The fuel system for gas turbine and reciprocating engines is the same. It delivers to the engine fuel metering system a uniform flow of clean fuel at the proper pressure and in the necessary quantity to operate the engine. Despite widely varying atmospheric conditions, the fuel supply must be adequate and continuous to meet the demands of the engine during flight.
The fuel system is one of the more complex aspects of a gas turbine engine. The variety of methods used to meet turbine engine fuel requirements makes reciprocating engine carburation seem simple by comparison.
It must be possible to increase or decrease the power at will to obtain the thrust required for any operating condition. In turbine-powered aircraft, this control is provided by varying the flow of fuel to the combustion chambers. However, turboprop and turboshaft aircraft also use variable-pitch propellers or helicopter rotors; thus, the selection of thrust is shared by two controllable variables, fuel flow and propeller blade or rotor blade pitch.
The fuel supply must be adjusted automatically to correct for changes in ambient temperature or pressure. An excessive fuel supply in relation to mass airflow through the engine can cause the limiting temperature of the turbine blades to be exceeded. It can produce compressor stall and surge.
The fuel system must deliver fuel to the combustion chambers in the right quantity and in the right condition for satisfactory combustion. Fuel nozzles form part of the system and atomize or vaporize the fuel so that it tiff ignite and burn efficiently. The fuel system must also supply fuel so that the engine can easily be started on the ground and in the air. This means fuel must be injected into the combustion chambers in a combustible condition when the engine is being turned over slowly by the starting system. Combustion must be sustained while the engine is accelerating to its normal running speed.
Another critical condition to which the fuel system must respond occurs during a quick acceleration. When the engine is accelerated, energy must be furnished to the turbine in excess of that necessary to maintain a constant RPM. However, if the fuel flow increases too rapidly, an overrich mixture may be produced causing a surge.
Turbojet, turbofan, turboprop, and turboshaft engines are equipped with a fuel control unit which automatically satisfies the requirements of the engine. Although the basic requirements apply generally to all gas turbine engines, the way in which individual fuel controls meet these needs cannot be conveniently generalized. Each fuel control manufacturer has its own way of meeting the engine demands.
Main fuel pumps deliver a continuous supply of fuel at the proper pressure during operation of the aircraft engine. Engine-driven fuel pumps must be able to deliver the maximum flow needed at high pressure to obtain satisfactory nozzle spray and accurate fuel regulation.
Fuel pumps for turbojet engines are generally positive displacement gear or piston types. The term "positive displacement" means that the gear or piston will supply a fixed quantity of fuel to the engine for every revolution of the pump gears or for each stroke of the piston.
These fuel pumps may be divided into two distinct system categories constant displacement and variable displacement. Their use depends on the system used to regulate the flow of fuel. This maybe a pressure relief valve (barometric unit) for constant displacement (gear) pump or a method for regulating pump output in the variable displacement (piston) pumps.
Gear-type pumps have approximately straight-line flow characteristics, whereas fuel requirements fluctuate with flight or ambient air conditions. A pump of adequate capacity at all engine operating conditions will have excess capacity over most of the range of operation. This is the characteristic which requires the use of a pressure relief valve for disposing of excess fuel. The impeller, which is driven at a greater speed than the high-pressure elements, increases the fuel pressure from 15 to 45 psi, depending on engine speed.
This fuel is discharged from the boost element (impeller) to the two high-pressure gear elements. Each of these elements discharges fuel through a check valve to a common discharge port. High-pressure elements deliver approximately 51 gallons per minute at a discharge pressure of 850 psig.
Shear sections are incorporated in the drive systems of each element. If one element fails, the other element continues to operate. Check valves prevent circulation through the inoperative element. One element can supply enough fuel to maintain moderate aircraft speeds.
A relief valve is incorporated in the discharge port of the pump. This valve opens at approximately 900 psi and is capable of bypassing the total flow at 960 psi. Excess fuel is recirculated. The bypass fuel is routed to the inlet side of the two high-pressure elements.
