Beechcraft D. Cessna TB Nov 19, A90, Quincy Municipal Airport, May 12, Issue Flight ManualSupplement. FMS Issue 5- Compliance. Record S COMP Issue Instructions for Continuing. Airworthiness S Flight Manual Supplement. FMS Issue 5The. Dec 19, Charlotte-Douglas International Airport, Charlotte, Feb 13, King Air. Type Certificate Holder Data Recorder installation. Line Fusion installed. Abstract: This report explains the accident involving United Express flight , a.
Beechcraft C, and a All Rights Reserved. Designed by Templatic. Home Copyright Privacy Contact. The collector tank sump drain is located in the center wing adjacent to the fuselage; the inboard leading-edge tank sump drain is on the underside of the wing just outboard of the nacelle; and the integral wet wing fuel tank sump drain is located approximately midway on the underside of the wing aft of the main spar.
Other drains are the fuel filter drain, in the main landing gear wheelwell, and the center wing tank drain at the wing root forward of the flap. Since jet fuel and water are of similar densities, water does not settle out of jet fuel as easily as from aviation gasoline.
For this reason, the airplane must sit perfectly still, with no fuel being added, for approximately three hours prior to draining the sumps if water is to be removed.
Although water ingestion is not as critical for turbine engines as it is for reciprocating engines, water should still be removed periodically to prevent formations of fungus and contaminationinduced inaccuracies in the fuel gaging system.
When draining flush-mounted drains, do not turn the draining tool. The Beechcraft Airliner fuel system Figures and is designed with the pilot in mind; simple to use in normal and emergency conditions with one or more failures.
Simple, quick over-the-wing refueling is also incorporated to minimize ground turnaround time requirements. The wet wing fuel system is quite different from the previous series airliners, although pilot operation of the two systems is very similar. The wet-wing fuel system consists of two integral fuel tanks in each wing Figure A main tank extends from engine nacelle to wing tip.
An auxiliary tank is located between the engine nacelle and the fuselage. The maximum allowable fuel imbalance between the wings is pounds. A collector tank is contained within each main tank immediately outboard of the nacelle. Each collector tank is filled from its main tank by gravity feed and two jet transfer pumps, which maintain the fuel level in the collector tank at normal flight attitudes. Each auxiliary tank is filled through its own fill port located just inboard of the engine nacelle.
When auxiliary tank fuel is required for a planned flight, the main tanks should be full and the additional fuel to complete the flight placed in the auxiliary tanks. The auxiliary tank fuel should be used first. There is no gravity flow between the main and auxiliary tanks, therefore, each must be filled separately. These conditions occur at lower power settings when the auxiliary tank transfer pump is supplying fuel to the collector tank, and the high-pressure pump purge line is directing fuel back to a full main tank.
Since the main tank is already full, any excess fuel flows through the vent system back to the auxiliary tank. This condition can continue until the auxiliary tank is empty. The second condition which will allow fuel to flow from the main tank to the auxiliary tank is thermal expansion. Fuel will not vent outside from the wing unless the auxiliary tank is full.
The fuel system is vented through a float operated valve near each wing tip. The system contains a flush vent with flame arrester, a heated ram-air vent, to maintain a slight positive tank pressure during flight, and a recessed ram vent.
The recessed ram vent is coupled to the protruding ram vent on the underside of the wing tip. The recessed vent is naturally ice resistant, while the protruding vent is heated to prevent icing.
The check valves in the vent tubes allow the air to flow one way through the vents. The flame arrestors, on the flush vent and incoming line, prevent a flame front produced by a lightning strike or static discharge from traveling up the vent line into the tank system. The vent system also incorporates a pressure-activated relief tube which prevents an overpressure condition in the tank.
A valve in the tube opens when the pressure exceeds a set amount. Vent lines connect the main tank and auxiliary tank as we have discussed earlier. As fuel is used from the main tank, it is gravityfed and also pumped through motive flow to the collector tank.
The cross vents to the auxiliary tank then are open and equalize the pressure in all tanks. An anti-siphon valve is installed in each tank filler port to prevent loss of fuel through siphoning in the event of improper securing or loss of the filler cap.
