Multi-temperature fuel injector for a gas turbine engine

Abstract

A gas turbine engine includes a combustor, a plurality of fuel nozzles, and at least one fuel manifold. The combustor includes a combustion chamber. The plurality of fuel nozzles inject fuel into the combustion chamber of the combustor. The gas turbine engine may include a first fuel circuit and a second fuel circuit. The first fuel circuit includes a first fuel manifold fluidly connected to at least one of the plurality of fuel nozzles to distribute fuel to the at least one fuel nozzle at a first temperature. The second fuel circuit includes a second fuel manifold fluidly connected to at least one of the plurality of fuel nozzles to distribute fuel to the at least one fuel nozzle at a second temperature lower than the first temperature.

Description

Technical Field

[0001] This disclosure relates to a fuel system for a gas turbine engine, and more specifically, to a gas turbine engine for an aircraft. Background Technology

[0002] In a gas turbine engine, fuel is injected into the combustion chamber using fuel nozzles, where it mixes with air and burns to produce combustion products that drive the turbine of the gas turbine engine. For example, in an aircraft, this combustion generates thrust to propel the aircraft. The fuel can be heated upstream of the fuel nozzles and before being injected into the combustion chamber to improve the efficiency of the gas turbine engine (i.e., reduce the amount of fuel burned), especially during operational conditions such as takeoff, climb, and cruise. Attached Figure Description

[0003] The features and advantages of this disclosure will become apparent from the following description of various exemplary embodiments, as illustrated in the accompanying drawings, wherein similar reference numerals generally denote the same, functionally similar, and / or structurally similar elements.

[0004] Figure 1 This is a schematic perspective view of an aircraft with a gas turbine engine according to an embodiment of the present disclosure.

[0005] Figure 2 yes Figure 1 The gas turbine engine of the aircraft shown Figure 1 A schematic cross-sectional view taken from line 2-2 in the figure.

[0006] Figure 3 This is a schematic diagram of a fuel system for a gas turbine engine according to an embodiment of the present disclosure.

[0007] Figure 4 This is a schematic diagram of a fuel system for a gas turbine engine according to an embodiment of the present disclosure.

[0008] Figure 5 It shows that it can be used Figure 3 and Figure 4 The arrangement of heat exchangers in the fuel system is shown.

[0009] Figure 6 It shows that it can be used Figure 3 and Figure 4 Another arrangement of the heat exchanger in the fuel system shown.

[0010] Figure 7 It shows that it can be used Figure 3 and Figure 4 Another arrangement of the heat exchanger in the fuel system shown. Detailed Implementation

[0011] The features, advantages, and embodiments of this disclosure will be apparent or obvious upon consideration of the following detailed description, accompanying drawings, and claims. Furthermore, it should be understood that the following detailed description is exemplary and intended to provide further explanation without limiting the scope of the claimed disclosure.

[0012] Various embodiments are discussed in detail below. Although specific embodiments are discussed, they are for illustrative purposes only. Those skilled in the art will recognize that other components and configurations can be used without departing from the spirit and scope of this disclosure.

[0013] As used herein, the terms “first,” “second,” and “third” are used interchangeably to distinguish one component from another and are not intended to indicate the location or importance of the individual components.

[0014] The terms "front" and "rear" refer to relative positions within a gas turbine engine or vehicle and to the normal operating posture of the gas turbine engine or vehicle. For example, for a gas turbine engine, "front" refers to the position closer to the engine inlet, while "rear" refers to the position closer to the engine nozzle or exhaust port.

[0015] The terms "upstream" and "downstream" refer to the relative directions of fluid flow within a fluid path. For example, "upstream" refers to the direction from which the fluid flows, and "downstream" refers to the direction from which the fluid flows.

[0016] The terms "direct upstream" or "direct downstream" when used to describe the relative positions of components in a fluid path refer to components placed adjacent to each other in the fluid path without any intermediate components between them, except for appropriate fluid connections such as pipes, tubes, valves, etc., to fluidly connect the components. These components may be spaced apart from each other by intermediate components not in the fluid path.

[0017] Unless otherwise specified herein, the terms “connection,” “fixation,” “attachment,” “linkage,” etc., refer to direct connection, fixation, attachment, or linking, as well as indirect connection, fixation, attachment, or linking through one or more intermediate components or features.

[0018] Unless the context clearly indicates otherwise, the singular forms “a,” “one,” and “the” include plural references.

[0019] As used throughout this specification and claims, approximate language is applied to modify any quantitative expression that allows for variation without altering its underlying function. Therefore, values ​​modified by one or more terms such as “approximately,” “about,” and “substantially” are not limited to specified exact values. In at least some cases, approximate language may correspond to the precision of the instrument used to measure the value, or the precision of the method or machine used to construct or manufacture the component and / or system. For example, approximate language may refer to a margin of one percent, two percent, four percent, ten percent, fifteen percent, or twenty percent at the endpoints of a single value, a range of values, and / or a defined range of values.

[0020] Scope limitations are combined and interchanged herein and throughout the specification and claims. Unless the context or language otherwise indicates, such scopes are identified and include all subscopes contained herein. For example, all scopes disclosed herein include endpoints, and endpoints may be combined independently of each other.

[0021] As mentioned above, the fuel efficiency of a gas turbine engine can be improved by heating the fuel upstream of the fuel injector and before it is injected into the combustion chamber, especially during operating conditions such as takeoff, climb, and cruise. However, injecting fuel into the combustion chamber at high temperatures can lead to other problems. One such problem is the risk of flameout during rapid throttle closure, a phenomenon known as chop. An example is when a pilot chops the throttle from cruise to idle for a descent. During cruise, the fuel, fuel injectors, and other components are at elevated temperatures. When the mass flow rate of fuel through the injector decreases rapidly during chop to idle, the hot fuel may boil and evaporate, causing fuel flow oscillations and potentially combustor flameout. To avoid this, fuel can be supplied to the injectors from the secondary fuel circuit at a reduced temperature.