The variable-displacement pump system differs from the constant-displacement pump system. Pump displacement is changed to meet varying fuel flow requirements; that is, the amount of fuel discharged from the pump; of variable flow, the applicable fuel control unit can automatically and accurately regulate the pump pressure and delivery to the engine.
Where variable-displacement pumps are installed, two similar pumps are provided and connected in parallel. Each pump can carry the load if the other fails during normal parallel operations. At times one pump may be insufficient to meet power requirements. Pump duplication increases safety in operation, especially during takeoff and landing.
The positive-displacement, variable-stroke-type pump incorporates a rotor, piston, maximum speed governor, and relief valve mechanism.
All gas turbine engines have several fuel filters at various points along the system. It is common practice to install at least one filter before the fuel pump and one on the high-pressure side of the pump. In most cases the filter will incorporate a relief valve set to open at a specified pressure differential to provide a bypass for fuel when filter contamination becomes excessive.
Paper Cartridge. This filter is usually used on the low-pressure side of the pump (Figure 4-1). It incorporates a replaceable paper filter element which is capable of filtering out particles larger than 50 to 100 microns (the size of a human hair). (One micron equals 0.000039 inch, or 25,400 microns equal one inch.) The cartridge protects the fuel pump from damage due to fuel contamination.
Screen Disc. This disc (Figure 4-2) is located on the outlet side of the pump. This filter is composed of a stack of removable fine wire-mesh screen discs which must be disassembled and cleaned periodically in an approved solvent.
Screen. This screen is generally used as a low-pressure fuel filter. Some of these filter screens are constructed of sinter-bonded stainless steel wire cloth and are capable of filtering out particles larger than 40 microns (Figure 4-3).
In addition to the mainline filters, most fuel systems will incorporate several other filtering elements. They may be located in the fuel tank, fuel control, fuel nozzles, and any other place deemed desirable by the designer.
The pressurizing and drain valve prevents flow to the fuel nozzles until sufficient pressure is reached in the main fuel control. Once pressure is attained, the servo assemblies compute the fuel-flow schedules. It also drains the fuel manifold at engine shutdown to prevent post-shutdown fires. But it keeps the upstream portion of the sure and delivery to the engine. system primed to permit faster starts.
Flow Divider. The pressurizing and dump valve used on some engines has a somewhat different function. In addition to the draining or dumping function, this unit also serves as a flow divider. At the beginning of an engine start, the fuel control supplies a pressure signal to the pressurizing and dump valve. This causes the valve to close the manifold drain and open a passage for fuel flow to the engine. On engine shutdown, fuel flow is cut off immediately by a valve in the fuel control. The pressure signal drops, the dump valve opens, and fuel drains from the manifold. The flow divider allows fuel to flow to the primary and secondary manifolds depending on fuel pressure. (See Figure 4-3.)
Pressure-Operated Valve. Some manufacturers, such as Allison and Lycoming, install a pressure-operated valve in the combustion chamber section. When the pressure in the burners drops below a specified minimum, usually a few pounds per square inch, this valve will open and drain any residual fuel remaining after a false start or normal shutdown. (Refer to Figure 4-4 to see where this drain valve fits into the system.)
Gas turbine engine fuel systems are very susceptible to the formation of ice in the fuel filters. When the fuel in the aircraft fuel tanks cools to 32°F or below, residual water in the fuel tends to freeze when it contacts the filter screen.
A fuel heater operates as a heat exchanger to warm the fuel. The heater can use engine bleed air, an air-to-liquid exchanger, or an engine lubricating oil, a liquid-to-liquid exchanger, as a source of heat.
A fuel heater protects the engine fuel system from ice formation. However, should ice form, the heater can also be used to thaw ice on the fuel screen.
In some installations the fuel filter is fitted with a pressure-drop warning switch which illuminates a warning light on the cockpit instrument panel. If ice begins to collect on the filter surface, the pressure across the filter will slowly decrease. When the pressure reaches a predetermined value, the warning light flashes on.