Power for the aux transfer pumps is supplied through their respective generator busses. All other fuel system functions on the control panel require only battery power. During normal operation, fuel flow to each engine is provided by the engine-driven fuel pumps high pressure and boost which draw fuel from the collector tank in the same wing Figure The collector tank draws from its respective main tank unless fuel is being supplied from the auxiliary tank. Any fuel contained in the auxiliary tanks is to be used prior to using fuel from the main tanks.
The auxiliary tank fuel will be used first. A loss of electrical power or failure of the transfer pump will prevent the use of auxiliary tank fuel. The auxiliary tank will not gravity feed into the main tank fuel system. The auxiliary tank transfer pump uses fuel for cooling. A thermal cutout switch has been incor-. Engine fuel is supplied from the collector tank through the motive-flow system operated by either the engine-driven boost pump, or the standby pump within the collector tank.
The motive-flow system is supported by a series of three transfer jet pumps main, forward, and aft. The main jet pump is located within the collector tank. It picks up fuel from the collector tank and sends fuel to the engine.
The aft jet pump is also located within the collector tank but it draws fuel from the main tank and fills the collector tank. The forward jet pump draws fuel from the main tank, forward of the wing spar, and sends it to the collector.
There are also three flapper valves and three upper wing stringer cutouts that will allow the collector to be gravity fed from the main tank.
The engine-driven boost pump draws fuel from the collector tank using the jet pump principle. The components we are concerned with are the motive-flow line, the primary jet pump, two transfer jet pumps, the standby electric boost pump, and the fuel supply line.
Fuel from the motive-flow line passes through the primary jet pump, which is actually a venturi. In order to pull fuel from the collector tank, a venturi effect is used. As a mass of fuel is accelerated through a small opening or venturi, it causes a drop in pressure. At this low-pressure point, fuel from the collector tank enters the fuelsupply line through a filter and low-pressure fuel is supplied as needed. The transfer jet pumps draw fuel into the collector tank from the main tank in the same manner.
The firewall shutoff valves are motor-driven and. If the primary engine-driven boost pump fails, fuel can be supplied to the system by the standby electric boost pump. This pump draws fuel directly from the collector tank and passes it to the fuel supply line. As fuel is drawn from the collector tank, it flows through the manual shutoff valve, a fuel filter and through the firewall shut-off valve. When fuel in the main tanks reaches a level allowing approximately 30 minutes of flight time at maximum continuous power, the right, or left, or both FUEL QTY annunciators will be actuated by fuel level sensors mounted on the forward side of the tank.
If the fuel in the collector tank area drops to a reserve of two minutes flight time at maximum continuous power, the right, left or both, FUEL FEED annunciators will light up.
A switch-controlled cross-transfer valve in the left wing is externally connected into the line. A manually operated cross-transfer control switch is mounted on the upper fuel control panel, just above the fuel quantity gages.
When the cross-transfer control switch is actuated, the cross-transfer valve opens to allow the standby fuel boost pump to transfer fuel to the opposite collector tank. In addition to the cross-transfer function, the electric boost pump can provide fuel to the engine should the engine-driven boost pump fail. Power for the switches is drawn through the circuit breakers at the bottom of the fuel panel.
During single-engine operation, it may become necessary to supply fuel to the operative engine from the fuel system on the opposite side. The simplified cross-transfer system is placarded for fuel selection with a diagram on the upper fuel control panel. This opens the cross-transfer valve, energizing the standby pump on the side from which cross transfer is desired. In the event one of the electric boost pumps fail, cross-transfer can only be accomplished from the side of the operative pump.
To discontinue fuel transfer operations, the transfer flow switch need only be placed in the center OFF position.
During engine. The indicators are marked in pounds. Once the switch is released, the fuel indicator will return to read only the main tank quantity. Fuel quantity probes, which are part of the fuel gaging system, are capacitance-type probes. These probes measure the density of fuel on board, and are calibrated to read pounds of fuel.
