Fuel.docx

2A-28-10: General Fuel (See Figure 1. Fuel System Block Diagram.) The fuel system for the G650ER consists of two (2) integral wing tanks. Total usable capacity is 48,200 pounds (approximately 7,194 U.S. gallons). The integral tanks are baffled to minimize aircraft center of gravity shift with chan...

2A-28-10: General Fuel (See Figure 1. Fuel System Block Diagram.) The fuel system for the G650ER consists of two (2) integral wing tanks. Total usable capacity is 48,200 pounds (approximately 7,194 U.S. gallons). The integral tanks are baffled to minimize aircraft center of gravity shift with changes in aircraft attitude. Each integral tank contains a hopper in which the pumping system for each wing is located. Adequate fuel quantity in each hopper is maintained during flight by gravity, supplemented by ejector pumps which pump fuel from tank areas outside of the hopper. A crossflow line, connecting the two (2) pumping systems, is provided to allow fuel to flow from one tank to either or both engines. A cockpit-controlled intertank valve permits fuel balancing in flight. The fuel system for both engines are self-bleeding to minimize maintenance. The fuel tanks can be filled from a single-point pressure fueling adapter or through tank fill openings on top of each wing. Fuel is drawn from the tanks and pressurized for distribution to the airplane engines and/or APU by boost pumps mounted on the rear wall of each tank. Sensors mounted within the fuel tanks provide information for cockpit displays, enabling the flight crew to monitor fuel quantity and fuel tank temperature. The fuel system is divided into the following subsystems: 2A-28-20: Wing Fuel Tanks System 2A-28-30: Fuel Distribution System 2A-28-40: Fuel State Indication Figure 1. Fuel System Block Diagram FUEL GAUGING PROBES AND TANK SENSOR L Wing 1 R Wing 1 L Wing 5 L Wing 4 P T L Wing 3 P P L Wing 2 P P P FCS K D T P P P P FCS K D P T P P R Wing 2 P P R Wing 3 P P R Wing 4 P T R Wing 5 L Wing 6 P P Hopper EACH TANK: Hopper P R Wing 6 P P P L Wing 7 P P P HLS 14 STANDARD TANK GAUGING PROBES 1 DUAL TANK GAUGING PROBE (HOPPER) 1 FUEL CHARACTERISTICS SENSOR (FCS) 1 HIGH LEVEL SENSOR (HLS) 1 FUEL TEMPERATURE SENSOR P P R Wing 7 P P P HLS LEFT WING TANK RIGHT WING TANK CHANNEL 1 1 CHANNEL 2 CHANNEL 1 CHANNEL 2 1 7 6 5 4 3 2 1 HOP 7 6 5 4 3 2 1 HOP 5 7 6 5 4 3 2 1 HOP 7 6 5 4 3 2 1 HOP 5 HLS P P P P P P P P FCS K T D P P P P P P P P T HLS P P P P P P P P FCS K T D P P P P P P P P T LEFT REMOTE DATA CONCENTRATOR (RDC) A LEFT REMOTE DATA CONCENTRATOR (RDC) B RIGHT REMOTE DATA CONCENTRATOR (RDC) A RIGHT REMOTE DATA CONCENTRATOR (RDC) B TIL-006748 Figure 2. Fuel System Diagram TIL-037355F 2A-28-20: Wing Fuel Tanks System General Description: The wing fuel tanks are internal compartments bounded by upper and lower wing skin and the front and rear wing spars. Six ribs in each wing are partially sealed at the bottom and are equipped with flapper flow valves which inhibit fuel "sloshing" during aircraft maneuvers. The wing fuel tanks are integral to the wing structure. Fuel is contained within most of the interior of the wing, with the tank dimensions defined by the front wing spar, rear wing spar and the upper and lower wing skin. The interior of the wing is coated with a sealant during manufacturing to prevent fuel leakage. Four electrically-driven, brushless DC fuel pumps, two in each fuel tank hopper, are provided to maintain the required fuel pressure at the engine. These pumps are mounted on the wing rear beam for ease of replacement without defueling the aircraft. They are canister mounted such that they can be replaced without draining or entering the tank. The pump assembly is mounted into the canister through holes on the rear spar in the collector compartment. The shape of the wing accommodates the installations necessary for efficient operation of the fuel system. The tank area near the wing root has the largest volume and houses the fuel boost pumps and fuel feed lines. With wing dihedral of 2.81°, fuel within the wing will always flow towards the wing root, ensuring that the fuel boost pump inlets will be adequately supplied until all usable fuel has been consumed. To prevent the outward movement of fuel during flight maneuvers involving turns, the six wing ribs within the tank that form the contour of the wing are fitted with baffles hinged to open only in the direction of the wing root. The wing ribs also contain a series of large holes above the baffles to allow fuel to flow outboard during single-point pressure fueling, since the single-point access is located near the wing root. A second set of 3/8" holes penetrate the wing ribs below the baffles to allow residual fuel and any accumulated water to drain inboard to the wing root as fuel quantity decreases. The following installations are contained within each fuel tank structure: Left and Right Fuel Hoppers Over-Wing (Gravity) Fueling System Gravity Water / Fuel Drain System Fuel Ventilation System Fuel Outlet Strainer System Description of Subsystems, Units and Components: Left and Right Fuel Hoppers: Each wing tank has an internal compartment at the wing root, termed a hopper, at the center aft portion of each tank. The hopper is fed from the main wing tank by an ejector pump. Motive flow for the ejector is taken from boost pump output. During all phases of flight (except takeoff), the ejector pump flow exceeds engine demand, and excess fuel overflows the hopper back to the main portion of the wing tank. The outboard wing rib and the forward wall have baffles directing fuel into the hopper, and drain holes for residual fluids. During a max power lightweight takeoff, the hopper is partially depleted during climbout until the aircraft reaches 25,000 ft. (The climbout in this scenario lasts approximately five [5] minutes, and the hopper is depleted by 15%.) When the aircraft exceeds 25,000 ft or throttles are reduced, the hopper will steadily replenish to its 100% value. The hopper will remain full if the aircraft is in straight and level flight at any altitude because engine consumption in straight and level flight does not exceed ejector pump supply. The aft wing spar serves as the mounting surface for the boost pumps, fuel shutoff valves, the crossflow and intertank valves and temperature sensors. The outside of the aft wing spar is accessible from the main landing gear wheel wells and the fuel system components