Fuel deicing systems are designed to be used intermittently. The system may be controlled manually by a switch in the cockpit or automatically using a thermostatic sensing element in the fuel heater to open or close the air or oil shutoff valve. An automatic fuel heater is illustrated in Figure 4-5.
On most gas turbine engines, fuel is introduced into the combustion chamber through a fuel nozzle. This nozzle creates a highly atomized, accurately shaped spray of fuel for rapid mixing and combustion with the primary airstream under varying conditions of fuel and airflow. Most engines use either the single (simplex) or the dual (duplex) nozzle.
Simplex Nozzle Figure 4-6 (A) illustrates a typical simplex nozzle. Its chief disadvantage is that it is unable to provide a satisfactory spray pattern in bigger engines because of the large changes in fuel pressures and airflows.
Duplex Nozzle. The chief advantage of the duplex nozzle is its ability to provide good fuel atomization and proper spray pattern at all rates of fuel delivery and airflow. At starting and low RPM and at low airflow, the spray angle needs to be fairly wide to increase the chances of ignition and good mixing of fuel and air. At higher RPM and airflow, a narrow pattern is required to keep the flame of combustion away from the walls of the combustion chamber (Figure 4-6[B]). The small fuel flow used in idling is broken up into a fine spray after being forced through a small outlet formed by the primary holes. The secondary holes are larger but still provide a fine spray at higher RPM because of the higher fuel pressure.
For the duplex nozzle to function, there must be a device to separate the fuel into low-(primary) and high( secondary) pressure supplies. This flow divider maybe incorporated in each nozzle, as with the single-entry duplex type (Figure 4-6[C]), or a single-flow divider may be used with the entire system (Figure 4-6[D]).
Single-entry duplex nozzles incorporating an internal flow divider require only a single fuel manifold (Figure 4-7). Dual-entry fuel nozzles require a double fuel manifold. Some dual fuel manifolds may not be apparent as such.
The flow divider, whether self-contained in each nozzle or installed in the manifold, is usually a spring-loaded valve set to open at a specific fuel pressure. When the pressure is below this value, the flow divider directs fuel to the primary manifold or nozzle orifice. Pressures above this value cause the valve to open, and fuel is allowed to flow in both manifolds or nozzle orifices.
Most modern nozzles have their passages drilled at an angle. The fuel is discharged with a swirling motion to provide low axial air velocity and high flame speed. In addition, an air shroud surrounding the nozzle cools the nozzle tip. It also improves combustion by retarding the accumulation of carbon deposits on the face. The shroud also provides some air for combustion and helps to contain the flame in the center of the liner (Figure 4-8).
Extreme care must be exercised when cleaning, repairing, or handling the nozzles. Even fingerprints on the metering parts may produce a fuel flow which is out of tolerance.
The engine fuel shutoff valve is installed in the main fuel supply line or tank outlet to the engine. It is controlled from the pilot's compartment. A fuel shutoff valve is usually installed between the fuel control unit and the fuel nozzles. When the throttle is placed in the closed position, this ensures positive shutoff of fuel to the engine.
The fuel system consists of the fuel boost pump, fuel filter, hydromechanical unit (HMU), and overspeed and drain valve. Integral to the HMU are the high-pressure vane pump, variable geometry (VG) actuator, and compressor inlet temperature sensor. These fuel system components are mounted on the accessory gearbox (AGB) and oriented as shown in Figure 4-9. At a 100 percent gas generator speed (Ng) of 44,700 RPM, the fuel boost pump is driven at 10,678 RPM and the HMU at 9947 RPM. This fuel-oil heat exchanger is located in the metered fuel flow path to the engine between the HMU discharge and the overspeed and drain valve inlet. Except for one external line between the HMU discharge (metered fuel) and the fuel-oil heat exchanger inlet passage, all the fuel transfer is through cored passages within AGB. This includes the boost pump inlet flange (engine-aircraft interface) to the overspeed and drain valve where the fuel discharges into the primer and the main manifolds.