[0022] The fuel systems discussed in this article are particularly applicable to engines, such as gas turbine engines used in aircraft. Figure 1 This is a perspective view of an aircraft 10 that can be implemented in various preferred embodiments. The aircraft 10 includes a fuselage 12, wings 14 attached to the fuselage 12, and a tail 16. The aircraft 10 also includes a propulsion system that generates the thrust required to propel the aircraft 10 during flight, taxiing, and other operations. Figure 1 The propulsion system of the aircraft 10 shown includes a pair of engines 100. In this embodiment, each engine 100 is attached to one of the wings 14 in an underwing configuration via a pylon 18. Although the engines 100 are... Figure 1The engine 100 is shown attached to the wing 14 in an underwing configuration, but in other embodiments, the engine 100 may have an alternative configuration and may be coupled to other parts of the aircraft 10. For example, the engine 100 may additionally or alternatively include one or more aspects coupled to other parts of the aircraft 10 (e.g., tail 16 and fuselage 12).

[0023] As will be referenced below Figure 2 Further description, Figure 1 The engine 100 shown is a gas turbine engine capable of selectively generating propulsive thrust for the aircraft 10. The amount of propulsive thrust can be at least partially based on the thrust generated via fuel system 200 (see fuel system 200). Figure 3 The volume of fuel supplied to the gas turbine engine 100 is controlled. The aviation turbine fuel in the embodiments discussed herein is a combustible hydrocarbon liquid fuel with the desired carbon number, such as kerosene-based fuel. The fuel is stored in the fuel tank 210 of the fuel system 200. Figure 1 As shown, at least a portion of the fuel tank 210 is located within each wing 14, and a portion of the fuel tank 210 is located within the fuselage 12 between the wings 14. However, the fuel tank 210 may be located at other suitable locations within the fuselage 12 or the wings 14. The fuel tank 210 may also be entirely located within the fuselage 12 or the wings 14. The fuel tank 210 may also be a separate tank, rather than a single integral body; for example, two tanks, each located within a corresponding wing 14.

[0024] although Figure 1 The aircraft 10 shown is an airplane, but the embodiments described herein are also applicable to other aircraft, including, for example, helicopters and unmanned aerial vehicles (UAVs). Preferably, the aircraft discussed herein is a fixed-wing or rotary-wing aircraft that generates lift through aerodynamic forces acting on, for example, a fixed wing (e.g., wing 14) or a rotating wing (e.g., the rotor of a helicopter) and is heavier than air rather than lighter than air (e.g., an airship). Furthermore, although not described herein, in other embodiments, the gas turbine engine can be any other suitable type of gas turbine engine, such as an industrial gas turbine engine incorporated into a power generation system, a marine gas turbine engine, etc.

[0025] Figure 2 yes Figure 1 A schematic cross-sectional view of one of the engines 100 used in the propulsion system of the aircraft 10 shown. Figure 2 The cross-sectional view is along Figure 1 It was cut from line 2-2 in the middle. For Figure 2In the depicted embodiment, engine 100 is a high-bypass turbofan engine. Engine 100 may also be referred to herein as turbofan engine 100. Turbofan engine 100 has an axial direction A (extending parallel to the longitudinal centerline 101, in...) Figure 2 (As shown in the diagram for reference), the radial direction R (extending perpendicularly to the longitudinal centerline 101, in...) Figure 2 (Shown for reference) and circumferential direction. Circumferential direction ( Figure 2 (Not depicted) extends in a direction that rotates about the axial direction A. The turbofan engine 100 includes a fan section 102 and a turbine 104 disposed downstream of the fan section 102.

[0026] Figure 2 The turbine 104 depicted includes a tubular housing 106 (also referred to as a casing or nacelle) defining an inlet 108. In this embodiment, the inlet 108 is annular. The housing 106 surrounds an engine core that includes, in series flow relationships, a compressor section comprising a turbocharger or low-pressure (LP) compressor 110 and a high-pressure (HP) compressor 112; a combustion section 150 (also referred to herein as a combustor 150); a turbine section comprising a high-pressure (HP) turbine 116 and a low-pressure (LP) turbine 118; and an exhaust nozzle section 120. The compressor section, combustion section 150, and turbine section together at least partially define a core airflow path 121 extending from the inlet 108 to the exhaust nozzle section 120. The turbofan engine 100 also includes one or more drive shafts. More specifically, the turbofan engine 100 includes a high-pressure (HP) shaft or spool 122 that drives an HP turbine 116 to an HP compressor 112 and a low-pressure (LP) shaft or spool 124 that drives an LP turbine 118 to an LP compressor 110.

[0027] Figure 2 The fan section 102 shown includes a fan 126 having multiple fan blades 128 coupled to a disk 130. The multiple fan blades 128 and the disk 130 can rotate together via an LP shaft 124 about a longitudinal centerline (axis) 101. The LP compressor 110 can also be directly driven by the LP shaft 124, as... Figure 2 As depicted, the disc 130 is covered by a rotatable front hub 132, which has an aerodynamic profile to facilitate airflow through multiple fan blades 128. Furthermore, an annular fan housing or outer nacelle 134 circumferentially surrounds at least a portion of the fan 126 and / or turbine 104. The nacelle 134 is supported relative to the turbine 104 by multiple circumferentially spaced outlet guide blades 136. A downstream section 138 of the nacelle 134 extends above the outer portion of the turbine 104 to define a bypass airflow passage 140 therebetween.

[0028] The turbofan engine 100 can operate in conjunction with and receive a fuel flow from the fuel system 200. As will be further described below, the fuel system 200 includes a fuel delivery assembly 202 that supplies a fuel flow from a fuel tank 210 to the turbofan engine 100, and more specifically, to a plurality of fuel nozzles 152 that inject fuel into a combustion chamber 154 of a burner 150. Each fuel nozzle 152 may be positioned at the front end of the combustion chamber 154, and each fuel nozzle 152 may be part of a swirler / fuel nozzle assembly. In some embodiments, the burner 150 is an annular burner 150, and the plurality of fuel nozzles 152 are arranged in an annular configuration, wherein the plurality of fuel nozzles 152 (swirler / fuel nozzle assemblies) are aligned in the circumferential direction of the burner. The swirler / fuel nozzle assembly may also be referred to as a cup, and the cups may be arranged adjacent to each other, as will be referenced below. Figure 3 and Figure 4 Further discussion is needed.

