Patent Publication Number: US-2020300168-A1

Title: Backup system for demand fuel pumping system

Description:
CROSS-REFERENCE TO RELATED APPLICATIONS 
     This application claims priority to U.S. Provisional Application No. 62/821,025 which was filed on Mar. 20, 2019. 
    
    
     BACKGROUND 
     A gas turbine engine typically includes a fan section, a compressor section, a combustor section and a turbine section. Air entering the compressor section is compressed and delivered into the combustion section where it is mixed with fuel and ignited to generate a high- energy exhaust gas flow. The high-energy exhaust gas flow expands through the turbine section to drive the compressor and the fan section. The compressor section typically includes low and high pressure compressors, and the turbine section includes low and high pressure turbines. 
     Fuel supplied to the combustor is provided by a mechanical pump driven by a rotating shaft of the engine. The mechanical pump is reliable and supplies fuel in proportion to engine speed. The minimum capacity of the mechanical pump is sized such that sufficient fuel is provided for high power conditions and/or engine starting. Excess fuel not needed is recirculated within the fuel system or back to the fuel tank. The fuel is further utilized as a coolant for other systems of the engine. Recirculation of fuel increases the temperature of the fuel and thereby reduces the available capacity to absorb heat from other systems. The capacity of the fuel to absorb heat from other systems is further limited by the characteristics of the fuel. At a certain temperature, the fuel begins to degrade and create deposits in the fuel system that can degrade engine performance. Reducing the amount of fuel that is recirculated during engine operation may improve the capacity of the fuel to absorb heat from other systems. 
     Turbine engine manufacturers continuously seek improvements to engine performance including improvements to thermal, transfer and propulsive efficiencies. 
     SUMMARY 
     A fuel system for a gas turbine engine according to an exemplary embodiment of this disclosure includes, among other possible things, a primary fuel pump providing fuel flow during engine operation and a primary electric motor coupled to the primary fuel pump for driving the primary fuel pump during engine operation. A secondary system provides fuel flow in the absence of fuel flow from the primary fuel pump. 
     In a further embodiment of a fuel system for a gas turbine engine, the secondary system comprises a secondary drive coupled to the primary fuel pump for driving the fuel pump instead of the electric motor. 
     In another embodiment of any of the foregoing fuel systems for a gas turbine engine, the secondary drive comprises a hydraulic turbine coupled to the primary pump. 
     In another embodiment of any of the foregoing fuel systems for a gas turbine engine, a clutch means for selectively coupling the hydraulic turbine is included to drive the primary pump. 
     In another embodiment of any of the foregoing fuel systems for a gas turbine engine, a hydraulic control valve controls a flow of hydraulic fluid to the hydraulic turbine to control a speed of the hydraulic turbine and thereby the speed of the primary pump and the flow of fuel. 
     In another embodiment of any of the foregoing fuel systems for a gas turbine engine, secondary system comprises a secondary pump powered by a secondary drive. 
     In another embodiment of any of the foregoing fuel systems for a gas turbine engine, the secondary drive comprises a hydraulically powered turbine. 
     In another embodiment of any of the foregoing fuel systems for a gas turbine engine, the secondary drive comprises an air cycle machine driven by a bleed airflow from the engine. 
     In another embodiment of any of the foregoing fuel systems for a gas turbine engine, a bleed air control valve controls a speed of the air cycle machine. 
     In another embodiment of any of the foregoing fuel systems for a gas turbine engine, a first valve is included upstream of the secondary pump and a second valve is included downstream of the secondary pump for controlling fuel flow from the secondary pump. 
     A gas turbine engine according to an exemplary embodiment of this disclosure includes, among other possible things, a fan rotatable within a fan nacelle and a core engine that includes a compressor communicating compressed air to a combustor where compressed air is mixed with fuel and ignited to generate a high-energy gas flow expanded through a turbine. A primary fuel pump provides fuel to the combustor during engine operation. A primary electric motor is coupled to the primary fuel pump for driving the primary fuel pump during engine operation, and a secondary system providing fuel flow in the absence of fuel flow from the primary fuel pump. 