A schematic diagram of the fuel system component arrangement is shown in Figure 4-10. An aircraft fuel system is depicted to show the overall aircraft-engine fuel-handling systems.
The primary requirements of the fuel system are--
These requirements must be met over the full engine operating envelope and environment.
The fuel boost pump is mounted on the AGB forward side (Figure 4-11). It is designed--
At 100 percent Ng (44,700 RPM) the fuel boost pump operates at 10,678 RPM. Pump discharge pressure ranges from 45 to 90 psi at maximum continuous speed and 20 psi minimum at ground idle speed.
The pump inlet is the engine-airframe fuel interface. The pump is mechanically designed as a cantilevered pumping element on a rigid shaft running in two oil-lubricated sleeve bearings. Oil is supplied from the engine oil system through a face port mating with gearbox forbearing lubrication. Oil and fuel are separated by two dynamic carbon seals with a center vent to the engine overboard drain manifold. Shaft splines are lubricated by oil mist which is pumped through the splints by radial pumping holes in the pump quill shaft and the mating engine gearbox shaft.
The pumping elements are an ejector or jet pump, a mixed-flow centrifugal inducer, and a radial-flow impeller with a flow path from impeller discharge to provide ejector pump motive flow. This bypass flow is approximately equal to twice maximum engine flow at pump rated speed.
Refer to Figures 4-12 and 4-13 and refer back to 4-1. The fuel filter provides 30-micron absolute filtration for engine fuel prior to entering the high-pressure valve pump in the HMU. After passing through a core in the accessory gearbox, engine fuel flows to the filter through its inlet port from the engine boost pump discharge. The flow then is directed across the 30-micron pleated, barrier-type filter element (outside to inside direction), out the discharge port through another gearbox core, and to the HMU vane pump inlet.
The fuel differential pressure across the element is sensed across a magnet assembly piston on both the impending and actual bypass indicators. At 8-10 psi of pressure differential, the impending button piston assembly moves, carrying the magnet with it. As the magnet of the impending indicator moves away from the button, the magnetic field holding the red button in the inward position is reduced, allowing the button to pop out aided by a spring.
The electrical filter bypass sensor (or actual bypass indicator) is activated by a pressure signal as the bypass valve opens at 18-22 psi differential pressure. However, in this case the reduced magnetic field allows the button return spring inside the microswitch to actuate the switch button through the action of the pivoted stitch lever arm. This provides a cockpit indication of filter bypass.
Impending indicator mechanisms are locked when the wire and half-ball element and the magnet piston assembly move towards the center of the filter. This causes the half ball to drop and catch on the edge of the ramp. This latching prevents the magnet piston assembly from returning to its original position. In the impending by-pass position, the button can be pushed in its recessed position but will not remain since the magnet has not returned to provide the necessary magnetic field. The filter bypass sensor does not latch.
The impending bypass button cannot be reset until the filter element and bowl assembly are removed and the reset piston pushed upwards. The reset piston upper land then trips the half-spherical ball at the end of the locking wire, allowing the spring-loaded piston to return the magnet piston assembly to its original position.
The filter bypass valve is a poppet type. As the filter differential pressure reaches 18-22 psi, the bypass valve opens, and the sensor is activated for remote indication of filter bypassing. The filter element is disposable when loaded.
The HMU is mounted on the aft center of the AGB (Figure 4-14). It provides fuel pumping, fuel metering fuel flow computation, fuel pressurization, and fuel shutoff. It also provides gas generator speed control, compressor VG scheduling and actuation, and anti-icing and starting bleed valve actuation. The unit responds to PAS input for fuel shutoff, start, ground idle, up to maximum permissible gas generator speed, vapor venting, and electrical control unit override capability. The HMU also responds to an externally supplied load demand input via the LDS which is proportional to the power absorber load. This initially and directly coordinates gas generator speed and power to closely approximate the power required for the rotor or shaft power absorber. The HMU then responds to input from an electrical control unit (ECU) via the HMU torque motor. This trims gas generator speed as directed for both power turbine speed control and turbine temperature limiting for more exact load share control. In the event of an ECU failure, the HMU has the capability to mechanically deactivate the ECU. It also vents the unit case to the overboard drain in the event of excessive air or vapor at the inlet because of overtravel in the PAS. The HMU responds to sensed engine parameters (T2, P3 and N3) which influence fuel flow and VG position.