[0029] As described above, the compressor section, combustion section (burner) 150, and turbine section at least partially form a core airflow path 121 extending from inlet 108 to the exhaust nozzle section 120. Air entering through inlet 108 is compressed by the blades of multiple fans of the LP compressor 110 and HP compressor 112. At least a portion of the compressed air enters (as primary air) the front end of the combustion chamber 154 of the burner 150. Fuel is injected into the compressed air by fuel nozzle 152 and mixes with the compressed primary air. As described above, in this embodiment, fuel nozzle 152 is part of a swirler / fuel nozzle assembly. The swirler / fuel nozzle assembly includes a swirler (not shown) for generating turbulence in the primary air. Fuel nozzle 152 injects fuel into the turbulent flow of primary air, and the turbulence promotes rapid mixing of fuel and primary air. The mixture of fuel and compressed air is burned in the combustion chamber 154, producing combustion gases (combustion products), which are accelerated as they exit the combustion chamber 154. As the combustion products are discharged through the outlet of combustion chamber 154 to drive engine 100, the combustion products are accelerated. More specifically, the accelerated combustion products through the outlet rotate the turbines of HP turbine 116 and LP turbine 118 (e.g., to drive turbine blades). As discussed above, HP turbine 116 and LP turbine 118, along with other components, drive LP compressor 110 and HP compressor 112.

[0030] The turbofan engine 100 also includes various accessory systems to assist in the operation of the turbofan engine 100 and / or the aircraft (including the turbofan engine 100). For example, the turbofan engine 100 may include a main lubrication system 162, a compressor cooling air (CCA) system 164, an active thermal gap control (ATCC) system 166, and a generator lubrication system 168, each of which is... Figure 2 The diagram is schematically depicted. The main lubrication system 162 is configured to provide lubricant to various bearings and gear meshes, such as in the compressor section, turbine section, HP shaft 122, and LP shaft 124. The lubricant provided by the main lubrication system 162 can increase the service life of these components and can remove a certain amount of heat from these components by using one or more heat exchangers. The compressor cooling air (CCA) system 164 supplies air from one or both of the HP compressor 112 or LP compressor 110 to one or both of the HP turbine 116 or LP turbine 118. The active thermal gap control (ATCC) system 166 is used to minimize the gap between the turbine blade tip and the casing wall as the casing temperature changes during flight missions. The generator lubrication system 168 provides lubrication to the electric generator (not shown) and provides cooling / heat removal for the electric generator. The electric generator can provide power to, for example, the starter motor of the turbofan engine 100 and / or various other electronic components of the turbofan engine 100 and / or the aircraft including the turbofan engine 100. The lubrication system 100 for the engine 100 (e.g., the main lubrication system 162 and the generator lubrication system 168) can be lubricated using hydrocarbon fluids (e.g., oil), wherein the oil circulates through the inner surfaces of the oil-cleaning lines.

[0031] Of course, the turbofan engine 100 discussed herein is provided merely as an example. In other embodiments, any other suitable engine may be used in conjunction with aspects of this disclosure. For example, in other embodiments, the engine may be any other suitable gas turbine engine, such as a turboshaft engine, a turboprop engine, a turbojet engine, a ductless single-fan engine, etc. In this way, it should also be understood that in other embodiments, the gas turbine engine may have other suitable configurations, such as other suitable numbers or arrangements of shafts, compressors, turbines, fans, etc. Furthermore, although the turbofan engine 100 is shown as a directly driven, fixed-pitch turbofan engine 100, in other embodiments, the gas turbine engine may be a geared gas turbine engine (i.e., a gearbox including a fan 126 and a shaft (e.g., LP shaft 124) driving the fan 126), or a variable-pitch gas turbine engine (i.e., a fan 126 including a plurality of fan blades 128 rotatable about their respective pitch axes). Furthermore, still in alternative embodiments, aspects of this disclosure may be incorporated into or otherwise used with any other type of engine (e.g., a reciprocating engine). Additionally, in other exemplary embodiments, the exemplary turbofan engine 100 may include or be operatively connected to any other suitable accessory system. Additionally or alternatively, the exemplary turbofan engine 100 may not include or may be operatively connected to one or more of the accessory systems 162, 164, 166, 168 discussed above.

[0032] Figure 3 This is a schematic diagram of a fuel system 200 according to an embodiment of the present disclosure, configured to store hydrocarbon fuel for engine 100 in a fuel tank 210 and deliver the hydrocarbon fuel to engine 100 via a fuel delivery assembly 202. In the following discussion, various components are described as being fluidly connected to or in fluid communication with the fuel delivery assembly 202. These components are also fluidly connected or coupled to each other, for example, via the fuel delivery assembly 202.

[0033] Fuel delivery assembly 202 includes pipes, conduits, ducts, etc., to fluidly connect various components of fuel system 200 to engine 100. As described above, fuel tank 210 is configured to store hydrocarbon fuel, and hydrocarbon fuel is supplied from fuel tank 210 to fuel delivery assembly 202. Fuel delivery assembly 202 is configured to deliver hydrocarbon fuel between fuel tank 210 and engine 100, thereby providing a flow path (fluid path) for hydrocarbon fuel from fuel tank 210 to engine 100. As described above, the terms "downstream" and "upstream" as used herein can be used to describe the position of a component relative to the flow direction of hydrocarbon fuel in the flow path of fuel delivery assembly 202. A component located downstream of another component is configured to receive fuel from another component, and similarly, a component located upstream of another component is configured to supply fuel to another component. Fuel delivery assembly 202 may also include various valves and other components for delivering hydrocarbon fuel to engine 100, these components being in... Figure 3 Not shown in the image.

[0034] The fuel system 200 includes at least one fuel pump, and Figure 3 The illustrated embodiment includes multiple fuel pumps fluidly connected to the fuel delivery assembly 202 to direct fuel flow through the fuel delivery assembly 202 to the engine 100. One such pump is a main fuel pump 212. The main fuel pump 212 is a high-pressure pump and is the primary source of pressure increase in the fuel delivery assembly 202 between the fuel tank 210 and the engine 100. The main fuel pump 212 can be configured to increase the pressure in the fuel delivery assembly 202 to a level greater than that in the combustion chamber 154 of the burner 150 (see [link to documentation]). Figure 2 The pressure within.