in the hoppers are mounted through openings in the spar. This design allows the components to be replaced without emptying the fuel tanks. Each hopper contains up to 190 U.S. gallons or 1,283 pounds of fuel (the international equivalents are 719 liters or 583 kilograms). The intake lines supplying the boost pumps are installed along the bottom of the hoppers ensuring that all possible fuel can be extracted from the tanks. Warming the fuel in the hopper and cooling the hydraulic fluid is accomplished by a heat exchanger located and housed within the bottom of the wing tank hopper. Left and right tanks cool their counterpart's system fluid with hydraulic system lines (feed and return) routed through the aft wing beam. Heat exchangers are contained in baffled coiled tubing and function on the same principle as radiators to transfer heat from the hydraulic fluid to fuel. Heat is generated by thermal dynamics via pressurization of the hydraulic fluid; generated heat is then transferred to fuel. Since fuel is cooled by transference of low temperature across the wing surface from high altitude flight, the exchange of heat is essential for fuel viscosity and to preserve pressure transmission of the hydraulic fluid. Over-Wing (Gravity) Fueling System: An over-wing gravity fueling adapter (filler cap) assembly is installed on the top of each wing forward and inboard of the wing flaps. The adapter has a locking fuel cap and a sleeve intake into the wing tank. The sleeve is fitted with a screen filter to prevent foreign objects from entering the tank during fueling. A grounding jack is installed on the wing trailing edge near the over-wing opening for prevention of electrical sparks or shorts during the fueling process. As fuel is pumped into the tank from a truck or underground facility, the fuel flows inboard through the baffles in the wing ribs to the hopper at the lowest point of the tank. The tank is filled from inboard to outboard and fuel quantity must be monitored at the fuel truck or hose outlet at underground tanks unless the airplane is powered and manned, in which case the airplane fuel gages may be used to monitor quantities. If the airplane gages are used to determine fuel quantity, the airplane should be level, otherwise an imbalance between wing tanks will occur. Both wing tanks should be filled simultaneously. If both tanks cannot be filled at the same time, fueling should alternate from wing tank to wing tank to avoid the maximum fuel imbalance limit of 5,000 pounds. (See Figure 3. Fueling Adapters and Drains.) Procedures for over-wing (gravity) fueling are presented in Chapter 9: Handling and Servicing. The maximum fuel load from over-the-wing gravity servicing is 43,650 lbs (19,799 kg), though attitudes outside nominal ground attitude (-1 degree pitch, 0 degrees roll) may reduce achievable fuel load to as little as 39,000 lbs (17,690 kg). Gravity Water / Fuel Drain System: Each wing tank is equipped with three drain valves located in the lower skin of each wing. Drain valves are provided for each tank: one in the main portion of the wing tank, one in each hopper, and one in each vent plenum. This design allows replacement of the seal without draining the tank, should conditions warrant seal replacement. The drains are operated manually and serve to drain any water that has infiltrated the airplane fuel due to rain seepage or contamination. (Any water would accumulate at the bottom of the fuel tank since the specific gravity of water is heavier than the specific gravity of petroleum-based fuels.) (See Figure 3. Fueling Adapters and Drains.) Fuel Ventilation System: The fuel ventilation system consists of a system of valves and tubes that allow air into the aircraft fuel tanks as fuel is being consumed and keeps fuel from venting overboard under normal operations. If fuel tank pressure exceeds the limit specified, the vent system will open and allow excess fuel to vent overboard. Venting of fuel is accomplished by a single pressure relief valve located internally on the vent stringer on the inboard area of the wing. Several non-pressure relief valves are located throughout the fuel tank to minimize air pockets within the tank. Operation of the venting system is fully passive and does not require aircraft power for operation. There is a fail-safe pressure valve that will open before the wing reaches overpressure limits in the event of valve failure. Two vent stringers are part of the wing structures. These stringers are designed to run from the vent plenum to the inboard portion of the wing, providing an air entry path and fuel exit path during overpressure conditions. The ventilation system spans the length of each wing fuel tank. At the wing tip, just outboard of the fuel tank, is an empty chamber or plenum, plumbed to an air vent scoop on the lower section of the outboard wing leading edge. Two vent tubes connected to the plenum run the length of the fuel tank at the top, just below the wing upper skin. The inboard ends of the vent tubes are joined together at a common manifold just forward of the fuel hopper. The common manifold is connected to two vent / float valves (one inboard, one outboard); both being 1.75" in diameter. When the inboard section of the fuel tank is not full of fuel, the hinged floats of the valves drop down, providing an air passage from the inboard section of the fuel tank out through the vent tubes to the plenum at the wing tip and out to the ram air vent on the wing leading edge. If the inboard section of the tank is to be filled with fuel during servicing (inboard fuel levels are dependent upon total fuel carried in the tank), the hinged float of the valves will remain open to permit the flow of air out of the inboard section of the tank until the rising fuel level pushes the float closed. Since inboard or lower section of the wing tank is filled first, another set of vent / float valves positioned adjacent to the air plenum at the wing tip are necessary to continue venting the wing during refueling. Each of the two vent / float valves is connected to one of the vent tubes along the top of the wing, and both act in the same manner as the two inboard vent / float valves. They provide a passage for air to exit the tank as fuel fills the tank. The outboard location of the vent / float valves enables air ventilation until the tanks are full, at which time the floats of the valves close. The four vent / float valves operate in the opposite manner as fuel is consumed from the tank. As the airplane takes off with full fuel tanks, the engines are supplied with fuel at a high rate by the boost pumps drawing fuel from the bottom of the tank hopper and pressurizing the fuel into the engine supply lines. The two inboard vent / float valves incrementally open to admit air to the tank as fuel is withdrawn from the tanks by the boost pumps. Forward airplane speed forces air into the vent scoop on the wing leading edge and the air pressure flows through the plenum at the wing tip, inboard through the vent tubes and to the vent / float valves, assisting in opening the valves. As fuel is consumed from the inboard section of the tank, fuel in the outboard section of the tank flows inboard through the rib baffles towards the hopper as space becomes available. As fuel levels in the outboard section of the wing fall, the two outboard vent / float valves open to provide additional air to fill the tank space vacated by the fuel. Airplane turns and banks after takeoff will result in either inboard or outboard vent / float valves opening or closing depending upon fuel levels and airplane attitudes but ventilation air will remain available. The ventilation system also provides a secondary function for the fuel tank. If a fueling malfunction overfills the tank, the pressure of the increasing volume of fuel will overcome the hinge action of the vent / float valves forcing them open. Fuel will then flow into a 3" spring-loaded pressure relief valve and out the vent tubes to the plenum chamber. If fuel continues to over-pressurize the tank, the plenum will fill with fuel and then spill out of the ram air inlet in the wing leading edge, providing a relief from the overpressure and preventing structural damage to the wing tank. The outboard vent plenum also acts as an expansion chamber for wing fuel. If the airplane fuel tanks are filled and the airplane subsequently is exposed to high temperatures and/or solar heating prior to departure, the fuel volume within the tanks will expand even though the weight of the fuel remains unchanged. The expanding fuel will force open the vent / float valves and expand out to the plenum that provides sufficient space for at least a 2% increase in fuel volume. If circumstances result in fuel filling the vent system, a drain valve in the bottom of the plenum can be opened to empty the plenum. Any fuel in the plenum should be drained prior to takeoff to prevent environmental damage from fuel spilling from the ram air opening. Any fuel remaining in the vent tubes will drain back into the fuel tank through three vent / float drains located on the underside of the tubes and positioned within the inboard 1/3 of the fuel tank. Fuel Outlet Strainer System: Three mesh strainers prevent contaminants from entering the fuel system. Over the wing fueling is protected from contaminants by one filter and the remaining two filters are mounted to the intakes of the fuel boost pumps, near the bottom of the fuel hopper. These meshes are able to prevent solid contaminants smaller than 0.25" from entering the fuel system but cannot prevent foreign liquid agents from entering the system. Procedural guidelines for fueling operations ensure liquid contaminants are not introduced to the fuel system. Controls and Indications: Crew Alerting System (CAS) Messages: The following CAS message is displayed if the fuel quantity in the two wing tanks differs by 1,000 pounds: Area Monitored CAS Message Fuel Quantity inputs to MAU #1 and #2 Fuel Imbalance (advisory) Limitations: Pressure Refueling Capacities: When pressure refueling, the usable fuel capacities for G650ER S/N 6001 and subsequent are: Right Tank 24,100 lb (10,931 kg) 3,597 gal (13,616 lit) Left Tank 24,100 lb (10,931 kg) 3,597 gal (13,616 lit) Total 48,200 lb (21,863 kg) 7,194 gal (27,233 lit) It may be possible to upload fuel in excess of 48,200 lb (21,863 kg). This is permitted as long as the maximum ramp weight and / or the maximum takeoff weight is not exceeded, and the loaded airplane center of gravity is within limits. If either fuel tank quantity exceeds 25,100 lb (11,385 kg), the Fuel Quantity digital readout on the Engine Instruments and synoptics displayed on the Display Unit(s) will have amber dashes on the affected side(s) and the total fuel quantities. Accurate fuel quantity values are available on the FMS and the REFUEL page within the UTILITY function of the SMC. Gravity Refueling Capacities: When gravity refueling, the total useable fuel capacity for G650ER is 43,650 lb (19,799 kg) / 6,515 gal (24,661 lit). Figure 3. Fueling Adapters and Drains OVERWING (GRAVITY) FUELING ADAPTER (2 PLACES) TOP VIEW FUEL / WATER GRAVITY DRAINS (6 PLACES) FUELING COMPARTMENT ACCESS DOOR BOTTOM VIEW TIL-003216 2A-28-30: Fuel Distribution System General Description: Fuel distribution for airplane operation involves several distinct features directed toward supplying the engines with the correct amount of usable fuel. The airplane tanks may be filled using a single point pressure fueling system. If fuel quantity needs to be reduced (or the tanks emptied), three methods are available for removal of fuel from the wing tanks. Fuel within the tanks is supplied to the engines by pressurizing the fuel with tank boost pumps. To compensate for an in-flight engine failure, either engine can be supplied with fuel from either tank. Fuel balance between the tanks is accomplished though opening a valve common to both tanks. The fuel system includes a means to maintain the temperature in the fuel tanks within a specified range. Fuel distribution incorporates the following subsystems (see Figure 4. Cockpit Fuel System Control Panel and Figure 5. Ground Service Control Panel): Storage Subsystem / Fuel Storage Single Point Pressure Fueling Defueling Fuel Crossflow and Intertank Transfer Engine and APU Fuel Supply Valves Fuel Boost Pumps Fuel Temperature Indication and Heated Fuel Return Description of Subsystems, Units and Components: Storage Subsystem / Fuel Storage: The storage system provides the necessary storage tanks for the fuel system, along with fill, drain and vent capabilities. Fuel is stored in two wing tanks (left and right). The tanks can be fueled from a single-point pressure fueling adapter or from two over-the-wing (gravity) fuel ports. The tanks are integral to the wing structure and each is