The HMU provides three major engine functions: fuel pumping, fuel metering, and VG positioning.
Fuel Pumping System (High Pressure). The main fuel pump is housed within the HMU. It is a vane-type, pressure-balanced pump which is self-priming, contamination-resistant, and interchangeable (at depot-level maintenance). The gearbox drives the pump spline at 9947 RPM at 100 percent Ng. The opposite end of the drive shaft serves as the rotational input to the HMU. The vane pump design was specifically developed with tungsten carbide vanes, pressure plates, and cam ring for contamination resistance.
The vane inlet pressure is internally boosted above engine boost discharge pressure by a fuel injector and injector bypass valve. This ensures positive charging of the vane element under all normal operating conditions and allows limited operation with a failed engine boost pump. The injector bypass valve maintains a differentail pressure across the injector of 12.5-135 psi. This results in a vane pump innlet pressure up to 30 psi above boost pump discharge.
A spring-loaded, high-pressure relief valve on the vane pump discharge prevents system overpressurization by limiting vane pump differential pressure to 832 psi.
Fuel Metering System. The HMU fuel metering system controls fuel flow to the engine. The components in this system are the--
High-pressure fuel from the pump passes around the outside of the servo supply wash filter and goes to the metering valve. The metering valve, in conduction with the pressure-regulating valve, meters required engine flow proportionally to the metering valve area and bypasses excess fuel through the PRV back to the pump inlet.
Variable Geometry Positioning System. A stator vane actuator is an integral part of the HMU. It provides positioning of the variable stator vanes and the anti-ice and the start bleed valve for efficient engine operation throughout the engine speed (Ng) range. A servo piston within the HMU provides an external output (force and position) to set the engine VG as a function of Ng and T2. This schedule is contoured on the Ng speed servo 3D cam. Motion of the schedule-follower lever moves a link pivoted on the actuator piston. This action strokes a double-acting, spring-loaded, rotating pilot valve. The valve then provides high-and low-fuel pressure signals to opposite sides for the VG actuator piston. Motion of the actuator returns the pilot valve to a null position, which corresponds to the input position being scheduled. The actuator provides minimum force output of 280 pounds and is positioned as a function of corrected speed.
Overspeed and Drain Valve. The T7-GE-701 valve Engine Shutdown controls the sequencing of fuel between the main manifold and the primer manifold (Figure 4-15). The valve also provides shutoff, drain, primer manifold purge functions, and Np overspeed protection. The assembly consists of the following components:
With the engine shut down, the pressure in the overspeed and drain valve is equal to ambient pressure, and the inlet and pressurizing valve is spring-loaded closed. When the engine is cranked over and the PAS is advanced to the stopcock in the HMU, the HMU delivers, metered flow to the overspeed drain valve inlet. This metered flow passes through awash-flow filter, and line pressure builds up until the inlet and pressurizing valve opens. Fuel boost pressure is directed to the top of a selector valve. Higher-pressure Wf passes through the open selector valve and flows to the 12 fuel injectors for engine lightoff and normal operation.
When the PAS is moved to the OFF position, the HMU is stopcocked, fuel flow to the overspeed and drain valve is stopped, and the inlet and pressurization valve is forced down by spring action. This opens the main fuel line in the overspeed and drain valve to overboard drain pressure. The residual P3 pressure in the combuster forces the fuel injectors and main manifold back through the overspeed and drain valve and out the overboard drain line.