[0035] The fuel system 200 may also include other auxiliary pumps, such as inlet pump 214. Inlet pump 214 is a low-pressure pump configured to provide initial pressurization to direct the hydrocarbon fuel flow through fuel delivery assembly 202. Inlet pump 214 may be configured to provide a smaller pressure rise within fuel delivery assembly 202 than the main fuel pump 212. Inlet pump 214 may be configured to provide a pressure rise less than 80 percent of the main fuel pump 212, for example, less than 70 percent, less than 60 percent, less than 50 percent, less than 40 percent, less than 30 percent, less than 20 percent, or at least 5 percent of the main fuel pump 212's pressure rise.

[0036] exist Figure 3 In the illustrated embodiment, inlet pump 214 is downstream of fuel tank 210 and upstream of main fuel pump 212. Although inlet pump 214 is shown as being located within engine 100, inlet pump 214 may also be suitably located in other parts of aircraft 10 (e.g., fuselage 12, wing 14, or pylon 18 (see [reference]). Figure 1In the process, inlet pump 214 directs the fuel flow from fuel tank 210, and then the fuel is heated by preheater 216.

[0037] The preheater 216 is in fluid communication with the fuel delivery assembly 202 and can be any suitable heater, such as a resistance heater, catalytic heater, or burner. In some embodiments, for example... Figure 3 In one embodiment depicted, the preheater 216 may be a heat exchanger in thermal communication with any suitable heat source (e.g., any suitable engine and / or aircraft heat source). Such an engine heat source may include, for example, a main lubrication system 162, and the preheater 216 may be a fuel oil heat exchanger (HX) fluidly connected to the main lubrication system 162 and configured to extract heat from the oil in the main lubrication system 162 and heat the hydrocarbon fuel flowing through the preheater 216. The preheater 216 is preferably configured to heat the fuel to a temperature that avoids ice formation in the fuel and cool the oil in the main lubrication system 162. The preheater 216 may be configured to heat the fuel to a temperature preferably from zero degrees Fahrenheit to two hundred degrees Fahrenheit, as measured at the outlet of the preheater 216. Although the preheater 216 is shown as being located within the engine 100, the preheater 216 may also be suitably located in other parts of the aircraft 10 (e.g., fuselage 12, wing 14, or pylon 18 (see [link to relevant documentation]). Figure 1 ))middle.

[0038] The fuel system 200 also includes a main filter 218 in fluid communication with the fuel delivery assembly 202. The main filter 218 is configured to remove contaminants that may be present in the fuel supply and is therefore preferably located close to the fuel tank 210 and upstream of many of the main components of the fuel system 200, such as the main fuel pump 212, the first fuel metering unit 222, the second fuel metering unit 232, and the deoxygenation system 219. Figure 3 In the depicted embodiment, the main filter 218 is located downstream of the fuel tank 210, inlet pump 214, and preheater 216. Although the main filter 218 is shown as being located within the engine 100, it may also be suitably located in other parts of the aircraft 10 (e.g., fuselage 12, wing 14, or pylon 18 (see [link to relevant documentation]). Figure 1 The main filter 218 can be any suitable filter, including, for example, a mesh filter. The main filter 218 preferably has a nominal micron level of from ten to fifty microns to remove potential contaminants.

[0039] Fuel system 200 includes multiple fuel circuits: a first fuel circuit 220 and a second fuel circuit 230. Each of the first fuel circuit 220 and the second fuel circuit 230 is configured to supply fuel to each fuel nozzle 152. The first fuel circuit 220 is configured to supply fuel to the fuel nozzle 152 at a first temperature, and the second fuel circuit 230 is configured to supply fuel to the fuel nozzle 152 at a second temperature. Each fuel nozzle 152 may have multiple fuel inlets, one of which is configured to receive fuel from the first fuel circuit 220, and another of which is configured to receive fuel from the second fuel circuit 230. In some embodiments, the fuel nozzle 152 may have a first orifice (or a first set of orifices) fluidly connected to the first fuel circuit 220 and a second orifice (or a second set of orifices) fluidly connected to the second fuel circuit 230, such that fuel from each circuit is injected into the combustion chamber 154 through a separate orifice. In other embodiments, other arrangements of fuel injection orifices may be used.

[0040] Each of the first fuel circuit 220 and the second fuel circuit 230 includes a fuel metering unit (first fuel metering unit 222 and second fuel metering unit 232) in fluid communication with the fuel delivery assembly 202. Any suitable fuel metering unit 222, 232 can be used, including, for example, a metering valve. Both the first fuel metering unit 222 and the second fuel metering unit 232 are located downstream of the main fuel pump 212. The first fuel metering unit 222 is located upstream of the first fuel manifold 224, and the second fuel metering unit 232 is located upstream of the second fuel manifold 234. The first fuel manifold 224 is configured to distribute fuel from the first fuel circuit 220 to each fuel nozzle 152, and the second fuel manifold 234 is configured to distribute fuel from the second fuel circuit 230 to each fuel nozzle 152.

[0041] Fuel system 200 is configured to supply fuel to each of a first fuel metering unit 222 and a first fuel manifold 224, and fuel metering units 222 and 224 are configured to receive fuel from fuel tank 210. The first fuel metering unit 222 and the first fuel manifold 224 work together to supply fuel flow to engine 100 in a desired manner. More specifically, the first fuel metering unit 222 is configured to meter fuel and supply a desired volume of fuel to the first fuel manifold 224 at, for example, a desired flow rate, and the second fuel metering unit 232 is configured to meter fuel and supply a desired volume of fuel to the second fuel manifold 234 at, for example, a desired flow rate. The first fuel manifold 224 and the second fuel manifold 234 are fluidly connected to fuel nozzles 152 and distribute (supply) the received fuel to a plurality of fuel nozzles 152, in which, as discussed above, fuel is injected into combustion chamber 154 and burned. Adjusting the first fuel metering unit 222 and the first fuel manifold 224 changes the volume of fuel supplied to the combustion chamber 154, and thus changes the amount of propulsive thrust generated by the engine 100 to propel the aircraft 10.

