Patent Publication Number: US-2018050812-A1

Title: Aircraft fuel pump systems

Description:
BACKGROUND 
     1. Field 
     The present disclosure relates to fuel pumps, more specifically to aircraft fuel pump systems. 
     2. Description of Related Art 
     Aircraft gas turbine engines receive pressurized fuel from fuel gear pumps. The gear pump must be compact, light-weight, and robust. The gear pump must perform over a wide operational range while providing critical fuel flows and pressures for various engine performance functions. Typically these gear pumps receive rotational power from an accessory gearbox through an input drive shaft. These gear fuel pumps are often oversized in order to satisfy the high-flow, high pressure fuel flow requirements at take-off engine power and/or low-speed windmill starts/re-starts. Subsequently, during the climb and cruise phases of the flight, the fuel flow to the engine is much reduced resulting in unnecessary additional pump power that remains unused. 
     The current practice includes bypassing a significant portion of the pressurized fuel flow past the fuel nozzles and back into the main fuel tanks. This is undesirable from a thermal management perspective and is a waste of energy. This bypassing increases the temperature of the fuel and limits the capability of fuel to be a heat sink. This fuel bypassing also wears out the fuel pumps, thus shortening their operational life, and introduces possible gas (air, oxygen, nitrogen, etc.) entrainment into the fuel. This is undesirable from an operational perspective. 
     Such conventional methods and systems have generally been considered satisfactory for their intended purpose. However, there is still a need in the art for improved aircraft fuel pump systems. The present disclosure provides a solution for this need. 
     SUMMARY 
     An aircraft fuel system includes a first pump system that is mechanically driven and it is in selective fluid communication with a fuel tank and one or more fuel nozzles of an engine. The first pump system is configured to pump fuel to the one or more fuel nozzles in a high flow rate condition and to be starved or nearly starved of fuel in a low flow rate condition. The aircraft fuel system includes a second pump system including an electric motor. The second pump system is in fluid communication with the fuel tank and the one or more fuel nozzles to pump fuel from the fuel tank to the one or more fuel nozzles. The second pump system is driven by the electric motor and is configured to pump flow in both the high-flow rate condition and the low-flow rate condition. 
     The second pump system can include a total-flow pump and a main pump attached to the electric motor, wherein the total-flow pump is configured to boost the main pump and/or the mechanically driven first pump system. The aircraft fuel system can include a first valve configured to shut-off or otherwise limit fuel flow to the first pump system, the first valve positioned between the total-flow pump and the first pump system. 
     The aircraft fuel system can include a heat exchanger disposed between the first pump system and the second pump system. The heat exchanger can be a fuel-oil heat exchanger, for example, that is configured to cool engine oil with the fuel in the aircraft fuel system. 
     The aircraft fuel system can include an ejector pump disposed between the first pump system and the second pump system and configured to evacuate the first pump system in the low fuel flow condition. The ejector pump can include a venturi, for example. 
     The aircraft fuel system can include a system shut-off valve disposed upstream of the fuel nozzle and downstream of the first and second pump systems. The aircraft fuel system can include a throttle valve disposed between the first pump system and the fuel nozzle. 
     The aircraft fuel system can include a controller operatively connected to the electric motor and to one or more sensors disposed in the aircraft fuel system to control the electric motor as a function of output of the one or more sensors. The controller can be operatively connected to one or more valves of the aircraft fuel system to actuate the valves as a function of the output of the one or more sensors. The one or more sensors can include at least one of a flow meter disposed upstream of the fuel nozzle, a pressure sensor disposed downstream of the first pump system, or a pressure sensor disposed downstream of the fuel nozzle. 
     The main pump can include one of a vane pump or gear pump. The second pump system can include a cruise pump, wherein the cruise pump is configured to boost the main pump. The first pump system can include a take-off pump. The total-flow pump, the cruise pump, or the take-off pump can include a centrifugal pump. 