The turbine overspeed function is designed to divert fuel flow from the combustor and flameout the engine to protect against destructive Np overspeed (Figure 4-16). When Np reaches 25,000 RPM (119.6 percent), the solenoid valve in the ODV is energized and opens a ball valve. This causes a bleed-off of equalizing fuel pressure on the bottom part of the selector valve. The high pressure W4 causes the selectro valve to move down and all W4 to bypass back to the HMU inlet. The instantaneous loss of W4 to the combustor causes the engine to flameout.
Twelve fuel injectors, installed in the midframe, receive fuel from the main fuel manifold and supply it to the combustion liner swirl cup (swirler) subassemblies (Figure 4-17). The swirlers consist of counter-rotating-flow primary and secondary vanes with venturi sleeves between the two radial inflow swirler vanes. Fuel from the injector is forced into these sleeves, and the counterrotating flow creates shear layers which break up the fuel film into finely atomized sprays.
The injector is a simplex nozzle with a single spin chamber fed by two spin holes. The swirled fuel from the spin holes exits through an orifice and discharges onto a 30° primary cone. This cone flares outward at the exit. The contoured exit provides a good-quality fuel spray at very low fuel pressure in the starting regime. The primary cone is housed in an air shroud which has six air holes. These six air holes feed air to augments the atomization of the fuel. This hydraulic design is simple; contains no valves, flow dividers, or secondary orifices and is purged by compressor discharge air during shutdown to prevent plugging.
The injector has a military standard fitting at the inlet with a last-chance, 0.009-inch diameter mesh screen to protect the orifice spin slots. A locating pin in the mounting flange assures correct orientation of the injector within the midframe casing. It is not necessary to remove the injectors before removing or replacing the combustion liner.
Modern fuel controls can be divided into two basic groups: hydromechanical and electronic. Hydromechanical controls are used most often. Fuel controls are extremely complex devices composed of speed governors, servo systems and feedback loops, valves, metering systems and various sensing mechanisms. Electronic fuel controls contain thermocouples, amplifiers, relays, electrical servo system switches, and solenoids. The discussion of fuel-control theory will address mainly the hydromechanical type. The simplest control is a plain metering valve to regulate fuel flow to the engine. This type of control could be installed on an engine used for thrust or gas generation.
Fuel controls may sense some or all of the following engine operating variables:
Some refinements might include a --
Like fuel controls for turbojet and turbofan engines, the fuel control for a turboprop or a turboshaft engine receives a signal from the pilot for a given level of power. The control then takes certain variables into consideration. It adjusts the engine fuel flow to provide the desired power without exceeding the RPM and turbine inlet temperature limitations of the engine. But the turboprop or turboshaft engine control system has an additional job to do that is not shared by its turbojet and turbofan counterparts. It must control the speed of the propeller or the free turbine, and it usually governs the pitch angle of the propeller blades.
Many turboprop and turboshaft engines in production today are the free turbine type. Engines of this kind act principally as gas generators to furnish high-velocity gases that drive a freely rotating turbine mounted in the exhaust gas stream. The free turbine turns the propeller through a system of reduction gears. If the engine is a turboshaft model, the free turbine rotates (also through reduction gears) a helicopter rotor or powers a machine. The General Electric T-700 and the Lycoming T-55-L-712 fuel control system follow.
The engine control system incorporates all control units necessary for complete control of the engine. The system provides for the more common functions of fuel handling, computation, compressor bleed and VG control, power modulation for rotor speed control, and overspeed protection. The system also incorporates control features for torque matching of multiple engine installations and overtemperature protection. The T-700 control system was designed for simple operation requiring a low level of pilot attention. The system performs many of the controlling functions formerly performed by the pilot. This has been done by providing--
Basic system operation is governed through the interaction of the electrical and hydromechanical control units. In general, the HMU provides for gas generator control in the areas of acceleration limiting, stall and flameout protection, gas generator speed limiting rapid response to power demands, and VG actuation. The ECU trims the HMU to satisfy the requirements of the load to maintain rotor speed, regulate load sharing, and limit engine power turbine inlet temperature.