subdivided in seven compartments communicated by flapper valves to restrict fuel "slosh" during maneuvers and to allow gravity (over-the-wing) refueling. The total usable fuel in the fuel storage system is 48,200 pounds (approximately 6,558 U.S. gallons). Each wing contains a surge tank where fuel entering the vent system is caught and drawn back into the main tank. Each wing tank has a fuel hopper that is an isolated compartment within the tank. Fuel is removed from the fuel hoppers for each engine and APU. The hoppers are continuously filled by fuel ejector pumps and gravity flow. The fuel tanks can be filled from a single-point pressure fueling adapter or through tank fill openings on top of each wing. Fuel is drawn from the tanks and pressurized for distribution to the airplane engines and/or APU by boost pumps mounted on the rear spar wall of each tank. Sensors mounted within the fuel tanks provide information for cockpit display windows enabling the flight crew to monitor fuel quantity and fuel tank temperature. Tank ventilation provides sufficient venting while the aircraft is on the ground and during flight. In flight, the ventilation system slightly pressurizes each wing tank. For cooling hydraulic fluid, a hydraulic fluid heat exchanger is installed in each hopper. A total of six water drain valves are incorporated on the wings at the lowest point of each tank (main, collector, surge). Their purpose is to drain accumulated water from the tanks and to complete the removal of residual fuel when accomplishing a complete aircraft defuel. Single Point Pressure Fueling: Pressure refueling is via a single adapter on the right side of the aircraft just forward of the right wing. Refueling is controlled automatically by a fuel load pre-selector taking its signal from the fuel gauging system. Fuel load can be entered from the flight deck Standby Multifunction Controller (SMC) or the Ground Service Control Panel (GSCP) near the refueling port. Overwing filler caps are provided in each wing for gravity refueling. The caps are positioned such that the wing cannot be filled to greater than certified capacity. Both left and right fuel tanks may be filled from a single fueling point within a panel located on the right side of the aircraft just forward of the right wing. When opened, the panel face swivels forward to gain access to the fueling receptacle, and contains the printed operating instructions for pressure fueling, a red high level warning light and the GSCP. Due to variations in fuel source pressure and system response time, the automatic refuel may not stop exactly at the preselected quantity. Up to a +/-200lb difference between the preselected quantity and the final quantity is normal. Optimum fuel pressure at the fueling nozzle (fuel flowing) is from 35-55 psig. If the pressure is below 35 psig, then the refuel may fail, and LEFT (OR RIGHT) REFUEL VLV CLSD FAULT will be indicated in the CMC. In this situation, the FQSC must be reset to clear the LEFT (OR RIGHT) REFUEL VLV CLSD FAULT before a subsequent refuel can be attempted. The GSCP controls allow the required total fuel load to be preselected and automatically controlled by the Fuel Quantity Signal Conditioner (FQSC). Overfilling of the tanks is protected by the gauging system by shutting off refueling valves when maximum fuel quantity is achieved. An automated test will be performed using the panel to check the operation of the refuel shutoff control. This allows the refueling valves to be actuated in order to verify their closure. Arming of the refuel system is controlled by means of a toggle switch on the refuel panel. After the desired fuel quantity has been selected with the GSCP, the shutoff control features of the pressure fueling system are operationally checked prior to fueling. During pressure fueling, the truck nozzle is attached and locked to the airplane adapter and fuel is routed from the adapter into a manifold flowing into each tank. Within each tank the fuel passes through the pressure fueling shutoff valve and into the tank, filling the tank from inboard to outboard. The valve may be closed by electrical signals from three sources: The fuel quantity system when the preselected total fuel quantity has been reached. The GSCP when the left TEST/RESET and AUTO REFUEL switch is placed in the center OFF position. The REMOTE FUELING SHUTOFF pushbutton above the FUEL SYSTEM panel on the Cockpit Overhead Panel (COP) (see Figure 4. Cockpit Fuel System Control Panel). The Ground Service Bus is turned OFF at any one of the switch locations. If the valve is closed via electrical signals from any of these sources, pressure will build up in the fuel control line, which closes the fueling shutoff valve. The cockpit control switch is then checked: The SHUTOFF pushbutton on the REMOTE FUELING section of the FUEL SYSTEM panel is depressed, resulting in the illumination of the blue CLSD legend within the pushbutton. The flow of pressurized fuel into the tanks is interrupted by electrically closing the solenoid valve in the control fuel line, and the resulting increase in pressure will close the fueling shutoff valve. The pushbutton is then selected to the open position (legend in the pushbutton extinguished) and the flow of fuel into the left tank resumes. Once all pressure fueling shutoff features have been successfully tested, pressure fueling can be completed, loading fuel on board the airplane until reaching the preselected level. When fuel reaches the preselected level, an electrical signal is sent by the GSCP to close the solenoid valve. If an unlikely series of malfunctions occurs and none of the shutoff features of the pressure fueling system stop the flow of fuel into the airplane tanks, a high level sensor in each fuel tank plenum will signal the illumination of the red warning light on the inboard face of the fueling panel. The pressure fueling operator should then immediately stop the flow of fuel into the airplane by using the control lanyard or emergency shutoff at the fuel truck. Defueling: Removal of fuel from the airplane may be accomplished by three methods. Depending upon the circumstance, a combination of the methods may yield the most satisfactory results. The methods are: Attaching the pressure fueling nozzle into the adapter at the fueling panel and applying suction instead of pressure at the fueling truck. Two defueling lines (one for each tank) are