     The aircraft fuel system can include a first heat exchanger disposed between the total-flow pump and the first pump system and a second heat exchanger disposed between the cruise pump and the main pump. The aircraft fuel system can include a system shut-off valve disposed upstream of the first and second pump systems. 
     A method includes adjusting volume of fuel pumped to one or more fuel nozzles in response to a change in fuel demand of an engine. 
     These and other features of the systems and methods of the subject disclosure will become more readily apparent to those skilled in the art from the following detailed description taken in conjunction with the drawings. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       So that those skilled in the art to which the subject disclosure appertains will readily understand how to make and use the devices and methods of the subject disclosure without undue experimentation, embodiments thereof will be described in detail herein below with reference to certain figures, wherein: 
         FIG. 1  is a schematic view of an embodiment of a system in accordance with this disclosure; 
         FIG. 2  is a schematic view of the system of  FIG. 1 , shown in low flow mode, e.g., for cruise and/or startup operations with lower fuel consumption requirements; 
         FIG. 3  is a schematic view of the system of  FIG. 1 , shown in high flow mode, e.g., for take-off or other modes with higher fuel consumption requirements; and 
         FIG. 4  is a schematic view of another embodiment of a system in accordance with this disclosure. 
     
    
    
     DETAILED DESCRIPTION 
     Reference will now be made to the drawings wherein like reference numerals identify similar structural features or aspects of the subject disclosure. For purposes of explanation and illustration, and not limitation, an illustrative view of an embodiment of an aircraft fuel system in accordance with the disclosure is shown in  FIG. 1  and is designated generally by reference character  100 . Other embodiments and/or aspects of this disclosure are shown in  FIGS. 2-4 . The systems and methods described herein can be used to improve efficiency of fuel systems. 
     Referring to  FIG. 1 , an aircraft fuel system  100  includes a first pump system  101  that is mechanically driven (e.g., via a gearbox  103  connected to an input shaft from an engine). The first pump system  101  is in selective fluid communication with a fuel tank  105  and one or more fuel nozzles  109  of an engine. The first pump system  101  is configured to pump fuel to the fuel nozzle  109  in a high flow rate condition (e.g., take-off, climb) and to be starved or nearly starved of fuel in a low flow rate condition (e.g., cruise, descent, start-up). The first pump system  101  can include a first pump  107  (e.g., a take-off pump configured to supply suitable flow during take-off or other high fuel flow conditions) and/or any other suitable pumps. 
     The aircraft fuel system  100  also includes a second pump system  111  that has an electric motor  113 . The second pump system  111  is in fluid communication with the fuel tank  105  and the fuel nozzle  109  to pump fuel from the fuel tank  105  to the fuel nozzle  109 . The second pump system  111  is driven by the electric motor  113  and is configured to pump flow in both the high flow rate condition (e.g., take-off, climb) and the low flow rate condition (e.g., cruise, descent, start-up). 
     In certain embodiments, the second pump system  111  can include a total-flow pump  115  and a main pump  117  attached to the electric motor  113 . The total-flow pump  115  can be configured to boost the main pump  117  and/or the mechanically driven first pump system  101 , for example. 
     The system  100  can include a first valve  118  configured to shut-off or otherwise limit fuel flow to the first pump system  101 . The first valve  118  can be positioned between the total-flow pump  115  and the first pump system  101 , for example, or in any other suitable location. 
     The system  100  can include a heat exchanger  119  disposed between the first pump system  101  and the second pump system  111 . The heat exchanger  119  can be a fuel-oil heat exchanger, for example, that is configured to cool engine oil with the fuel in the fuel system  100 . Any other suitable heat exchanger type is contemplated herein. The heat exchanger  119  can be placed downstream of the second pump system  111 . This ensures there is always fuel flow through the heat exchanger regardless of the selected flight phase and pump(s) operation mode. The heat exchanger  119  and/or any suitable additional heat exchanger(s) can be placed in any other suitable location(s). 