[0042] In some embodiments, the first fuel metering unit 222 and the second fuel metering unit 232 can operate cooperatively such that the proportion of fuel flowing from each of the first fuel circuit 220 and the second fuel circuit 230 remains constant as the total flow rate increases or decreases. In other embodiments, the first fuel metering unit 222 and the second fuel metering unit 232 can be adjusted independently. Figure 3 In the illustrated embodiment, each of the first fuel circuit 220 and the second fuel circuit 230 includes a fuel metering unit, but other embodiments may be used, including, for example, a single fuel metering unit located upstream of both the first fuel circuit 220 and the second fuel circuit 230.

[0043] The fuel downstream of the first fuel metering unit 222 and the second fuel metering unit 232 can be further heated to improve the efficiency, performance, and durability of the gas turbine. The fuel can be used as a cooling source to improve the durability of aircraft or engine components or to extract heat from the core airflow path 121 or the CCA system 164 to improve engine thermodynamic efficiency. Each of the first fuel circuit 220 and the second fuel circuit 230 may include a heat exchanger, referred herein as a performance heat exchanger (HX), to heat the fuel in the respective circuit 220, 230. The first fuel circuit 220 may include a first performance heat exchanger 226, and the second fuel circuit 230 may include a second performance heat exchanger 236. The first performance heat exchanger 226 is located upstream of the fuel nozzle 152, and more specifically, upstream of the first fuel manifold 224; the second performance heat exchanger 236 is located upstream of the fuel nozzle 152, and more specifically, upstream of the second fuel manifold 234.

[0044] The first performance heat exchanger 226 and the second performance heat exchanger 236 can each be a heat exchanger in thermal communication with any suitable heat source (e.g., any suitable engine and / or aircraft heat source). For example, each of the first performance heat exchanger 226 and the second performance heat exchanger 236 can be in thermal communication with the hot gas path of the engine 100. Such an engine heat source can include, for example, a flow path for heating air through the engine 100, such as the core air flow path 121. Each of the first performance heat exchanger 226 and the second performance heat exchanger 236 can also be fluidly connected, for example, to a CCA system 164 to cool the HP turbine 116. Each of the first performance heat exchanger 226 and the second performance heat exchanger 236 can be thermally connected to other portions of the core air flow path 121, including the exhaust nozzle section 120. Additionally, or alternatively, in other embodiments, each of the first performance heat exchanger 226 and the second performance heat exchanger 236 can be thermally coupled to an intermediate heat transfer system, which in turn is thermally coupled to one or more systems of the engine 100 or to the flow path for air through the engine 100. Each of the first performance heat exchanger 226 and the second performance heat exchanger 236 can be thermally connected to an intermediate heat transfer system to receive heat from these heat sources.

[0045] As described above, each of the first fuel circuit 220 and the second fuel circuit 230 is configured to supply fuel to each fuel nozzle 152, wherein the first fuel circuit 220 supplies fuel to the fuel nozzle 152 at a first temperature, and the second fuel circuit supplies fuel to the fuel nozzle 152 at a second temperature. In this embodiment, the first temperature is greater than the second temperature. The first fuel circuit 220, and more specifically, the first performance heat exchanger 226, may be configured to heat the fuel to a temperature above 200 degrees Fahrenheit, and more preferably, from 300 degrees Fahrenheit to 900 degrees Fahrenheit. As described above, supplying fuel to the fuel nozzle 152 at such temperatures improves the efficiency of the gas turbine engine by reducing fuel consumption, particularly during operating conditions such as takeoff, climb, and cruise; however, supplying fuel at such temperatures can lead to problems such as an increased risk of flameout. To mitigate the risk of flameout, the second fuel circuit 230 supplies fuel to the fuel nozzle 152 at a temperature (second temperature) lower than the temperature of the fuel supplied to the fuel nozzle 152 via the first fuel circuit 220 (the first temperature). In some embodiments, the second temperature is 20 to 500 degrees Fahrenheit lower than the first temperature. The second fuel circuit 230, and more specifically, the second performance heat exchanger 236, can be configured to heat the fuel to different temperatures for different phases of flight. In some embodiments, the second fuel circuit 230 supplies fuel to the fuel nozzle 152 without any additional heating. In this case, the fuel temperature at the outlet of the main fuel pump 212 is sufficient, and the second performance heat exchanger 236 can be omitted from the second fuel circuit 230 or otherwise not operated.

[0046] The second temperature is preferably low enough that when the pilot closes the throttle valve, the heat from the fuel nozzle 152 and other components of the burner 150 will not cause the fuel supplied by the second fuel circuit 230 to evaporate. In this way, even when the throttle valve is closed, fuel can still be supplied to the combustion chamber 154 (see...). Figure 2 This provides a liquid fuel flow and can prevent flameout. In this case, the amount of liquid fuel does not need to be large. In some embodiments, the first fuel circuit 220 supplies most of the fuel to the fuel nozzles 152, and the ratio of the amount of fuel supplied by the first fuel circuit 220 to each fuel nozzle 152 to the amount of fuel supplied by the second fuel circuit 230 to each fuel nozzle 152 can range from two-tenths to twenty-tenths, and can vary for different phases of flight. For cruise, this ratio may be five to ten. The reduced temperature of the fuel supplied by the second fuel circuit 230 contradicts the beneficial effect of heating the fuel, and this effect can be mitigated by supplying fuel to the fuel nozzles 152 at these ratios.

[0047] like Figure 3As shown, the fuel system 200 may also include a deoxygenation system 219 configured to reduce the oxygen content in the fuel. Oxygen in the fuel can be a cause of thermal oxidation and coking, especially at temperatures above 300 degrees Fahrenheit, thus coking can be a problem in the first fuel circuit 220 and the second fuel circuit 230. In this embodiment, the deoxygenation system 219 is in fluid communication with the fuel delivery assembly 202 upstream of each of the first fuel circuit 220 and the second fuel circuit 230 (and more specifically, each of the first performance heat exchanger 226 and the second performance heat exchanger 236), such that the deoxygenation system 219 reduces the oxygen content of the fuel supplied to each of the first performance heat exchanger 226 and the second performance heat exchanger 236. Figure 3 As shown, the deoxygenation system 219 is also upstream of the first fuel metering unit 222, the second fuel metering unit 232, and the main fuel pump 212. The deoxygenation system 219 is downstream of the main filter 218, and more specifically, directly downstream of the main filter 218. A suitable deoxygenation system 219 is the fuel deoxygenation unit shown and described in U.S. Patent Application Publication No. 2020 / 0140114, the disclosure of which is incorporated herein by reference in its entirety. The fuel supplied by the deoxygenation system 219 may have an oxygen content of less than five parts per million (“ppm”), for example, less than three ppm, for example, less than two ppm, for example, less than one ppm, and for example, less than half a ppm.

