with altitude; from a 5-foot skid height, do not attempt collective reduction but use the available rotor energy and collective to cushion touchdown; above 5-foot skid heights, a partial reduction of collective will maintain rotor rpm until up collective is initiated to cushion touchdown.

Note: The best glide airspeed is 72 KIAS. The minimum rate of descent airspeed is 50 KIAS. Do not exceed 100 KIAS in sustained autorotation. If time and altitude permit, engine restart may be attempted. The decision to attempt a restart is the pilot’s responsibility and is dependent upon the pilot’s experience and operating altitude. All autorotative landings should be made into the wind to a suitable landing site.

Indications: Nr decrease Rapid settling Left yaw ROTOR LOW RPM caution light and audio ENGINE OUT caution light and audio

Should an engine failure occur at high airspeed and low altitude, a rapid loss of Nr accompanied by a severe nose-tucking tendency will occur.
 * 14.1.1 Engine Failure at High Airspeed and Low Altitude

Procedures:
 * 1.Cyclic – Immediately apply aft
 * 2.Autorotate

Warning: Rapid cyclic movement should be avoided to preclude mast bumping.

In the event of an engine failure in flight, a safe landing can be accomplished, provided that altitude and airspeed combination is within safe limits and altitude is sufficient to permit selection of a suitable landing area. Consideration should be given to an engine restart in flight.
 * 14.1.2 Engine Failure in Flight

Procedures:
 * 1.Autorotate
 * 2.Should harness – Lock

If time and altitude permit:
 * 3.Mayday – Transmit on guard
 * 4.Transponder – Emergency

3. *14.5 Engine Restart in Flight An engine flameout in flight would most likely result from a malfunction of the fuel control unit or fuel system. The decision to attempt an engine restart during flight is the pilot’s responsibility and is dependent upon the pilot’s experience and operating altitude. Consideration must be given to the cause of the failure prior to attempting restart. If attempting an engine restart, proceed as follows:

Procedures:
 * 1.Autorotate
 * 2.Fuel valve – Check ON
 * 3.Start – Engage

Caution: If Ng is allowed to fall below a minimum of 12 percent Ng, then close the twist grip and perform a normal start.

Note: Ng will not decrease below minimum starting speed within 10 seconds because of rotational inertia plus possible ram air effect. The twist grip can be left in the full open position since fuel flow during the start will be on the normal acceleration schedule.

If light-off occurs:
 * 4.Land as soon as possible

4. Engine chip clearing procedures (NATOPS 12-2) Procedures:

ENGINE INSTRUMENTS — Check engine instruments for secondary indications of impending failure. If these indications exist, LAND AS SOON AS POSSIBLE. If no secondary indications exist, proceed as follows:

- First chip light — Press clear chip; if eng chip light goes out, note the time and continue flight. If eng chip remains illuminated, LAND AS SOON AS POSSIBLE. - Second chip light — If within 30 minutes of the first, LAND AS SOON AS POSSIBLE. If more than 30 minutes have elapsed since the first light, press clear chip and proceed as with the first light. - Any subsequent chip light — If within 50 flight hours of the first, LAND AS SOON AS POSSIBLE and make no attempt to clear chip. All chip lights shall be documented on a VIDs/MAF.

Chip detector system:

The Chip Detector System incorporates a way to clear nuisance chips and continuously monitors the integrity of all chip detector circuits. The system goes through a self--test each time electrical power is turned on. A CLEAR CHIP switch is located adjacent to the caution--warning panel to allow the pilot to manually attempt to clear all chips, which will extinguish the corresponding chip light if successful. The Power Unit is located on the aft side of the pilot’s seat back and is easily accessible by the pilot during preflight and post flight checks and by maintenance personnel. Chip detector elements are installed in the drain plugs of the transmission sump, engine sump, freewheeling unit, and the tail rotor gearbox. These plugs provide a means of detecting metal particles in the oil. Filings in the oil bridge the gap across the detector plug and complete an electrical circuit to the cockpit that illuminates the appropriate caution light. The caution lights are marked ENG CHIP, TRANS CHIP, and T/R CHIP, and are located on the caution panel. In the case of a chip light, the CLEAR CHIP indicator switch will also illuminate. The chip lights are protected by the CAUTION LT circuit breaker. The aircraft chip detecting system is equipped with a continuity sensor that provides an automatic continuity check. When the battery switch is closed, the chip detector caution lights will illuminate for approximately 5 seconds confirming continuity of the electrical circuits for these devices. Interruption of the electrical power in excess of 2 seconds or activation of the caution panel test switch for 2 seconds or longer will recycle the circuit continuity check.

5. *14.9 Compressor Stall

Indications: Popping or rumbling noise Vibrations Rapid rise in TOT Ng fluctuation Loss of power

Warning: Be prepared for a complete power loss.

Procedures:


 * 1.Collective – Reduce

Note: Slight power (collective) reduction will often eliminate compressor stalls.


 * 2.ENG Anti-Ice Switch – OFF
 * 3.Cabin Heat Valve – OFF
 * 4.Land as soon as possible

Warning: When accelerating the rotor system during the initial rotor engagement or after a full autorotation, exceeding 40 percent torque may induce compressor stall or engine chugging.

Note: Mild compressor stalls may occur that will allow powered flight if TOT is within operating limits

The TH-57 power plant is equipped with a bleed valve. "This valve bleeds 5th stage pressure at low pressure ratios to unload the compressor in order to prevent compressor stall and surge. The compressor bleed air system is an entirely automatic system." (systems)

After power off maneuvers: Caution - To prevent compressor stall, stabilize Ng at flight idle for a minimum of 15 seconds with the collective full down. (NATOPS)

6. Rotor Drop (NATOPS 11-11) Droop is a term used to denote a change in power turbine speed (Nf) and rotor speed that occurs with a demand for increased power with the governor at a constant speed setting. Droop may be further categorized as either transient or steady state. Transient droop is the momentary change on power turbine speed and rotor speed resulting from increased power demand, and it is compensated for by the power turbine governor control. Steady-State droop is the decrease in power turbine speed and rotor speed the results from an increased power demand when the engine is already operating at maximum gas producer speed. This condition should be avoided.

7. CRM/Situational awareness -Self explanatory if not look it up.