     In certain embodiments, the system  100  can include an ejector pump  121  disposed between the first pump system  101  and the total flow pump  115  of second pump system  111 . The ejector pump  121  can be configured to evacuate the first pump system  101  in the low fuel flow condition so as to reduce power consumption. The ejector pump  121  can include a venturi, for example. 
     In certain embodiments, the system  100  can include a system shut-off valve  123  disposed upstream of the fuel nozzle  109  and downstream of the first and second pump systems  101 ,  111 . System shut-off valve  123  can shut down the engine by preventing fuel flow to the fuel nozzle  109 . The valve  123  can prevent any accidental fuel dripping into the fuel nozzles  109 . 
     The system  100  can include a throttle valve  125  disposed between the first pump system  101  and the fuel nozzle  109  in certain embodiments. The throttle valve  125  can include feedback systems for control. Any other suitable valves (e.g., check valves  127 ) can be included as is appreciated by those having ordinary skill in the art in view of this disclosure. The throttle valve  125  can be placed upstream of the check valve  127 . This provides ability to limit the first pump system  101  with the throttle valve  125  and to close it or otherwise limit it with the first valve  118 . 
     The system  100  can include a controller  129  (e.g., an EEC, FADEC, any other distributed control architecture for example) operatively connected to the electric motor  113  and to one or more sensors disposed in the fuel system to control the electric motor as a function of output of the one or more sensors. The controller  129  can be operatively connected to one or more suitable valves (e.g., as described above) to actuate the valves as a function of the output of the one or more sensors. In certain embodiments, the one or more sensors can include at least one of a flow meter  131  disposed upstream of the fuel nozzle  109 , a pressure sensor  133  disposed downstream of the first pump system  101 , or a second pressure sensor  135  disposed downstream of the fuel nozzles  109 . 
     In certain embodiments, a pressure sensor  133  can be placed directly upstream of throttle valve  125 . The second pressure sensor  135  which can sense burner pressure of the combustor (not shown) can be placed downstream of the fuel nozzles  109 . The first and second pressure sensors  133 ,  135 , as well as the throttle valve  125 , can be operatively connected to and/or controlled by the controller  129  to provide accurate feedback for active control of the fuel flow meter  131  in real time. This can ensure optimum engine TSFC during all flight phases of the aircraft. 
     The main pump  117  can include one of a vane pump or gear pump, or any other suitable pump. The total-flow pump  115  and/or the take-off pump  107  can include a centrifugal pump, in certain embodiments. 
     Referring to  FIG. 2 , the system  100  is shown in a low fuel flow condition (e.g., cruise, start-up, descent). Notionally, the direction of fuel flow is shown with white arrows. As shown, the first valve  118  is closed, preventing fuel from traveling to the first pump system  101 . The first pump system  101  is evacuated of fuel by the ejector valve  121 , thereby reducing wear on the pump  107  by reducing power and heat load. This pump  107  may continue to rotate (e.g., if it is connected rigidly to the gearbox  103  shaft) however it is not pumping any liquid fuel and the pump load is minimal. Fuel is still allowed to flow from the second pump system  111  (e.g., at a rate controlled by the speed of the electric motor  113 , for example, as a function of sensor readings). 
     Referring to  FIG. 3 , the system  100  is shown in a high fuel flow condition (e.g., take-off). Notionally, the direction of fuel flow is shown with white arrows. As shown, the first valve  118  is open and allowing fuel from the tank  105  to the mechanically driven pump system  101  (e.g., which is boosted by the total-flow pump  115 ). Fuel is also pumping from the electric motor-driven second pump system  111 . In this regard, maximum fuel is being supplied to the engine for high power scenarios (e.g., take-off). The electric vane/gear main pump  117  is sized to provide a maximum fuel flow (e.g., 3000 pph, 150 psia for example) at 100% pump rotational speed. The combined output of all the fuel pumps ensures sufficient fuel flow and fuel pressure is provided to the fuel nozzles  109  of the engine during the take-off phase of the flight. This configuration may be re-activated during transient high-fuel-flow settings (e.g. step climb, acceleration, etc.) as needed. 