This system is necessary on this airplane because the engines operate on weight flow of fuel rather than gallons. The gages in the cockpit therefore read in pounds of fuel flow and pounds of fuel remaining in the tanks. To provide the pilot with such a readout, it is necessary to use a system which compensates for changes in the specific gravity of the fuel in use. Therefore, each probe is designed to compensate for differences in specific gravity. Each main tank contains six fuel quantity probes, and each auxiliary tank two probes.
Information from these probes is relayed to the fuel panel in the cockpit to show fuel remaining in each tank. These pumps are activated. The standby pumps are also used for all cross-transfer operations. In the event one of the electric pumps fail, cross-transfer can only be accomplished from the side of the operative pump. With fuel in the auxiliary tank and the transfer switch in AUTO, the auxiliary tank transfer pump will run once the 10 psi low-pressure switch is activated.
It will continue to run until the boost pressure falls below 10 psi, or the float switch and transfer line low-pressure switches open; in either case the auxiliary tank transfer pump will shut down. This activates the electric transfer pump in each auxiliary tank and pumps fuel to the collector tank of the same wing. Fuel will continue to be transferred until the auxiliary tank is empty, at which time the pump will automatically shut off. In the event of a transfer system failure, it is permissible to temporarily operate the airplane with fuel in the auxiliary tanks providing fuel imbalance and fuel reserve requirements can be met.
Normal procedures call for the pump to be left in the AUTO position. In AUTO, there are four additional non-pilot operated control features. The gauges, when not submerged in fuel, are red with a black dot; when they are submerged they are totally black. The outboard probe, when red, indicates less than 1, pounds of fuel, the inboard probe, when red, indicates less than pounds of fuel. During preflight, the fuel sumps on the tanks, pumps and filters should be drained to check for fuel contamination.
There are six five in Series UE sump drains in each wing Figure The two one in Series UE collector tank sump drains are located below the wing on the outboard side of the nacelle; the two main tank drains are located on the underside of the wing, outboard of the nacelle, one forward and one aft of the main wing spar: the other drains are the fuel filter drain located on the underside of the wing, outboard of the nacelle under a springloaded access panel and the auxiliary tank drain at the wing root forward of the flap.
Although water ingestion is not as critical for turbine engines as it is for reciprocating engines, water should still be removed periodically to. The quantity of water contained in the fuel depends on its type and temperature. Kerosene, with its higher aromatic content, tends to absorb and suspend more water than aviation gasoline.
Along with water, kerosene will suspend rust, lint and other foreign materials longer. Given sufficient time, suspended contaminants will settle to the bottom of the tank. The settling time for kerosene is five times that of aviation gasoline; therefore, jet fuels require good fuel handling practices to ensure servicing with clean fuel. If recommended ground procedures are carefully followed, solid contaminants will settle, and free water can be reduced to 30 parts per million ppm , a value considered acceptable by the major airlines.
Dissolved water has been found to be the major potential fuel contaminant. Its effects are multiplied in aircraft that operate primarily in humid regions and in warm climates. Since most suspended matter, including water, can be removed from the fuel by allowing sufficient settling time and by proper filtration, fuel contamination is usually not a major problem.
Dissolved water cannot be filtered from the fuel by micronic-type filters used in the fuel system; however, water in the fuel can be released by lowering fuel temperature, which occurs in flight. The difference of 40 ppm will have been released as super-cooled water droplets which need only a piece of solid contaminant or an impact shock to convert them into ice crystals.
Tests indicate that released, super-cooled water droplets will not settle during flight. Droplets are pumped freely through the system.
If they become ice crystals in the tank, they will not settle since the specific gravity of ice is approximately equal to that of kerosene. Sludge and other fuel contaminants can cause corrosion of metal parts in the fuel system and clogging of the fuel filters. Although the Airliner uses integral wet wing fuel cells in each wing, and all metal parts except the standby boost pumps and jet transfer pumps are mounted above the settlement areas, consistently using contaminated fuels can cause filters to clog and fuel pumps to corrode.
This applies not only to maintaining a clean fuel supply, but to keeping the aircraft system clean. The following is a list of steps that may be taken to recognize and prevent contamination problems. Know your supplier. It is impractical to assume that contaminant-free fuel will always be available. But, it is feasible to exercise caution and be watchful for signs of fuel contamination. Be sure, as much as possible, that fuel has been properly stored. Fuel should be filtered as it is pumped to the truck, and again as it is pumped from the truck to the aircraft.
