Aircraft system with thermal management system

Abstract

An aircraft system is provided. An aircraft system includes: a gas turbine engine; a fuel system having a fuel tank and a fuel delivery system fluidly connecting the fuel tank to the gas turbine engine; the electric driving assembly comprises a component module; and a thermal management system, the thermal management system comprising: a thermal fluid loop; a heat source heat exchanger in thermal communication with the component module and the thermal fluid loop; a gas flow radiator heat exchanger in thermal communication with the thermal fluid loop and further configured to be in thermal communication with the cooling gas flow during operation of the gas turbine engine; and a fuel radiator heat exchanger in selective thermal communication with the thermal fluid loop, the fuel tank, or both, the fuel radiator heat exchanger being in fluid communication with the fuel tank independently of the fuel delivery system during at least a first operating condition.

Description

[0001] Priority information

[0002] This application claims priority from Italian patent application serial number 102023000023157 filed on November 3, 2023. Technical Field

[0003] The present disclosure relates to an aircraft system having a thermal management system and a control method thereof. Background Art

[0004] Conventional commercial aircraft generally include a fuselage, a pair of wings, and a propulsion system that generates thrust. Such a propulsion system generally includes at least two aircraft engines, such as turbofan jet engines. Each turbofan jet engine is generally mounted to one of the wings of the aircraft, such as in a suspended position below the wing that is separate from the wing and the fuselage.

[0005] Hybrid electric propulsion systems are being developed to improve the efficiency of such conventional commercial aircraft. A hybrid electric propulsion system typically includes one or more propellers. For example, a propeller may include an electric motor operably coupled to an aircraft engine. Although many advances have been made, further efficiency improvements and integrated solutions for propellers of hybrid electric propulsion systems are desired. BRIEF DESCRIPTION OF THE DRAWINGS

[0006] A full and enabling disclosure of the present disclosure, including the best mode thereof, to one of ordinary skill in the art is set forth in the specification with reference to the accompanying drawings, in which:

[0007] Figure 1 A schematic top view of an aircraft is provided according to various exemplary embodiments of the present disclosure.

[0008] Figure 2 A schematic diagram of a gas turbine engine according to an exemplary aspect of the present disclosure is provided.

[0009] Figure 3A A schematic diagram of an aircraft system according to an exemplary aspect of the present disclosure is provided.

[0010] Figure 3B A schematic diagram of an aircraft system according to another exemplary aspect of the present disclosure is provided.

[0011] Figure 4A A schematic diagram of an aircraft system according to yet another exemplary aspect of the present disclosure is provided.

[0012] Figure 4B A schematic diagram of an aircraft system according to yet another exemplary aspect of the present disclosure is provided.

[0013] Figure 5A schematic diagram of an aircraft system according to yet another exemplary aspect of the present disclosure is provided.

[0014] Figure 6 A schematic diagram of an aircraft system according to yet another exemplary aspect of the present disclosure is provided.

[0015] Figure 7 A flow chart of a method of operating an aircraft system according to an exemplary aspect of the present disclosure is provided.

[0016] Figure 8 A schematic diagram of a gas turbine engine according to another exemplary aspect of the present disclosure is provided.

[0017] Fig. 9 A schematic diagram of a gas turbine engine according to yet another exemplary aspect of the present disclosure is provided. DETAILED DESCRIPTION

[0018] Reference will now be made in detail to the present embodiments of the present disclosure, one or more examples of which are illustrated in the accompanying drawings. The detailed description uses numerical and letter references to refer to features in the drawings. Like or similar reference numbers in the drawings and description have been used to refer to like or similar parts of the present disclosure.

[0019] The word "exemplary" is used herein to mean "serving as an example, instance, or illustration." Any implementation described herein as "exemplary" is not necessarily to be construed as preferred or advantageous over other implementations. In addition, unless expressly stated otherwise, all embodiments described herein should be considered exemplary.

[0020] The singular forms "a," "an," and "the" include plural references unless the context clearly dictates otherwise.

[0021] The term "at least one" in a context such as "at least one of A, B, and C" means only A, only B, only C, or any combination of A, B, and C.

[0022] The term “turbomachine” refers to a machine that includes one or more compressors, a heat generating section (eg, a combustion section), and one or more turbines that together generate a torque output.

[0023] The term "gas turbine engine" refers to an engine having a turbine as all or part of its power source. Example gas turbine engines include turbofan engines, turboprop engines, turbojets, turboshaft engines, etc., as well as hybrid-electric versions of one or more of these engines.

[0024] The term "combustion section" refers to any heat addition system for a turbomachine. For example, the term combustion section may refer to a section that includes one or more of a deflagration combustion assembly, a rotating detonation combustion assembly, a pulse detonation combustion assembly, or other suitable heat addition assemblies. In certain example embodiments, the combustion section may include an annular combustor, a can combustor, a tubular combustor, a trapped vortex combustor (TVC), or other suitable combustion systems, or a combination thereof.

[0025] As used herein, the terms “first,” “second,” and “third” may be used interchangeably to distinguish one component from another, and are not intended to indicate the position or importance of each component.