     In a further embodiment of the foregoing gas turbine engine, the secondary system comprises a hydraulically powered turbine coupled to drive the primary fuel pump instead of the primary electric motor. 
     In another embodiment of any of the foregoing gas turbines engines, a hydraulic control valve controls a flow of hydraulic fluid to the hydraulic turbine to control a speed of the hydraulic turbine, and thereby the speed of the primary pump and the flow of fuel to the combustor. 
     In another embodiment of any of the foregoing gas turbines engines, secondary system comprises a secondary pump powered by a secondary drive. 
     In another embodiment of any of the foregoing gas turbines engines, the secondary drive comprises a hydraulic turbine driven by a flow of fluid. 
     In another embodiment of any of the foregoing gas turbines engines, the hydraulic turbine comprise an air cycle machine driven by a flow of air bled from the compressor. A bleed air control valve controls a speed of the air cycle machine by varying the flow of air bleed from the compressor. 
     In another embodiment of any of the foregoing gas turbines engines, a first valve is upstream of the secondary pump and a second valve is downstream of the secondary pump for controlling fuel flow from the secondary pump. 
     A method of supplying fuel to a combustor of a gas turbine engine according to an exemplary embodiment of this disclosure includes, among other possible things, a primary fuel pump driven by an electric motor to provide a first fuel flow that varies independent of a speed of shaft driven by a turbine of the engine. A second fuel flow is generated with a secondary system in response to the electric motor not driving the primary fuel pump sufficiently to power the engine. 
     In another embodiment of any of the foregoing gas turbines engines, the second fuel flow is generated with a secondary drive driving the primary fuel pump. 
     In another embodiment of any of the foregoing gas turbines engines, the second fuel flow is generated with a secondary fuel pump driven by a secondary drive independent of the primary fuel pump. 
     In another embodiment of any of the foregoing gas turbines engines, the secondary drive comprises an air cycle machine powered by a flow of air bleed from a compressor of the engine. 
     Although the different examples have the specific components shown in the illustrations, embodiments of this invention are not limited to those particular combinations. It is possible to use some of the components or features from one of the examples in combination with features or components from another one of the examples. 
     These and other features disclosed herein can be best understood from the following specification and drawings, the following of which is a brief description. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         FIG. 1  is a schematic view of an example gas turbine engine. 
         FIG. 2  is schematic view of an example fuel system embodiment. 
         FIG. 3  is a schematic view of another example fuel system embodiment. 
     
    
    
     DETAILED DESCRIPTION 
       FIG. 1  schematically illustrates a gas turbine engine  20 . The gas turbine engine  20  is disclosed herein as a two-spool turbofan that generally incorporates a fan section  22 , a compressor section  24 , a combustor section  26  and a turbine section  28 . The fan section  22  drives air along a bypass flow path B in a bypass duct defined within a nacelle  18 , and also drives air along a core flow path C for compression and communication into the combustor section  26  then expansion through the turbine section  28 . Although depicted as a two-spool turbofan gas turbine engine in the disclosed non-limiting embodiment, it should be understood that the concepts described herein are not limited to use with two-spool turbofans as the teachings may be applied to other types of turbine engines including three-spool architectures and low bypass engines. 
     The exemplary engine  20  generally includes a low speed spool  30  and a high speed spool  32  mounted for rotation about an engine central longitudinal axis A relative to an engine static structure  36  via several bearing systems  38 . It should be understood that the various bearing systems  38  may alternatively or additionally be provided at different locations, and the location of bearing systems  38  may be varied as appropriate to the application. 