Perform filter inspections to determine if sludge is present. Maintain good housekeeping by periodically flushing the fuel tankage system. The frequency of flushing will be determined by the climate and the presence of sludge.
Aviation gas is an emergency fuel. If avgas has been used, observe the requirement for hours maximum operation on aviation gasoline before engine overhaul. The time should be logged in the aircraft engine operation records as gallons of avgas added to the fuel system. Use only clean fuel servicing equipment. After refueling, allow a settling period of at least three hours, whenever possible; then drain a small amount of fuel from each drain. Fuel spills on airplane tires have a deteriorating effect.
Be sure to remove spilled fuel from the ramp area immediately to prevent tire damage. Even if the fuel does not contain water, or if water has been drained, the possibility of fuel icing still exists at some very low temperatures. The oil-to-fuel heat exchanger prevents fuel icing during most normal operating conditions; however, in extremely cold temperatures at some cruise altitudes, anti-icing fuel additives must be used.
This chart is used as a guide in preflight planning to determine operating temperatures where icing at the fuel control unit could occur Figure Enter the graph with the known or forecast outside air temperature at cruise, and plot vertically to the expected cruise pressure altitude.
Since no fuel temperature measurement is available prior to the heat exchanger, fuel temperature must be assumed to be the same as outside air temperature. Next, plot horizontally to determine the minimum oil temperature required to prevent icing. The Airliner maintains a constant oil temperature, although the exact temperature varies from one airplane to another. Compare the minimum oil temperature obtained in the preceding example with the normal oil temperature of the airplane to be used for the flight to determine if anti-icing additive is needed.
When required, anti-icing additive conforming to specification MIL-I should be added during fueling. If the Airliner is fueled with aviation gasoline, some operational limitations must be observed. If use of aviation gasoline is necessary, operation is limited to hours before engine overhaul, and is prohibited if either of the two standby pumps is inoperative or if flight is conducted above 15, feet 18, feet in Series UE.
When avgas is used, lead deposits form on the turbine wheels causing power degradation; therefore, when operating on avgas, the lowest octane rating available should be used because its lead content is lowest. Since the aviation gas will probably be mixed with jet fuel already in the tanks, it is easier to record the number of gallons of avgas added than to note hours of operation.
If an engine has an average fuel consumption of 55 gallons per hour, each time 55 gallons of aviation gasoline are added, one hour of the hour limitation is being used. Because it is less dense, aviation gas delivery is much more critical than jet fuel delivery; therefore, operation on avgas is prohibited if above 15, feet 18, feet in Series UE.
Aviation gas feeds well under pressure but cannot suction feed as well, particularly at high altitudes. For this reason, two alternate means of pressure feed must be available. Standby pumps provide alternate pressure feed capability, and both are required to be operational when avgas is used.
Brand names are listed for easy reference and are not specifically recommended by Beech Aircraft Corporation. Any product conforming to the recommended specification may be used.
Make sure the aircraft and the servicing unit are both grounded to the ground, and that the aircraft is statically grounded to the serving unit. The filler caps are located in the main fuel tank on the leading edge of each wing near the wing tip and the auxiliary tank fuel caps are just inboard of each nacelle. Do not rest fuel nozzle in tank fillers because this may damage the filler neck. Allow a three-hour settling period whenever possible, then drain a sufficient amount of fuel from each drain point to remove water or contaminants.
The adapter contains a check valve to prevent fuel drainage when the plug is removed. Each wing fuel system may be drained as follows: a. Cut the safety wire and remove the plug.