[0048] Figure 4 This is a schematic diagram of a fuel system 200 according to another embodiment of the present disclosure. Figure 4 The embodiments shown include many of the same Figure 3 The embodiments shown use the same or similar components. In both embodiments, the same or similar components are referred to by the same reference numerals, and detailed descriptions of these components are omitted here. Figure 3 In the illustrated embodiment, each of the first fuel circuit 220 and the second fuel circuit 230 supplies fuel to each fuel nozzle 152. However, in Figure 4In the illustrated embodiment, a first fuel manifold 224 is fluidly connected to a first set of fuel nozzles, including a plurality of first fuel nozzles 152a. The first fuel manifold 224 dispenses (provides) received fuel to each of the first fuel nozzles 152a in the first set of fuel nozzles. A second fuel manifold 234 is fluidly connected to a second set of fuel nozzles, including a plurality of second fuel nozzles 152b. The second fuel manifold 234 dispenses (provides) received fuel to each of the second fuel nozzles 152b in the second set of fuel nozzles. In this embodiment, the first fuel nozzles 152a and second fuel nozzles 152b are arranged alternately such that one of the first fuel nozzles 152a is adjacent to two of the second fuel nozzles 152b, and one of the second fuel nozzles 152b is adjacent to two of the first fuel nozzles 152a. For example, in Figure 4 In the illustrated embodiment, the first fuel nozzle 152a is located in cups 1, 3, and N-1, and the second fuel nozzle 152b is located in cups 2, 4, and N. However, other arrangements can be used, in which two or three fuel nozzles from the same group may be positioned adjacent to each other. Although two fuel circuits are shown, other embodiments may include three or more fuel circuits.

[0049] The primary air flowing into and passing through combustion chamber 154, as well as the combustion process, exhibits pressure oscillations and heat release oscillations. Pressure and heat release oscillations in one cup may interact with oscillations in adjacent cups, forming a feedback loop that can increase the amplitude of these oscillations in the combustion chamber (so-called combustion dynamics). A significant increase in oscillations can lead to damage to components of engine 100 (and more specifically, combustor 150). Figure 4 The illustrated embodiment mitigates these combustion kinetics. The fuel flowing from the first fuel nozzle 152a has a different temperature than the fuel flowing from the adjacent second fuel nozzle 152b. Therefore, the combustion and pressure oscillations between the two adjacent fuel nozzles 152a, 152b differ, thereby reducing the likelihood that the oscillations will interact in an amplified manner. To mitigate these combustion kinetics, the first temperature can be 20 to 500 degrees Fahrenheit lower than the second temperature. Furthermore, this temperature difference reduces the average NOx emissions from the combustor compared to a uniform temperature between the fuel nozzles 152.

[0050] Figure 5 An arrangement for a first performance heat exchanger 226 that can be used in the fuel system discussed above is shown. In this embodiment, the first performance heat exchanger 226 is a heat exchanger surrounding a conduit of a fuel delivery assembly 202 that supplies fuel to a first fuel manifold 224. Although this arrangement has been described with respect to a first fuel circuit 220 and a first performance heat exchanger 226, it can also be used with a second performance heat exchanger 236 in a second fuel circuit 230.

[0051] Figure 6 An arrangement is shown for a heat exchanger 242 on which a fuel manifold 244 can be installed. Figure 4 In the illustrated embodiment, multiple fuel circuits (first fuel circuit 220 and second fuel circuit 230) are used to supply fuel to adjacent fuel nozzles 152 at different temperatures to mitigate oscillations and combustion dynamics. In this regard, multiple heat exchangers 242 are used to regulate the temperature. The heat exchangers 242 in this embodiment are configured similarly to those referenced above. Figure 5 The first performance heat exchanger 226 discussed can be thermally coupled to any suitable heat source as discussed above for the first performance heat exchanger 226. Alternatively, one or more heat exchangers 242 in this embodiment can be thermally coupled to a cooling source to reduce the temperature of the fuel. Such cooling sources include, for example, air flowing through the core airflow path 121. The fuel manifold 244 is configured similarly to the first fuel manifold 224 and the second fuel manifold 234 discussed above, and is fluidly connected to the fuel nozzle 152 ( Figure 6 (Seen as a cup). Fuel manifold 244 distributes (provides) the received fuel to each cup, and in some embodiments, a single fuel circuit including fuel manifold 244 may be used. Fuel manifold 244 is a conduit between adjacent cups, and in this embodiment, each of the heat exchangers 242 surrounds the conduit (fuel manifold 244) between adjacent cups (fuel nozzles).

[0052] In this embodiment, the heat exchanger 242 can be placed between adjacent cups and used to heat the fuel supplied to the adjacent cups (fuel nozzles 152) to different temperatures, thereby mitigating the combustion dynamics discussed above. The heat exchanger 242 in this embodiment can be configured as described above... Figure 4 The temperature difference discussed provides fuel to each cup. For example... Figure 6 As shown, for example, one of the heat exchangers 242 can be fluidly positioned downstream of cup 1 (the first fuel nozzle) and upstream of cup 2 (the second nozzle). Fuel is supplied to cup 1 at the temperature of the fuel circuit. The fuel flowing through cup 1 is heated or cooled by the heat exchanger 242 and supplied to cup 2 at different temperatures. Figure 6 In the illustrated embodiment, the heat exchanger 242 is positioned between each cup (fuel nozzle 152), but other arrangements may be used, such as an arrangement in which the heat exchanger 242 is positioned with two or more cups therebetween.