connected to the pressure fueling inlet lines, but separated from the lines by one-way check valves that remain closed under fueling pressure. When suction is applied to the fueling inlet lines, the check valves open and fuel is siphoned from the airplane. The inlets of the defueling lines are located on the bottom of the fuel hoppers at the lowest point of the fuel tanks in order to remove the maximum amount of fuel possible from the tanks. However, approximately 11 gallons (42 liters) of fuel will remain within the tanks using this method. The water / fuel drains must be opened to remove residual fuel if the tank is to be totally emptied. Suction defueling is the common method of reducing the amount of fuel on board the airplane if operational circumstances require a reduction in airplane takeoff gross weight. Connecting a one inch hose from the drain fitting on the fuel supply line on the right side of the aircraft and using fuel truck suction to draw fuel through the boost pump intake lines. This method results in less residual fuel in the tanks, but is much slower, since a one inch line is used, and only one tank at a time may be emptied. Using the same one inch line connected to the engine fuel line drain fitting and fuel truck, but powering the boost pumps to pressurize the engine supply line. This method is faster than method number two, and leaves the least amount of fuel in the tanks, since the boost pumps operate the ejector pump to induce greater fuel scavenging. Prior to defueling the airplane into a fuel truck, ensure that the capacity of the truck will accommodate the amount of fuel to be removed from the airplane. Fuel Crossflow and Intertank Transfer: During normal flight operations, each engine consumes fuel from the respective side tank (i.e. left engine from left tank); however, malfunctions may require different fuel feed requirements. If both boost pumps in one tank fail, the engine corresponding to that tank must be fed from the opposite tank. Feeding both engines from one tank will cause an imbalance in the quantity between tanks, and if not corrected, exceed the airplane limitation of 2,000 lbs between tanks. The same imbalance would occur at a slower rate if one engine fails and the remaining operating engine is fed only from the tank corresponding to that engine. Two sets of lines and valves are installed between the fuel tanks to provide alternate fuel feed paths to maintain fuel flow to the engines while maintaining fuel balance between tanks. The crossflow line and valve is plumbed between the manifolds housing the two boost pumps within the hoppers of each tank, and allows an engine to be supplied with pressurized fuel from the opposite side tank. In the event of an engine failure, the remaining engine can be fed alternately from the onside fuel tank, then from the tank on the side of the inoperative engine to maintain fuel balance. Feeding an engine from the opposite side tank is accomplished by first opening the crossflow valve, turning on the boost pumps in the opposite side tanks, and then turning off the boost pumps in the onside tank. The crossflow valve can also compensate for the loss of both boost pumps within a tank. The valve is selected open with the X FLOW pushbutton on the FUEL SYSTEM panel on the COP. When the pushbutton is depressed, the white line legend within the switch is illuminated, completing the illustrated diagram between the tanks shown on the panel. The crossflow valve is powered by the Left Essential DC bus; valve position is monitored by MAU #1 and MAU #2. When the valve is open, MAU #1 relays the position to the Monitor and Warning System (MWS) that in turn generates a Crossflow Valve Open (advisory) CAS message and illustrates the direction of fuel feed with a green arrow on the Fuel synoptic display window (the green arrow will point towards the tank with the least number of operating boost pumps). If the crossflow is left open more than five minutes, the Fuel Crossflow Valve Open will turn amber in color. Likewise, the Crossflow valve on the fuel synoptic page will also turn amber. The five minute timer mentioned above can be reset by first ensuring that at least one fuel boost pump is ON for each engine, then cycling the Crossflow CLOSED and back OPEN again. To enable balancing fuel between tanks in the event of dual boost pump failure within a tank, an intertank line and valve are installed between the two fuel tank hoppers. The valve is opened with the INTER TANK pushbutton switch on the FUEL SYSTEM panel on the COP. Depressing the switch will open the valve and the white line legend will illuminate within the switch completing the diagram line between tanks shown on the panel. The intertank valve is also powered by the Left Essential DC bus, but is position monitored by MAU #2. When the valve is open, a Tank Valve Open (advisory) CAS message is displayed, and the valve is illustrated open on the Fuel synoptic window display. Opening the intertank valve only provides a path for fuel flow through the line connecting the hoppers of the fuel tanks. Since boost pump operation cannot provide the motive force to transfer fuel from tank to tank, the fuel balancing procedure requires that the airplane be flown slightly out of trim in the yaw axis to induce fuel migration through the intertank valve. The out-of-trim condition will produce sufficient lateral force to cause fuel from the tank with the greater quantity to flow into the tank with the lesser quantity. For a full description of the procedures for operation of the crossflow and intertank valves, see section 05-14-00: Fuel System Abnormal / Emergency Procedures. Engine and APU Fuel Supply Valves: The engines are supplied pressurized fuel from the manifolds housing the dual boost pumps in each tank. The fuel is routed through a shutoff valve located on the aft wing beam at each tank. The shutoff valves are powered by the Essential DC buses, Left Essential DC for the left engine and tank and the Right Essential DC for the right engine and tank. The operation of the valves is controlled by the engine fire handles on the forward section of the cockpit center pedestal. Pulling out a fire handle will close the shutoff valve, stopping fuel flow to the selected engine. The APU receives pressurized fuel from the left boost pumps and tank. A shutoff valve is installed in the APU fuel line at the left tank and is accessible through