This will seat the check valve. Thread an AN adapter into the drain, unseating the check valve to start the flow of fuel. Fuel will gravity drain. Overspeed Governor Engine Installation Engine Cutaway Engine Gas Flow Free-Turbine Reverse Flow Principle Engine Modular Concept Engine Start and Ignition Switches Typical PT6A Engine Engine Lubrication Diagram Engine Oil Dipstick Magnetic Chip Detector Simplified Fuel System Diagram Simplified Fuel Control System Fuel Flow Gages Fuel Pressure Annunciator Control Pedestal Control Levers Overtemperature Limits Starting View through Exhaust Duct In-Flight Engine Data Log Hartzell Propeller Propeller Tiedown Boot Installed Propeller Blade Angle Diagram Primary Governor Diagram Propeller Onspeed Diagram Propeller Overspeed Diagram Propeller Underspeed Diagram Beta and Reverse Control Beta Range and Reverse Diagram Propeller Postioning Diagram Overspeed Governor Diagram Power Levers Propeller Control Levers Propeller Synchrophaser Performance within the normal parameters of powerplant and propeller systems extends engine life and ensures safety.
This chapter describes basic engine components, limitations, and system checks. In-depth knowledge of the propeller system is essential to proper operation of the engine power system. Operating within safe parameters of the powerplant and propeller systems extends engine life and ensures safety. This chapter also describes the propeller system and its operational limits and preflight checks. The purpose of this chapter is to provide pilots with sufficient engine operating details to further understand normal, abnormal, and emergency procedures.
Descriptions include primary and overspeed governors, autofeather system, and synchrophaser. This chapter also presents a description and discussion of the propeller system. Location and use. The PT6AB reverse flow, free-turbine, turboprop engine Figure is flat-rated to 1, shaft horsepower. The PT6AD is flat-rated to 1, shaft horsepower. The engines are equipped with composite fourblade, full-feathering, reversing, constant-speed propellers mounted on the output shaft of the engine reduction gearbox.
Engine oil supply and single-action, engine-driven governors control propeller pitch and speed. When the engines are shut down, propellers automatically feather, and will unfeather when engines are started as engine oil is pumped into the propeller dome. Reference to the right or to the left side of the aircraft, propellers, or engines always assumes the pilot is looking from the rear of the aircraft forward Figure SHP is determined by propeller rpm and torque applied to turn the propeller shaft.
Hot exhaust gases leaving the engine also develop some kinetic energy similar to a turbojet engine. The engine specification tables show engine ratings and temperatures Tables and Engine Type Free Turbine Type of Combustion Chamber Annular Compression Ratio Counterclockwise Propeller Shaft Rotation looking forward Clockwise Propeller Shaft Gear Ratio To identify locations in the engine, it is common practice to establish engine station numbers at various points Figure To refer to pressure or temperature at a specific point in engine airflow path, the appropriate station number is used, such as P3 for Station 3 pressure or T5 for gas temperature at Station 5.
For instance, airflow temperature measured between the compressor and first-stage power turbine at Engine Station 5, is called interstage turbine temperature ITT or T5. Bleed air, located after the centrifugal compressor stage and prior to entering the combustion chamber, is commonly referred to as P3 or bleed air. Bleed air is used for cabin heat, pressurization, and the pneumatic system. These terms should be memorized since they are used often when describing PT6A engines.
N1 or Ng—Gas generator rpm in percent of turbine speed. Np—Propeller rpm. Also referred to as axial stage air or compressor interstage air. P3—Air pressure at Engine Station 3, the source of bleed air used for some aircraft systems. The compressor turbine drives the engine compressor and accessories.
Dual power turbines drive the power section and propeller through the planetary reduction gearbox. Compressor and power turbines are mounted on separate shafts and are driven in opposite directions by gas flow across them.
Inlet air enters the compressor at the aft end of the engine, moves forward through the combustion section and the turbines, and is exhausted at the front of the engine. The gas generator module includes the compressor and the combustion section. Its function is to draw air into the engine, add energy to it in the form of burning fuel, and produce the gases necessary to drive the compressor and power turbines.
The power module converts the gas flow from the gas generator into mechanical action to drive the propeller. This is done through an integral planetary gearbox, which converts the high-speed, low torque of the power turbine to low-speed, high torque required at the propeller.