     The low speed spool  30  generally includes an inner shaft  40  that interconnects, a first (or low) pressure compressor  44  and a first (or low) pressure turbine  46 . The inner shaft  40  is connected to a fan section  22  through a speed change mechanism, which in exemplary gas turbine engine  20  is illustrated as a geared architecture  48  to drive fan blades  42  at a lower speed than the low speed spool  30 . The high speed spool  32  includes an outer shaft  50  that interconnects a second (or high) pressure compressor  52  and a second (or high) pressure turbine  54 . A combustor  56  is arranged in exemplary gas turbine  20  between the high pressure compressor  52  and the high pressure turbine  54 . A mid-turbine frame  58  of the engine static structure  36  may be arranged generally between the high pressure turbine  54  and the low pressure turbine  46 . The mid-turbine frame  58  further supports bearing systems  38  in the turbine section  28 . The inner shaft  40  and the outer shaft  50  are concentric and rotate via bearing systems  38  about the engine central longitudinal axis A which is collinear with their longitudinal axes. 
     The core airflow is compressed by the low pressure compressor  44  then the high pressure compressor  52 , mixed and burned with fuel in the combustor  56 , then expanded over the high pressure turbine  54  and low pressure turbine  46 . The mid-turbine frame  58  includes airfoils  60  which are in the core airflow path C. The turbines  46 ,  54  rotationally drive the respective low speed spool  30  and high speed spool  32  in response to the expansion. It will be appreciated that each of the positions of the fan section  22 , compressor section  24 , combustor section  26 , turbine section  28 , and fan drive gear system  48  may be varied. For example, gear system  48  may be located aft of the low pressure compressor  44  and the fan blades  42  may be positioned forward or aft of the location of the geared architecture  48  or even aft of turbine section  28 . 
     The engine  20  in one example is a high-bypass geared aircraft engine. In a further example, the engine  20  bypass ratio is greater than about six (6), with an example embodiment being greater than about ten ( 10 ), the geared architecture  48  is an epicyclic gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3 and the low pressure turbine  46  has a pressure ratio that is greater than about five. In one disclosed embodiment, the engine  20  bypass ratio is greater than about ten (10:1), the fan diameter is significantly larger than that of the low pressure compressor  44 , and the low pressure turbine  46  has a pressure ratio that is greater than about five 5:1. Low pressure turbine  46  pressure ratio is pressure measured prior to inlet of low pressure turbine  46  as related to the pressure at the outlet of the low pressure turbine  46  prior to an exhaust nozzle. The geared architecture  48  may be an epicycle gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3:1 and less than about 5:1. It should be understood, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present invention is applicable to other gas turbine engines including direct drive turbofans. 
     A significant amount of thrust is provided by the bypass flow B due to the high bypass ratio. The fan section  22  of the engine  20  is designed for a particular flight condition—typically cruise at about 0.8 Mach and about 35,000 feet (10,668 meters). The flight condition of 0.8 Mach and 35,000 ft (10,668 meters), with the engine at its best fuel consumption—also known as “bucket cruise Thrust Specific Fuel Consumption (‘TSFCT’)”—is the industry standard parameter of 1 bm of fuel being burned divided by 1 bf of thrust the engine produces at that minimum point. “Low fan pressure ratio” is the pressure ratio across the fan blade alone, without a Fan Exit Guide Vane (“FEGV”) system. The low fan pressure ratio as disclosed herein according to one non-limiting embodiment is less than about 1.45. “Low corrected fan tip speed” is the actual fan tip speed in ft/sec divided by an industry standard temperature correction of [(Tram ° R)/(518.7 ° R)] 0.5 . The “Low corrected fan tip speed” as disclosed herein according to one non-limiting embodiment is less than about 1150 ft/second (350.5 meters/second). 
     The example gas turbine engine includes the fan section  22  that comprises in one non-limiting embodiment less than about  26  fan blades  42 . In another non-limiting embodiment, the fan section  22  includes less than about  20  fan blades  42 . Moreover, in one disclosed embodiment the low pressure turbine  46  includes no more than about 6 turbine rotors schematically indicated at  34 . In another non-limiting example embodiment, the low pressure turbine  46  includes about 3 turbine rotors. A ratio between the number of fan blades  42  and the number of low pressure turbine rotors is between about 3.3 and about 8.6. The example low pressure turbine  46  provides the driving power to rotate the fan section  22  and therefore the relationship between the number of turbine rotors  34  in the low pressure turbine  46  and the number of blades  42  in the fan section  22  disclose an example gas turbine engine  20  with increased power transfer efficiency. 