     Referring to  FIG. 4 , another embodiment of a fuel system  200  is shown. The second pump system  211  can additionally include a cruise pump  241  that is configured to boost the main pump  117  and is sized for fuel demands in a cruise flight condition. The total-flow pump  242  can be sized for a high fuel flow condition in such an embodiment, for example. The cruise pump  241  can include a centrifugal pump, in certain embodiments. 
     In certain embodiments, the system  200  can include a first heat exchanger  243  disposed between the total-flow pump  242  and the first pump system  101  and a second heat exchanger  245  disposed between the cruise pump  241  and the main pump  117 . The heat exchangers  243 ,  245  can be any suitable heat exchanger as described above, for example. The system  200  can include a system shut-off valve  247  disposed upstream of the first and second pump systems  101 ,  211 . 
     The system  200  can operate similarly to system  100  as described above. The system  200  includes additional pumping hardware and modified flow circuitry to provide additional pump pressure in the event of failure of the mechanically driven pump system  101  and/or allow additional fuel flow as needed. As shown, the total-flow pump  242  can evacuate the fuel flow during cruise flight conditions, e.g., in a more efficient way. When it is needed, the total-flow pump  242  is filled with fuel and provides fuel flow to the gearbox-driven take-off pump  107 . 
     In certain embodiments, total-flow pump  242  can also be mounted on the output shaft form the electric motor  113 . In certain embodiments, this pump  242  can be alternately mounted on a mechanical drive elsewhere and be driven by a mechanical pad rather than the motor. The gearbox driven take-off pump  107  can be multistage to improve specific speed and overall efficiency. 
     As described above, output of the gearbox-driven fuel pump  107  can flow through a throttle valve  125 , a check valve  127 , and can be eventually delivered to the fuel nozzles  109  of the engine. A fuel flow meter  131  can be placed downstream of check valve  127  and upstream of the fuel nozzles  109 . This can be used to calibrate the fuel flow vs. speed for the main pump  117  as well as the throttle valve  125  position vs. fuel flow speed for the gearbox-driven centrifugal take-off pump  107 . 
     A fuel pressure sensor  133  can be placed directly upstream of throttle valve  125 . A second pressure sensor  135  can be used to sense burner pressure of the combustor by being placed downstream of the fuel nozzles  109 . These sensors and the throttle valve  125  can be controlled by the controller  129  (e.g., an EEC/FADEC) to provide accurate feedback for active control of the fuel flow meter  131  in real time. This ensures optimum engine TSFC during all flight phases. 
     Embodiments allow metered fuel flow to be delivered to the fuel nozzles  109  based on the exact fuel demand set by the engine power settings (e.g., detected by the burner pressure sensor  135 ) as function of electric motor  113  speed (e.g., continuously variable) of the main fuel pump  117 . Embodiments can closely match the engine power settings and fuel demand with the fuel supply form the various fuel pumps. This optimizes the operation of the fuel pumps, thus extending their operational life (lower wear, heating, etc.). As a consequence, the overall fuel thermal management is improved (less/minimal fuel recirculation), while fuel remains a viable cooling sink due to its lower service temperature. This in turn, requires smaller/lighter/more compact heat exchangers which also saves weight. 
     Embodiments as described above eliminate wasteful fuel re-circulation during lower engine power settings, shuts off fuel flow to gearbox-driven pumps when not needed, lower power demand to drive fuel pumps, lower operational speed of fuel pumps, and allows all pumps to be controlled in real-time by a controller (e.g., the engine&#39;s EEC/FADEC for example). 
     The methods and systems of the present disclosure, as described above and shown in the drawings, provide for aircraft fuel systems with superior properties. While the apparatus and methods of the subject disclosure have been shown and described with reference to embodiments, those skilled in the art will readily appreciate that changes and/or modifications may be made thereto without departing from the spirit and scope of the subject disclosure.