Metering of fuel to the engine and basic engine control computations are performed in the hydromechanical control unit (Figure 4-18). The electrical and hydromechanical control units compute the fuel quantity to satisfy power requirements of the engine. The fuel and control system contains the following components:
For the T-700-GE-701/701A engines in a one engine inoperative (OEI) condition, the automatic contingency power switch, located in the ECU, resets the temperature limit amplifier to allow for more power is required for the operating engine.
Figure 4-19 lists the functional split between the two units mentioned above.
Fully integrated operation of the engine results from three inputs:
The HMU operates as a conventional gas generator power control when there is no input to the torque motor and with the load demand spindle set for maximum load (Figure 4-20). Engine schedules such as acceleration fuel flow and compressor VG position, are controlled by Ng and T2. The acceleration (or stall) limit is imposed by a maximum can schedule of Wf/P3 versus Ng and T2. A minimum flow schedule controls engine deceleration. Maximum and minimum metering valve stops provide absolute fuel-flow limits. An intermediate topping line controls maximum Ng speed. The topping line also limits excessive T4.5 in event of an electrical T4.5 control-system failure. Since these features are incorporated in the HMU, any power-available spindle motion will result in safe engine operation and will not cause engine damage. Except for intentional stopcocking of the control, no inadvertent shutdowns will occur from PAS motion.
A load-demand signal is introduced to the HMU through the LDS. When the LDS is reduced for its maximum setting with a reduction of aircraft rotor collective pitch setting the desired Ng is reset down from the prevailing PAS setting to provide immediate and accurate gas generator response. This new setting is trimmed by the ECU to satisfy the Np and load control functions established by the ECU.
The PAS sets a maximum available gas generator speed. The pilot usually sets this 120° PAS angle allowing the gas generator speed to reach a value which gives intermediate power. Through this collective pitch setting, the pilot positions the LDS to adjust available Ng to power level approximately equal to the rotor load demand power. The actual level of horsepower at 120° PAS angle will be more than required by the helicopter.
This is intentionally done for two reasons:
The ECU, powered from an engine-mounted alternator (Figures 4-21, 4-22)--
Before takeoff, the PAS lever is advanced from ground idle to 120°, which is the normal flight position. This allows the rotor head to accelerate to 100 percent Np/Nr with the collective still in flat pitch. Ng will rise as the PAS is advanced but will cut back after the Np governing speed of 100 percent is reached. (Np governing can be selected at 95-101 percent, but 100 percent is considered standard.)
To take off, collective pitch is increased, increasing the torque load on the power turbine. Simultaneously, the LDS rotates, calling for an increase in gas generator speed to keep Np/Nr from falling below 100 percent. The ECUs of each engine perform a fine trim of fuel flow to match torques and trim Np/Nr to 100 percent.
As Ng increases, the HMU schedules the bleed valve closed and the variable stators open to increase airflow through the combustor and turbine.
If collective pitch is increased to a very high angle, TGT may approach the temperature limiting value. When this occurs, the ECU prevents any further increase in fuel to the engine. If torque load is increased further, Np/Nr will droop below 100 percent because power turbine governing must be sacrificed to protect the engine against overtemperature.
At the end of the climb segment, less rotor lift is required and collective pitch is recked. The LDS rotates to reduce fuel flow and Ng; the variable stators will close slightly to optimize part-power fuel consumption and preserve stall margin.
Again, ECUs trim fuel flow. Upon entering a descent, the same sequence of events reduces Ng to the point that the bleed valve maybe opened. If the collective pitch is fully lowered, autorotation (power-off descent) is reached, and the torque drops to zero; overrunning clutches in the airframe transmission preclude a negative torque situation. Once the engines are uncoupled from the rotor, Nr is free to accelerate above 100 percent; and both engines will continue to govern Np at 100 percent. Nr and Np rotor speeds are held in reserve to pick up the rotor load when collective is again increased. This condition (100 percent Np zero torque) is known as flight idle.