     Fuel is delivered to the combustor  56  by a fuel system  62 . The example fuel system  62  includes a primary system  66  and a back-up or secondary system  68 . Fuel from a fuel tank  64  is pumped to a desired pressure and provided to the combustor  56 . The disclosed fuel system  62  tailors a flow of fuel  70  to the combustor  56  based on engine operating conditions. Instead of simply providing a fuel flow that provides for extremes of operating demands, the disclosed fuel system  62  varies the flow of fuel  70  according to a demand for fuel. By tailoring the flow of fuel to engine operating demand, a fuel recirculation loop for excess fuel can be reduced and/or eliminated. 
     Fuel is utilized as a heat sink to cool other flows within the engine such as lubricant and air flows. In this example, a heat fuel/oil heat exchanger  65  cools a flow of lubricant generated by a lubricant system  55 . Recirculation of fuel results in an increased temperature of the fuel and thereby a reduced capability to accept heat from other engine systems, such as the example lubricant system  55 . 
     The disclosed fuel system  62  varies the flow of fuel  70  based on demand to reduce and/or eliminate the recirculation of fuel and thereby increase the ability to accept heat from other engine systems. 
     Referring to  FIG. 2  with continued reference to  FIG. 1 , the example fuel system  62  includes a primary fuel pump  72  that generates the flow of fuel  70  during engine operation. The primary fuel pump  72  is driven by an electric motor  74  coupled to the primary fuel pump  72  by a shaft  76 . The electric motor  74  is controlled by a controller  88  that varies a speed of the motor  74  to drive the primary pump  72  in a variable manner to match engine fuel demand. It is desirable to provide a back-up system to provide an uninterrupted flow of fuel  70  in the event operation of the electric motor is disrupted. The example fuel system  62  includes the secondary system  68  to continue the flow of fuel  70  in the event the electric motor is unable to drive the primary fuel pump  72 . 
     The secondary system  68  includes a hydraulically powered turbine  78  coupled to the shaft  76  to drive the primary fuel pump  72  independent of the electric motor  74 . In the disclosed example, a clutch  108  is provided to selectively couple the turbine  78  to the shaft  76 . The clutch  108  may be decoupled while the electric motor  74  operates the primary fuel pump  72 . In instances where the electric motor  74  is not driving the primary fuel pump  72 , the clutch  108  may be actuated to couple the turbine  78  to the shaft  76  and drive the primary pump fuel  72 . 
     In the disclosed example, the turbine  78  is driven by hydraulic fluid flow produced by a hydraulic pump  84  receiving hydraulic fluid from a hydraulic system schematically indicated at  86 . The hydraulic system  86  may be a dedicated system for providing a back-up drive for the fuel system  62 . The hydraulic system  86  may also be part of other engine systems that provide hydraulic power to hydraulic actuators or other hydraulic devices existing on the engine or aircraft. The pump  82  maybe operated through a mechanical connection schematically shown at  84  to a rotating shaft driven by the turbine section  28 . The connection  84  may also be to an accessory gearbox or other mechanically coupled device provided on the engine  20 . 
     A hydraulic control valve  80  controls operation of the turbine  78 . The control valve  80  may normally be in a closed condition such that the turbine  78  is not driven. Opening of the control valve  80  generates hydraulic flow that powers the turbine  78 . Operation of the control valve  80  is governed by the controller  88  to vary a flow of hydraulic fluid to the hydraulic turbine  78  to thereby control a speed of the turbine  78  and thereby the speed of the primary pump  72  and the flow of fuel  70 . The hydraulic turbine  78  may be part of lubricant system or a hydraulic system already existing and part of the engine and aircraft. 