[0053] Figure 7 Another arrangement of the heat exchanger 242 is shown. Figure 6 In the arrangement shown, the heat exchanger 242 is located around the fuel manifold 244. Figure 7In this embodiment, fuel nozzle 152 is fluidly connected to fuel manifold 244 via at least one fuel supply line. In this embodiment, fuel nozzle 152 is fluidly connected to fuel manifold 244 via a first fuel supply line 246 and a second fuel supply line 248. Each fuel nozzle may include multiple fuel inlets, including a first inlet fluidly connected to the first fuel supply line 246 to receive fuel from the first fuel supply line 246 and a second inlet fluidly connected to the second fuel supply line 248 to receive fuel from the second fuel supply line 248. Instead of (or in addition to) placing heat exchanger 242 around fuel manifold 244, heat exchanger 242 may be placed around at least one fuel supply line downstream of fuel manifold 244 and upstream of fuel nozzle 152. In this embodiment, heat exchanger 242 is placed around the first fuel supply line 246, and heat exchanger 242 is not placed around the second fuel supply line 248. In addition to controlling the temperature of the fuel by the heat flowing into the heat exchanger 242, the temperature of the fuel flowing out of the fuel nozzle 152 can be controlled by controlling the distribution (proportion) of the fuel flowing between each of the first fuel supply line 246 and the second fuel supply line 248.

[0054] The embodiments discussed herein can be used to mitigate the effects of fuel flow oscillations and other combustion dynamics, thereby preventing combustion problems such as flameout. Further aspects of this disclosure are provided by the subject matter of the following clauses.

[0055] A gas turbine engine includes: a combustor, a plurality of fuel injectors, a first fuel circuit, and a second fuel circuit. The combustor includes a combustion chamber. The plurality of fuel injectors inject fuel into the combustion chamber of the combustor. The first fuel circuit includes a first fuel manifold fluidly connected to at least one of the plurality of fuel injectors to distribute the fuel to the at least one fuel injector at a first temperature. The second fuel circuit includes a second fuel manifold fluidly connected to at least one of the plurality of fuel injectors to distribute the fuel to the at least one fuel injector at a second temperature. The second temperature is lower than the first temperature.

[0056] According to the gas turbine engine described in the foregoing clause, the first temperature is above 200 degrees Fahrenheit.

[0057] According to any of the preceding clauses, the gas turbine engine wherein the second temperature is 20 to 500 degrees Fahrenheit lower than the first temperature.

[0058] According to any of the foregoing clauses, in a gas turbine engine, wherein the first fuel circuit and the second fuel circuit are fluidly connected to each of the plurality of fuel nozzles to supply fuel to each of the plurality of fuel nozzles. Each fuel nozzle has a plurality of fuel inlets. One of the plurality of fuel inlets is fluidly connected to the first fuel circuit to receive the fuel from the first fuel circuit, and another of the plurality of fuel inlets is fluidly connected to the second fuel circuit to receive fuel from the second fuel circuit.

[0059] According to any of the preceding clauses, in a gas turbine engine, wherein the second fuel circuit includes a heat exchanger thermally connected to a heat source to heat the fuel to the second temperature.

[0060] According to any of the foregoing clauses, in a gas turbine engine, the heat exchanger is located upstream of the second fuel manifold relative to the flow of the fuel in the second fuel circuit.

[0061] According to any of the preceding clauses, in a gas turbine engine, wherein the first fuel circuit includes a heat exchanger thermally connected to a heat source to heat the fuel to the first temperature.

[0062] According to any of the foregoing clauses, in a gas turbine engine, the heat exchanger is located upstream of the first fuel manifold relative to the flow of fuel in the first fuel circuit.

[0063] The gas turbine engine according to any of the foregoing clauses further includes a core airflow path, the core airflow path including a compressor section, the combustor and the turbine section, wherein one of the compressor section, the combustor or the turbine section is the heat source.

[0064] In a gas turbine engine according to any of the foregoing clauses, the plurality of fuel nozzles includes a first nozzle and a second nozzle. The second nozzle is adjacent to the first nozzle. The first nozzle is fluidly coupled to a first fuel manifold to receive fuel from the first fuel manifold. The second nozzle is fluidly coupled to a second fuel manifold to receive fuel from the second fuel manifold.

[0065] According to any of the preceding clauses, the first nozzle is one of a plurality of first groups of fuel nozzles, and the second nozzle is one of a plurality of second groups of fuel nozzles.

[0066] According to any of the preceding clauses, in the gas turbine engine, the fuel nozzles in the first group of fuel nozzles and the fuel nozzles in the second group of fuel nozzles are arranged alternately, such that one fuel nozzle in the first group of fuel nozzles is adjacent to two fuel nozzles in the second group of fuel nozzles and one fuel nozzle in the second group of fuel nozzles is adjacent to two fuel nozzles in the first group of fuel nozzles.

[0067] A gas turbine engine includes: a combustor, a plurality of fuel nozzles, a fuel manifold, and at least one heat exchanger. The combustor includes a combustion chamber. The plurality of fuel nozzles inject fuel into the combustion chamber of the combustor. The plurality of fuel nozzles includes a first fuel nozzle and a second fuel nozzle. The fuel manifold is fluidly connected to the plurality of fuel nozzles to distribute the fuel to the fuel nozzles and to supply the fuel to the first fuel nozzle at a first temperature. The at least one heat exchanger is located downstream of the first fuel nozzle and upstream of the second fuel nozzle. The heat exchanger is connected to the fuel manifold such that the fuel manifold supplies the fuel to the second fuel nozzle at a second temperature different from the first temperature.

[0068] The gas turbine engine according to any of the foregoing clauses further includes a plurality of heat exchangers. At least one of the plurality of fuel nozzles is located between adjacent heat exchangers.

[0069] In any of the preceding clauses, the gas turbine engine wherein the fuel manifold is a pipe and the heat exchanger surrounds the pipe.

[0070] According to any of the foregoing clauses, the gas turbine engine wherein the first temperature is above 200 degrees Fahrenheit.

[0071] According to any of the preceding clauses, the gas turbine engine wherein the second temperature is 20 to 200 degrees Fahrenheit higher than the first temperature.