The fuel control for the Lycoming T55-L-712 free power turbine engine is a hydromechanical type designated by Hamilton Standard as the JFC 31-22. It consists of the following main units:
Figure 4-23 schematically describes the fuel control system. Fuel control can be divided on a functional basis into two sections: the flow control section, consisting of the valving which meters engine fuel flow, and the computer section. The computer section includes the elements which schedule position of the flow control section metering valves as a function of the control input signals. The computer section also signals the closure of the compressor bleed valve.
|Figure 4-23. Fuel Control Schematic (T-55-L-11C/11D/11E/T55-L-712)|
|Editor's Note: This graphic is not viewable in HTML format. Check "Download Document" at top of this file for an alternate format or obtain a printed copy of the document.|
This computer operates primarily on four engine parameters gas producer speed (Ng), power turbine speed (Np), compressor discharge total pressure (Pt3), and compressor inlet total temperature (Tt2) for a given power lever position and power turbine speed-setting lever position. The ratio of engine fuel flow to compressor discharge total pressure is the manipulated variable used to control engine power output. This variable is biased by Tt2 during acceleration and Ng topping.
The position of the engine fuel flow metering valve is established by a multiplication of two positions representing compressor discharge pressure (P3) and fuel flow (Wf). The (Wf) engine fuel flow is made proportional to the metering valve position by maintaining a constant differential pressure across the valve flow area, which varies linearly with valve position. This constant pressure in maintained by the pressure regulating valve. This valve is positioned as necessary to maintain a 40-psi differential across the metering area. To accomplish this, all pump output not required to operate servos and supply engine flow requirements is bypassed back to the pump inlet.
The P3 servo (motor) is positioned by directing compressor discharge (P3) to a bellows, which positions the servo directly proportional to P3 air pressure. The servo is connected to the P3 ramp, which moves proportional to servo inputs. The ramp works directly against the ratio servo and in turn positions the metering valve.
The ratio servo may be positioned by anyone of four inputs: the Np governing linkage, the Ng governing linkage, acceleration limiting linkage, or the maximum flow limiter. The inputs, which schedule the lowest fuel flow, determine the position of the metering valve. The ratio servo and the Ng servo provide an input to the compressor bleed air signal mechanism.
The Np governor linkage input to the ratio servo is controlled by the Np servo piston whose position is established by a flyball-operated pilot valve. It functions as a droop-type governor, limiting fuel flow proportional to speed over its operating range, which is set by the power turbine speed setting lever position. The Np governing linkage input to the ratio servo is biased by Ng speed such that Np governing droop line position increases in speed as Ng speed decreases.
The Ng governing linkage input to the ratio servo is controlled by the Ng servo piston whose position is established in the same manner as the Np servo. It functions as a droop-type governor, limiting the ratio servo inversely proportional to Ng speed over its operating range. Its operating range (ground idle to maximum speed) is established by the power and shutoff lever position. This linkage is also biased by T2 through the speed reset linkage.
The acceleration limiting linkage input to the ratio servo is controlled by a contour on the 3D cam. The 3D cam is positioned longitudinally by the Ng servo piston and rotated by the T2 servo piston. This provides a unique cam positioned for each set of NG speeds and T2 conditions. The acceleration finding cam contour schedules the maximum allowable fuel of low P3 ratio at each speed and temperature setting. The 3D cam also positions the speed reset linkage through the speed reset contour and provides an input to the bleed air signal mechanism.
The position of the T2 servo is established by the liquid-filled bellows which positions the servo directly proportional to the compressor inlet temperature.
The maximum flow limiter input to the ratio servo limits the engine shaft torque by limiting fuel flow. This is done by positioning the ratio servo as necessary to reduce fuel flow at a safe value.
The control computer section also provides a pneumatic signal to the bleed air acutator which is used to open or close the bled air during appropriate phases of engine operation.