the left main wheel well. The valve is controlled by the APU MASTER switch on the COP and is also powered through the relay in the switch, via the Right Battery bus or Left Essential DC bus. The shutoff valves for the engines and APU are accessible from the wheel wells and can be removed without defueling the airplane. Fuel Boost Pumps: The airplane fuel tanks are equipped with four identical and interchangeable fuel boost pumps, two in each tank. The boost pumps are located within the tank hoppers to ensure a positive supply of fuel to the pumps. The boost pump intakes are covered with filter screens to prevent the ingestion of foreign objects or particles that could damage the pumps. Boost Pump Select Switches: Four boost pump select switches located on the Cockpit Overhead Panel (COP) provide operational control of the boost pumps. These switches are labeled as L PUMPS MAIN and ALT, and the R PUMPS MAIN and ALT. When a switch is depressed to the ON position, its associated control contactor actuates to provide power to the boost pump, turning it on. When the switch is depressed ON, the OFF label on the switch light extinguishes. When the switch is released to the extended position, the switch legend illuminates OFF. Boost Pump Control Contactors: The boost pump control contactors, located in the left and right Power Distribution Boxes (PDBs), provide power (on / off) control to the boost pumps. When actuated, power from a 25 amp circuit breaker source is connected directly to the boost pump. One of the contactor's auxiliary switches provides a 28V DC signal to the Annunciator Lights Dim and Test Controller that enables the illumination of the OFF legend on the on / off control switch. A second auxiliary switch provides a 28V DC signal to the MAU to indicate which boost pump is selected ON. The information is used to display boost pump status on the Fuel synoptic page. Boost Pump Power Source Left Main Left Essential DC Left Alternate Left Main DC Right Main Right Essential DC Right Alternate Right Main DC The pressure produced by the boost pumps is also used to provide fluid flow through the ejector pump in each tank. The ejector pump incorporates a small diameter line from the pressurized boost pump manifold plumbed to extend forward in the tank to a position in front of the intake baffle to the tank hopper. The ejector directs a stream of high pressure fuel into the mouth of a wider opening plumbed back into the hopper. The velocity of the fuel ejected from the pump induces the flow of a larger volume of fuel into the hopper, thus assisting in the movement of fuel into the boost pump intakes. Should both boost pumps fail, suction feed flows through the pumps, enabling the engine to siphon fuel from the bottom of the hopper using the engine-mounted fuel pumps. If the crossflow valve is open, boost pump pressure from the opposite side tank will be higher than engine suction pressure and the engine will be fed from the opposite tank. Fuel Temperature Indication and Heated Fuel Return: The temperature of the fuel in the tanks is monitored to ensure that it remains within limitations. Two Fuel Characteristic Sensors (FCSs), one in each tank at BL 16, provide a measurement of fuel density, fuel permittivity and temperature. Two Fuel Temperature Sensors (FTSs), one in each tank in the middle of Rib Station 242.66, are used for fuel temperature measurement within the fuel tanks. The sensors contain an element with an electrical resistance that varies with temperature. The resistance in the sensors is monitored by the MAUs (left tank by MAU #1, right tank by MAU #2) and converted to a proportional voltage that is then translated into a digital format for display by the MWS. The temperature display range is from -70°C to +300°C with an accuracy of ±1°C. Fuel tank temperature is displayed on the Fuel synoptic 2/3 window, Summary synoptic 2/3 window and the Secondary Engine 1/6 window displays. When operating at high altitudes or in extreme latitudes, the temperature of the airplane fuel supply must be increased in order to ensure that fuel viscosity remains low enough for the fuel to flow freely through tank components and the engine Fuel Metering Unit (FMU). Fuel in the wing tanks is warmed by a Heated Fuel Return System (HFRS) that diverts a part of the fuel supplied to the engine back to the tanks. The fuel returned to the tanks is drawn off after it has passed through a heat exchanger with the engine oil system where hot oil is cooled by engine fuel. As a result, the temperature of the fuel is raised to approximately 50°C. The hot fuel is returned to the fuel tanks and distributed throughout the wing by a system of pipes with multiple small holes. Although the volume of hot fuel returned is much smaller than the cold fuel within the tank, the temperature difference is significant enough to warm the tank fuel to within the range to maintain required fuel viscosity. The flow of hot fuel back to the tank is controlled by three switches: one on the COP and one at each engine. The FUEL RETURN OFF / AUTO pushbutton switch on the overhead FUEL SYSTEM panel provides flight crew control of the HFRS. When selected to the AUTO (or ON, switch light is DARK) position, the return of fuel from the engine is managed by the Full Authority Digital Engine Control (FADEC) on each engine. The engine FADECs receive fuel tank temperature from the temperature probes fitted into each wing tank through the aft wing beam at the hopper. If the temperature within a wing tank is :50°C, the FADEC will open a Fuel Return To Tank (FRTT) valve at the engine, porting hot fuel back to the tank. When tank fuel temperature rises 210°C, the FADEC will close the valve, retaining the fuel within the engine FMU circulation. The operation of the FRTT valves are monitored by the MAUs, with MAU #1 communicating with the cockpit overhead switch and the FRTT valve of the left engine, and MAU #2 linked to the FRTT valve of the right engine. HFRS status is reported by the MAUs to the MWS that in turn formulates CAS messages appropriate to the operating condition and generates the graphic display of HFRS operation on the Fuel synoptic window. The Fuel synoptic window will display the HFRS only when the system is operating. The system is represented by a line and a valve between the respective tank and engine, displayed in white when not in operation, and in green when in operation . If the system is in AUTO and selected ON by the FADEC and a malfunction exists, the line and valve are represented in amber. The operation of the HFRS is limited to fuel system normal operating parameters and also by airplane performance requirements, since the fuel returned to the tank must be in excess of the flow consumed by the engines as scheduled by the power levers and FADEC. For the FADEC to open the FRTT valve, the following conditions must be valid: The cockpit overhead switch selected to AUTO The engine fire handle stowed (not pulled) Engine low fuel pressure not indicated Engine low fuel quantity not indicated Engine fuel flow requirement less than 2,650 pounds per hour Fuel filter not blocked Fuel crossflow valve closed Fuel temperature should be at or below 0°C Electrical power to the FRTT valve available Controls and Indications: (See Figure 4. Cockpit Fuel System Control Panel.) Circuit Breakers (CBs): The following CBs protect the fuel distribution system: Circuit Breaker Name CB Panel Location Power Source L MAIN FUEL PUMP LPDB LEER Left Essential DC R MAIN FUEL PUMP RPDB REER Right Essential DC L MAIN PUMP CONT POP C-7 L ESS 28V DC R MAIN PUMP CONT CPOP C-7 R ESS 28V DC ALT FUEL PUMP L MCDU SSPC (#2808) L MAIN 28V DC Circuit Breaker Name CB Panel Location Power Source ALT FUEL PUMP R MCDU SSPC (#2809) R MAIN 28V DC FUEL X-FLOW VLV MCDU SSPC (#2813) L ESS 28V DC FUEL INTERTNK VLV MCDU SSPC (#2810) L ESS 28V DC FUEL S/O VLV L MCDU SSPC (#2811) L ESS 28V DC FUEL S/O VLV R MCDU SSPC (#2812) R ESS 28V DC FUEL RETURN L MCDU SSPC (#2805) L MAIN 28V DC FUEL RETURN R MCDU SSPC (#2806) R MAIN 28V DC Crew Alerting System (CAS) Messages: The following CAS messages are associated with the fuel distribution system: Area Monitored CAS Message Pressure in boost pump manifold L-R Fuel Pressure Low (warning) Fuel tank temperature below -37°C or above +54°C Fuel Tank Temperature (warning) Alternate fuel boost pump pressure L-R Alt Fuel Pump Fail (caution) Boost pump control logic (only one pump on with crossflow valve open above 41,000 ft.) Fuel Boost Pump (caution) Fuel tank temperature between -35°C and -36°C or tank temperature above -35°C at an altitude above 35,000 ft. with fuel quantity indicating more than 1,000 lbs in either or both tanks, but only 1,000 lbs. remaining in hopper Fuel Tank Temperature (caution) Main fuel boost pump pressure L-R Main Fuel Pump Fail (caution) Crossflow valve Fuel Crossflow Valve Open (caution) NOTE: This message is displayed after the crossflow valve has been open for five (5) minutes. Area Monitored CAS Message Crossflow valve Fuel Crossflow Valve Open (advisory) Fuel quantity differs between tanks by more than 1,000 lbs. Fuel Imbalance (advisory) Intertank valve Fuel Inter Tank Valve Open (advisory) Heated fuel return L-R Fuel Return Fail (advisory) Limitations: Boost Pumps: All operable boost pumps must be selected ON for all phases of flight unless fuel balancing is in progress. Fuel Tank Temperature: Maximum: Fuel temperatures of +54°C or greater cause a Temperature (warning) CAS message to be displayed. Minimum: Caution Range: Fuel temperatures of -34.5°C to -37°C will cause a Temperature (caution) CAS message to be displayed. Warning Range: Fuel temperatures less than -37°C cause a Temperature (warning) CAS message to be displayed. In-Flight Fuel Tank Temperature at or Below -30°C With Less Than 5000 lb (2268 kg) Total Remaining: When fuel tank temperature is at or below -30°C in flight with less than 5,000 lb (2268 kg) total fuel remaining, the airplane shall be descended to an altitude where SAT is -60°C or warmer and maintained at a minimum speed of Mach 0.80. Figure 4. Cockpit Fuel System Control Panel FUEL RETURN OFF / AUTO When selected on, FADEC controls the fuel return shutoff valve to control tank temperature. When selected OFF, the shutoff valve closes and the OFF legend illuminates. REMOTE FUELING SHUTOFF When selected to CLSD, the pressure fueling shutoff valves close and the CLSD legend illuminates. When selected to off, the pressure fueling shutoff valves open and the CLSD legend extinguishes. FUEL RETURN REMOTE FUELING INTER TANK When selected on, the intertank valve opens OFF/AUTO SHUTOFF FUEL SYSTEM INTER TANK R TANK and a horizontal dash legend illuminates. When selected off, the intertank valve closes and the horizontal dash legend extinguishes. L PUMPS R PUMPS ALT MAIN MAIN ALT OFF OFF OFF OFF APU X FLOW R ENG X FLOW When selected on, the crossflow valve opens and a horizontal dash legend illuminates. When selected off, the crossflow valve closes and the horizontal dash legend extinguishes. L (or R) MAIN (or ALT) PUMPS When selected on, the respective boost pump operates. When selected OFF, the respective boost pump shuts off and the OFF legend illuminates. TIL-001404 Figure 5. Ground Service Control Panel TIL-001405 2A-28-40: Fuel State Indication General Description: The fuel state indication system is a complete two channel system that provides fuel quantity measurement, fuel low and high level monitoring and fuel tank temperature indications. The indication system provides monitoring of the following conditions: Amount of fuel remaining in the left and right wing tanks Total fuel remaining Fuel temperature in the hopper Each wing fuel tank contains 14 standard tank gauging probes, one dual tank gauging probe (hopper), one Fuel Characteristics Sensor (FCS), one Fuel Temperature Sensor (FTS), one high-level sensor and two interconnect wiring harnesses. Fuel quantity in each tank is measured by a transistorized, capacitance type sensor. The independent sensors supply separate signals to the fuel quantity processor. The fuel quantity is shown in pounds on the Fuel synoptic page. The low fuel level sensors installed in each hopper supply a signal to turn on the L-R Fuel Level Low (caution) CAS message under either of the following conditions: Hopper fuel probe (1A side) senses a fuel quantity of less than 650 lb (-100 gal) Hopper fuel probe (1B side) wetted height is less than 303 mm A fuel temperature sensor installed in each hopper supplies signals to the fuel tank temperature display on the Fuel synoptic page. Airplanes configured in accordance with JAR specifications have the option of displaying fuel quantities in kilograms (Kg.) WHEN REFUELING THE AIRPLANE AT LOCATIONS WHERE THE FUEL VENDOR MAY BE UNFAMILIAR WITH AIRPLANE REQUIREMENTS, REVIEW AIRPLANE FLIGHT