The reduction ratio from power turbine shaft rpm to propeller rpm is approximately Since it is not necessary to remove the engine from the airplane to accomplish the HSI, inspection is both simple and fast. Because of modular design, the gas generator section or the combustion section can be completely replaced independently of each other. This feature permits easy maintenance, modular overhaul, and onwing HSI. Inlet air enters the rear of the engine through an annular plenum chamber, formed by the compressor inlet case, where it is directed forward to the compressor see Figure The compressor consists of four axial stages and a single centrifugal stage assembled as a single unit on a common shaft.
Rows of stator vanes between each stage of compression diffuse the air, raise its static pressure, and direct it to the next stage of compression. Now diffused, the air passes through straightening vanes to the annulus surrounding the combustion chamber liner. The combustion chamber liner contains perforations of varying size that allow entry of compressor delivery air.
For smoother engine starts, the PT6AB fuel is introduced into the combustion chamber liner in two stages through 14 simplex fuel nozzles. The nozzles are supplied by a dual fuel manifold consisting of primary and secondary transfer tubes and adapters.
In the PT6AD engine, fuel is introduced through 14 duplex nozzles. After combustion, expanding gases reverse direction in the exit duct zone, and pass through compressor turbine inlet guide vanes to the s i n g l e - s t a g e c o m p r e s s o r d r iv e t u r b i n e.
Expanding gases are then directed forward through power turbine inlet guide vanes to drive the power turbine section. The guide vanes ensure that expanding gases impinge on the turbine blades at the correct angle with minimum energy loss.
Exhaust gas from the power turbines is then directed to the atmosphere through abifurcated exhaust plenum to twin opposed-exhaust ports. Compressor and power turbines are located in the approximate center of the engine, with their respective shafts extending in opposite directions. This feature simplifies installation and inspection procedures. A compressor bleed valve compensates for excess air flow at low rpm by bleeding axial stage air P2.
Pressure relief helps prevent axial stage compressor stall. The compressor bleed valve is a pneumatic piston that references pressure differential between axial and centrifugal stages. The valve helps prevent. At low N1 rpm, the compressor bleed valve is open. If the compressor bleed valve were to stick closed at low N1 speeds, compressor stall could result from an attempt to accelerate the engine to higher power.
If the valve were to stick open at high N1 speeds, power output would be considerably reduced. With the valve open, and. The valve relieves excess P2. Two spark-type igniters in the combustion chamber provide positive ignition during engine start. Although the engine is equipped with two igniters, it needs only one for start.
The system is designed so that if one igniter malfunctions, the remaining igniter will continue to operate. A jet flap slot, which secures the accessory section to the engine compressor section, is machined into one side of each hollow strut.
A jet flap intake system Figure functions as a variable inlet guide vane without variable geometry. Compressor interstage air P 2. Spark ignition is effective for quick engine starting throughout a wide temperature range.
The system consists of an airframe-mounted ignition exciter, two individual high-tension cable assemblies, and two spark igniters. It is energized from the aircraft nominal volt DC supply, and will operate in the 9 to volt range. The igniter system can produce up to 3, volts.
The leverlocked ON position activates the starter and both igniters. Igniters are not energized in this position. The ignition system features automatic capability.
The auto-ignition system should be armed in turbulence, precipitation, and icing conditions. Accessories are driven by the compressor shaft N1 through a coupling shaft. One lubricating oil pressure pump and two scavenge oil pumps are mounted inside the accessory gearbox. Two additional oil scavenge pumps are externally mounted. Each mounting pad has its own specific gear ratio. It also provides oil to the propeller governor and propeller reversing control system.
The main oil tank houses a geartype engine-driven pressure pump, an oil pressure regulator, a cold pressure relief valve, and an oil filter. The engine oil tank, an integral part of the compressor inlet case, is located in front of the accessory gearbox. As oil is pumped from the tank, it passes through pressure- and temperature-sensing bulbs mounted on the rear accessory case. Oil is then delivered through an external oil transfer line below the engine to bearing compartments and to the nose case.
Gear-type scavenge pumps return the oil through external oil transfer lines and through an external oil cooler below the engine. The oil cooler is thermostatically controlled to maintain desired oil temperature. Another externally mounted unit, the oil-to-fuel heat exchanger, uses hot engine oil to heat fuel before it enters the engine fuel system.