[0072] A gas turbine engine includes: a combustor, a plurality of fuel nozzles, a fuel manifold, a first fuel supply line, a second fuel supply line, and a heat exchanger. The combustor includes a combustion chamber. The plurality of fuel nozzles inject fuel into the combustion chamber of the combustor. Each fuel nozzle includes a plurality of fuel inlets, including a first inlet and a second inlet. The fuel manifold is fluidly connected to the plurality of fuel nozzles to distribute the fuel to each of the plurality of fuel nozzles at a first temperature. The first fuel supply line is fluidly connected to the first inlet of the fuel manifold. The second fuel supply line is fluidly connected to the second inlet of the fuel manifold. The heat exchanger is fluidly connected to the first fuel supply line. The heat exchanger is positioned downstream of the fuel manifold and upstream of the fuel nozzles, and is thermally coupled to a heat source to heat the fuel to a second temperature different from the first temperature.

[0073] According to any of the preceding clauses, the gas turbine engine wherein the first fuel supply line is a pipe and the heat exchanger surrounds the pipe.

[0074] According to any of the preceding clauses, the gas turbine engine wherein the second temperature is 20 to 200 degrees Fahrenheit higher than the first temperature.

[0075] Although the foregoing description is directed to preferred embodiments, other variations and modifications will be apparent to those skilled in the art and can be made without departing from the spirit or scope of this disclosure. Furthermore, features described in connection with one embodiment may be used in conjunction with other embodiments, even if not explicitly stated above.

Claims

1. A gas turbine engine characterized by, include: A fuel conduit fluidly connected to a fuel source to receive fuel from the fuel source; A first fuel circuit, fluidly connected to the fuel conduit, for receiving a first portion of the fuel from the fuel source, the first fuel circuit including a first fuel manifold for dispensing the first portion of the fuel at a first temperature; A second fuel circuit, arranged in parallel with the first fuel circuit and fluidly connected to the fuel conduit, is provided to receive a second portion of the fuel from the fuel source. The second fuel circuit includes a second fuel manifold for distributing the second portion of the fuel at a second temperature, which is lower than the first temperature. A burner, the burner including a combustion chamber; and A plurality of fuel nozzles inject a first portion and a second portion of the fuel into the combustion chamber of the burner. Each of the plurality of fuel nozzles is fluidly connected via a first fuel manifold to a first fuel circuit to receive the first portion of the fuel at a first temperature, and fluidly connected via a second fuel manifold to a second fuel circuit to receive the second portion of the fuel at a second temperature. Each of the plurality of fuel nozzles is fluidly connected to both the first fuel circuit and the second fuel circuit to receive the first portion of the fuel at the first temperature and the second portion of the fuel at the second temperature.

2. The gas turbine engine according to claim 1, characterized in that, The first temperature is between 300 and 900 degrees Fahrenheit.

3. The gas turbine engine according to claim 1, characterized in that, The second temperature is 20 to 500 degrees Fahrenheit lower than the first temperature.

4. The gas turbine engine according to claim 1, characterized in that, Each of the plurality of fuel nozzles has a plurality of fuel inlets, one of the plurality of fuel inlets of each of the plurality of fuel nozzles being fluidly connected to the first fuel circuit to receive the first portion of the fuel from the first fuel circuit, and another of the plurality of fuel inlets of each of the plurality of fuel nozzles being fluidly connected to the second fuel circuit to receive the second portion of the fuel from the second fuel circuit.

5. The gas turbine engine according to claim 1, characterized in that, The second fuel circuit includes a heat exchanger thermally connected to a heat source to heat the second portion of the fuel to the second temperature.

6. The gas turbine engine according to claim 5, characterized in that, The heat exchanger is located upstream of the second fuel manifold relative to the second portion of the fuel flowing in the second fuel circuit.

7. The gas turbine engine according to claim 1, characterized in that, The first fuel circuit includes a heat exchanger thermally connected to a heat source to heat the first portion of the fuel to the first temperature.

8. The gas turbine engine according to claim 7, characterized in that, The heat exchanger is located upstream of the first fuel manifold relative to the first portion of the fuel flowing in the first fuel circuit.

9. The gas turbine engine according to claim 7, characterized in that, It further includes a core airflow path, which includes a compressor section, the burner, and a turbine section, wherein one of the compressor section, the burner, or the turbine section is the heat source.

10. The gas turbine engine according to claim 1, characterized in that, Each of the plurality of fuel nozzles has a plurality of orifices, one of which is fluidly connected to the first fuel circuit to inject fuel from the first fuel circuit into the combustion chamber, and another of which is fluidly connected to the second fuel circuit to inject fuel from the second fuel circuit into the combustion chamber.

11. The gas turbine engine according to claim 1, characterized in that, The burner mentioned above is an annular burner.

12. The gas turbine engine according to claim 11, characterized in that, The plurality of fuel nozzles are arranged in a ring configuration.

13. The gas turbine engine according to claim 11, characterized in that, The plurality of fuel nozzles are aligned along the circumferential direction of the burner.

14. The gas turbine engine according to claim 1, characterized in that, The first fuel circuit includes a first fuel metering unit fluidly connected to the first fuel manifold, and the second fuel circuit includes a second fuel metering unit fluidly connected to the second fuel manifold.

15. The gas turbine engine according to claim 14, characterized in that, The first fuel metering unit is located upstream of the first fuel manifold relative to the first portion of the fuel flowing in the first fuel circuit, and the first fuel metering unit is operable to provide a first flow rate of the first portion of the fuel to the first fuel manifold. The second fuel metering unit is located upstream of the second fuel manifold relative to the second portion of the fuel flowing in the second fuel circuit, and the second fuel metering unit is operable to provide a second flow rate of the second portion of the fuel to the second fuel manifold.

16. The gas turbine engine according to claim 15, characterized in that, The first fuel metering unit and the second fuel metering unit are configured to operate in concert to maintain a constant fuel ratio between the first flow rate and the second flow rate when each of the first flow rate and the second flow rate is changed.

17. The gas turbine engine according to claim 15, characterized in that, The ratio of the first flow rate to the second flow rate is between 20% and 20%.

18. The gas turbine engine according to claim 15, characterized in that, The ratio of the first flow rate to the second flow rate is five to ten.

19. The gas turbine engine according to claim 15, characterized in that, Each of the plurality of fuel nozzles includes a common fuel flow path fluidly connected to both the first fuel circuit and the second fuel circuit to receive a first portion of the fuel at the first temperature from the first fuel circuit and a second portion of the fuel at the second temperature from the second fuel circuit, respectively.

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