MANUAL / LIMITATIONS / AIRPLANE SERVICING PRIOR TO FUELING: Description of Subsystems, Units and Components: Ground Service Control Panel (GSCP): The GSCP (see Figure 6. Single Point Refuel and GSCP), is externally mounted in the fuel service panel. The GSCP is designed to receive power when the ground service bus is selected ON. The GSCP has three amber LED display modules capable of displaying the full ASCII character set. The upper display indicates the left tank fuel quantity and left high level warning annunciator. The middle display indicates the right tank fuel quantity and right high level warning. The lower display indicates the preselected refuel quantity, the unit of measurement (lbs/kg) or various fault messages. The fuel quantities are displayed in 25 lbs/10 kg increments. The preselect quantity can be set in 100 lbs / 50 kg increments. Fault display is controlled by the INC / DEC switch. Standby Multifunction Controller (SMC): The SMC provides a soft menu interface for refuel control from the flight deck. Two SMCs are installed; one in the pilot glare shield and one in the copilot glare shield. Both the pilot and copilot SMCs provide a menu for fuel quantity display and can be used for refuel control by being able to set REFUEL quantity. Fuel Quantity Signal Conditioner (FQSC): The FQSC provides dual-channel independent left tank and right tank fuel quantity processing after receiving the fuel gauging information from the four Remote Data Concentrators (RDCs) located at the fuel tank. The FQSC then transmits two-channel ARINC 429 data to the Honeywell system for flight deck display. Data is also transmitted to the GSCP on a bidirectional Controller Area Network (CAN) bus. The channels are monitored by the Modular Avionics Unit (MAU) and provide an Channel Fail (advisory) CAS message to the flight deck display in the event of a channel failure. Discrete signals provided by the FQSC include those to close the fueling shutoff valves and a signal to actuate the externally-mounted Fuel High Level warning light. Configuration pin strapping is provided to determine fuel load in pounds or kilograms. No FQSC or tank sensor calibration is required at aircraft installation, or when components are replaced. Extensive Built-In-Test (BIT) is provided in three operating modes: power up, continuous background and initiated. BIT checks all system functions including tank sensors, FQSC, GSCP and RDCs for proper operation and reasonable data. Remote Data Concentrator (RDC): The fuel tank sensor information is input to the microprocessor based-RDCs, mounted on the front spar of the wing tank. Each RDC includes all circuits required to connect to the gauging probes (to condition the signals coming back from them) and to transmit the information to both FQSC channels through a fiber optic CAN data bus. The 30V DC power supply to the RDCs is provided by the FQSC. Fuel Indication System: The fuel indication system provides the flight crew with a continuous indication of the amount of fuel in each wing tank (left or right) and the total fuel (left and right wing fuel tanks combined). The system also provides continuous indication of left and right wing fuel tank temperature. Fuel quantity in each tank and total quantity of fuel on board the aircraft are displayed on the Fuel synoptic page. Fuel system indication is provided on the 2/3 fuel system display. Fuel system information can also be accessed from the following pages: Summary Ground Service Compacted Engine MCDU #1 Backup Engine Controls and Indications: (See Figure 6. Single Point Refuel and GSCP.) Full descriptions of synoptic and system window displays are presented in Section 2B-07-00. The fuel quantity system may be tested with the FUEL pushbutton switch on the SYSTEM TEST panel on the upper left section of the cockpit overhead. Depressing this pushbutton will illuminate the amber TEST legend in the switch and result in the following indications being displayed on whichever of the above listed synoptic or system pages are currently selected for viewing on the Multi-function Control and Display Units (MCDUs): A fuel quantity of 7,000 lbs in each tank (left and right) and a total fuel quantity of 14,000 lbs Fuel tank temperatures (left and right) of -10° C Circuit Breakers (CBs): The following CBs protect the fuel indication system: Circuit Breaker Name CB Panel Location Power Source FUEL QTY L (AIR) MDAU SSPC (#2801) L DC MPT 2 FUEL QTY R (AIR) MDAU SSPC (#2803) R DC MPT 1 FUEL QTY L (GND) MDAU SSPC (#2802) R DC MPT 1 FUEL QTY R (GND) MDAU SSPC (#2804) R DC MPT 1 Crew Alerting System (CAS) Messages: The following CAS messages are associated with the fuel indication system: CAS Message Possible Cause(s): L-R Fuel Pressure Low (warning) Fuel pressure is less than 16 psi or both fuel boost pumps on one side have been selected to OFF with X FLOW valve CLOSED. Fuel Tank Temperature (warning) Temperature is below -37°C or above +54°C L-R Fuel Level Low (caution) Fuel level in hopper is 650 lbs (295 Kg) or less Fuel Tank Temperature (caution) Fuel tank temperature is -34.5°C to -37°C OR With fuel tank temperature warmer than -34.5°C, altitude above 35,000 ft and fuel quantity indicating more than 1000 lb in either or both tanks, there is only 1000 lb of fuel remaining in the hopper. L-R FQMS Degrade (advisory) An internal fault within the identified Fuel Quantity Measuring System has caused degraded accuracy. FQMS Maintenance Required (advisory) Fuel Quantity Measurement System maintenance is required. NOTE: Failure of the FQSC information is transmitted from the FQSC to the Flight Deck via ARINC 429 bus and triggers the FQMS MAINTENANCE REQUIRED blue CAS message. FQSC Channel Fail (advisory) A channel Fuel Quantity Signal Conditioner is failed. CAS Message Possible Cause(s): Fuel Imbalance (advisory) Fuel Imbalance between tanks has exceeded 1000 pounds. Refuel System Fail (advisory) Fuel system has determined that it is incapable of pressure refueling. Limitations: a. In-Flight Fuel Tank Temperature at or Below -30°C With Less Than 5000 lb (2268 kg) Total Remaining: When fuel tank temperature is at or below -30°C in flight with less than 5000 lb (2268 kg) of total fuel remaining, descend the airplane to an altitude where the Static Air Temperature (SAT) is -60°C or warmer and maintain a minimum speed of Mach 0.80. Figure 6. Single Point Refuel and GSCP SEE DETAIL A DETAIL A TIL-003215