Total oil system capacity is 3. Maximum oil consumption is 1 quart every 10 hours; however, normal oil consumption may be as little as 1 quart per 50 hours. Most PT6A engines normally seek an oil level of one to two quarts low.
When adding oil between oil changes, do not overfill, and do not mix types or brands of oil due to the possibility of chemical incompatibility. A placard inside the engine cover shows the brand and type of oil used in that particular engine. Although the preflight checklist calls for checking oil level, the best time to check oil quantity is shortly after shutdown, since oil levels are most accurately indicated at that time.
Oil level checks during preflight may require motoring the engine to obtain an accurate level indication. The oil tank is provided with a filler neck and integral quantity dipstick housing.
The cap and dipstick are secured to the filler neck, which passes through the gearbox housing and accessory diaphragm into the tank. Dipstick markings indicate the number of U. Engine parameters should be monitored for abnormal indications.
If parameters are. They are the primary low-pressure boost pump, oil-to-fuel heat exchanger, high-pressure fuel pump, fuel control unit, fuel cutoff valve, flow divider, and dual fuel manifold with 14 simplex nozzles, 14 duplex nozzles in the D. The low-pressure boost pump is engine driven, and operates when the gas generator shaft N1 is turning. It provides sufficient fuel head pressure approximately 45 psi maximum for proper cooling and lubrication of the high-pressure pump.
The oil-to-fuel heat exchanger regulates fuel temperature at the fuel pump inlet to prevent icing at the pump filter.
This is done automatically and requires no action by the pilot. After fuel passes through the oil-to-fuel heat exchanger, it flows into the high-pressure, engine-driven fuel pump and into the fuel control unit FCU. Prior to entering the FCU, a fuel purge line constantly directs a small amount of fuel back to wing fuel tanks to clear vapors and bubbles from the fuel control system. The high-pressure fuel pump is an engine-driven, gear-type pump that can supply fuel at psi maximum pressure to the fuel side of the FCU.
Its primary purpose is to supply sufficient pressure to fuel nozzles for adequate spray pattern during all modes of engine operation. Flow rates and pressures will vary with changes in gas generator N1 rpm.
The fuel cutoff valve is internal in the FCU. The valve is controlled by the condition lever, and is either open or closed; it has no intermediate position. When the fuel cutoff valve is open, fuel flows to the minimum pressurizing valve, which blocks fuel flow during start until fuel pressure is sufficient for proper spray pattern in the combustion chamber. As high-pressure fuel pump output.
If the high-pressure pump fails, the valve will close and combustion will cease. During start, fuel flows initially through the flow divider to seven primary fuel spray nozzles in the combustion chamber. All 14 nozzles then deliver atomized fuel to the combustion chamber.
The progressive sequence of primary and secondary fuel nozzle operation provides cooler starts. Increased acceleration in N 1 speed may be noticed when secondary fuel nozzles activate. The system consists of a P3 accumulator purge tank with P3 air input at one end and P3 discharge to the flow divider at the other end. As long as the engine is running, fuel pressure keeps the flow divider purge port closed.
As fuel pressure drops to zero during engine shutdown, P 3 air escapes through a check valve into the flow. As a result, the pilot may notice a one- or two-second delay in initial engine spooldown after the condition levers are moved into fuel cutoff. Flow rates are calibrated for starting, acceleration, and maximum power.
The FCU compares gas generator speed N1 with power lever setting and regulates fuel to engine fuel nozzles. The FCU also senses compressor discharge pressure and compares it to N1 rpm to establish acceleration and deceleration fuel flow limits.
A minimum flow stop, set to approximately 90 pounds per hour per engine, guarantees sufficient fuel flow at all operating altitudes to sustain engine operation at minimum power. Download the Actiontec CA info sheet. WiFi supported: 2. Works with these CenturyLink internet services:.
Note: Listed speeds reflect the maximum download speed the technology can deliver on CenturyLink services. In-home speeds may differ based on router setup, devices and other factors. Download the latest firmware for the CA. Tip: If the firmware link doesn't work, try another browser Chrome is known to have issues with this kind of link. Or, right-click the link , select "copy link address" , then paste the link into a new browser tab.
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