     Referring to  FIG. 3 , with continued reference to  FIG. 1 , another example fuel system embodiment is schematically shown and indicated at  62 ′. The example fuel system  62 ′ includes a secondary system  68 ′ that includes a secondary pump  92  powered by a secondary drive  98 . In this disclosed example, the secondary drive  98  is an air cycle machine driven by bleed airflow  108  drawn from a tap  110  from the compressor section  24  of the engine  20 . 
     The air cycle machine  98  is driven by the bleed airflow  108  obtained from the compressor location  110 . The air cycle machine uses the bleed airflow  108  for other aircraft systems including environmental control system and cooling systems for hot sections of the engine  20 . The air cycle machine  98  is coupled to a secondary pump  92  by a shaft  100 , shown schematically. The secondary pump  92  is disposed in a secondary passage  96  including a first control valve  104  upstream of the secondary pump  92  and a second control valve  106  downstream of the secondary pump  106 . Fuel flow within the secondary passage  96  is closed during routine operation of the primary fuel pump  90 . A bleed air control valve  12  controls operation and speed of the air cycle machine and thereby of the shaft  100  diving the secondary pump  92 . 
     Because the air cycle machine  98  is operable during engine operation to provide bleed airflow to other engine systems, the mechanical connection through the shaft  100  continually drives the secondary pump  92 . However, because the control valves  104  and  106  are normally in a closed position, the secondary pump  92  is not in communication with fuel and does not contribute to fuel flow to the combustor  56 . An evacuation system pump  114  is in selective communication with the secondary passage  96  to exhaust fuel from the secondary pump  92  when not in use. In one disclosed example, a control valve  112  controls communication between the evacuation system  114  and the secondary passage  96 . Once the control valves  104  and  106  are closed, the control valve  112  opens to enable the evacuation system  114  to evacuate fuel back to the fuel tank  64 . The evacuation system  114  could then be turned off once fuel is evacuated from the secondary pump  92  and the secondary passage  96 . 
     In the event that the electric motor  74  is no longer able to drive the primary fuel pump  90  and thereby provide fuel flow through the first fuel passage  94 , the first and second control valves  104  and  106  are opened to allow the secondary pump  92  to supply a flow of fuel to the combustor  56 . The flow of fuel provided by the secondary fuel pump  92  can be varied by adjusting a flow of bleed airflow with the bleed air control valve  102 . 
     Operation of the engine  20  with reference to the Figures is therefore provided according to a disclosed method of supplying fuel to the combustor  56  by driving the primary fuel pump  72 ,  90  with the electric motor  74  to provide the fuel flow  70  that varies independent of a speed of the shafts  40 ,  50  driven by the turbine section  28 . Because the pump  72 ,  90  is not mechanically coupled to the shafts  40 ,  50 , the pump  72 ,  90  can be operated in a variable manner based on a demand for fuel. Accordingly, the flow of fuel can be reduced for applicable engine operating conditions to reduce and/or eliminate the need for recirculation of excess fuel flow. The decoupling of the engine speed from fuel flow enables an improved match between pump operation and actual demand of fuel flow. The reduced or eliminated recirculation of excess of fuel reduces the overall temperature of the fuel flow and thereby increases the heat acceptance capacity of the fuel flow. 
     In the event that the electric motor  74  becomes incapable of driving the primary pump  72 ,  90 , the disclosed secondary systems  68 ,  68 ′ provide for the generation of a second fuel flow by either driving the primary fuel pump  72 , or driving a secondary fuel pump  90 . A secondary drive in the form of a hydraulically powered turbine  78  or an air cycle machine  98  provide the power to continue the flow of fuel  70  to the combustor  56 . 
     Accordingly, the disclosed fuel systems provide a varied flow to match engine demand during operation that enables an increased heat acceptance capacity of the fuel while maintaining operation with a secondary drive system to assure uninterrupted fuel flow. 
     Although an example embodiment has been disclosed, a worker of ordinary skill in this art would recognize that certain modifications would come within the scope of this disclosure. For that reason, the following claims should be studied to determine the scope and content of this disclosure.