[0026] The present disclosure generally relates to an aircraft system having an electric drive assembly and a thermal management system capable of managing the thermal load of the electric drive system under various operating conditions by using an auxiliary radiator. More specifically, the aircraft system of the present disclosure includes: a gas turbine engine; a fuel system having a fuel tank and a fuel delivery system that fluidly connects the fuel tank to the gas turbine engine; an electric drive assembly having a component module; and a thermal management system. The thermal management system includes: a thermal fluid loop; a heat source heat exchanger that is thermally connected to the component module and the thermal fluid loop; and an airflow radiator heat exchanger that is thermally connected to the thermal fluid loop and is further configured to be thermally connected to a cooling airflow during operation of the gas turbine engine. In addition, the thermal management system of the present disclosure also includes a fuel radiator heat exchanger that is selectively thermally connected to the thermal fluid loop, the fuel tank, or both. During at least a first operating condition, the fuel radiator heat exchanger is in fluid communication with the fuel tank independently of the fuel delivery system.

[0027] In this manner, it should be appreciated that the airflow radiator heat exchanger can be configured to provide a desired amount of heat rejection for the thermal management system to maintain the temperature of the component modules of the electric drive system within a desired operating temperature range during certain operating conditions (e.g., cruising). However, when the mass flow rate of the airflow reaching the airflow radiator heat exchanger is insufficient, when the component modules generate a large amount of heat, or both, the thermal management system can thermally connect the heat loop to the fuel flow via the fuel radiator heat exchanger to transfer heat to the fuel flow and supplement the heat rejection of the airflow radiator heat exchanger.

[0028] Referring now to the drawings, in which like numerals refer to like elements throughout, Figure 1 A schematic top view of an exemplary aircraft 100 is provided that may incorporate various embodiments of the present disclosure. Figure 1As shown, for reference, the aircraft 100 defines a longitudinal direction L1 and a lateral direction L2. The aircraft 100 further defines a longitudinal centerline 114 extending therethrough along the longitudinal direction L1. The aircraft 100 extends, for example, along the longitudinal direction L1 between a front end 116 and a rear end 118. In addition, the aircraft 100 includes a fuselage 112 that extends longitudinally from the front end 116 of the aircraft 100 to the rear end 118 of the aircraft 100. The aircraft 100 also includes a tail 119 at the rear end 118 of the aircraft 100. In addition, the aircraft 100 includes a wing assembly that includes a first port wing 120 and a second starboard wing 122. The first and second wings 120, 122 each extend laterally outward relative to the longitudinal centerline 114. The first wing 120 and a portion of the fuselage 112 together define a first side 124 of the aircraft 100, and the second wing 122 and another portion of the fuselage 112 together define a second side 126 of the aircraft 100. For the depicted embodiment, the first side 124 of the aircraft 100 is configured as the port side of the aircraft 100, and the second side 126 of the aircraft 100 is configured as the starboard side of the aircraft 100.

[0029] Aircraft 100 includes various control surfaces. For this embodiment, each wing 120, 122 includes one or more leading edge flaps 128 and one or more trailing edge flaps 130. Aircraft 100 further includes, or more specifically, the tail 119 of aircraft 100 includes: a vertical stabilizer 132, which has a rudder flap (not shown) for yaw control; and a pair of horizontal stabilizers 134, each horizontal stabilizer having a lift flap 136 for pitch control. Fuselage 112 additionally includes an outer surface or skin 138. It should be understood that in other exemplary embodiments of the present disclosure, aircraft 100 may additionally or alternatively include any other suitable configuration. For example, in other embodiments, aircraft 100 may include any other control surface configuration.

[0030] Figure 1 The exemplary aircraft 100 also includes a hybrid electric propulsion system 150. For this embodiment, the hybrid electric propulsion system 150 has a first propeller 160 and a second propeller 170, both of which are operable to generate thrust. The first propeller 160 is mounted to the first wing 120, and the second propeller 170 is mounted to the second wing 122. In addition, for the depicted embodiment, the first propeller 160 and the second propeller 170 are both configured as underwing mounted configurations. However, in other exemplary embodiments, one or both of the first and second propellers 160, 170 can be mounted in any other suitable location in other exemplary embodiments.

[0031] The first propeller 160 includes a gas turbine engine 162 and an electric motor 164 operably coupled to the gas turbine engine 162. The electric motor 164 may be a generator, an electric motor, or a combination generator / motor. For this example embodiment, the electric motor 164 is a combination generator / motor. In this way, when operating as a generator, the electric motor 164 can generate electricity when driven by the gas turbine engine 162. When operating as an electric motor, the electric motor 164 can drive or start the fan spool of the gas turbine engine 162. In addition, for this example embodiment, the gas turbine engine 162 is configured as a turbofan engine, and therefore, the first propeller 160 is configured as a hybrid electric turbofan engine.

[0032] Likewise, the second propulsor 170 includes a gas turbine engine 172 and an electric motor 174 operably coupled to the gas turbine engine 172. The electric motor 174 may be a generator, an electric motor, or a combined generator / motor. For this example embodiment, the electric motor 174 is a combined generator / motor. In this manner, when operating as a generator, the electric motor 174 may generate electricity when driven by the gas turbine engine 172. When operating as an electric motor, the electric motor 174 may drive or start a fan spool of the gas turbine engine 172. Furthermore, for this example embodiment, the gas turbine engine 172 is configured as a turbofan engine, and therefore, the second propulsor 170 is configured as a hybrid electric turbofan engine.

[0033] The hybrid electric propulsion system 150 also includes a power source 180, which can be electrically connected to the motors 164, 174, and in some embodiments, can be electrically connected to other electrical loads. The power source 180 can be configured as an electrical energy storage unit (e.g., one or more batteries, such as one or more lithium-ion batteries, or alternatively can be configured as any other suitable electrical energy storage device, such as a supercapacitor), one or more fuel cells (e.g., one or more proton electron membrane fuel cells and / or solid oxide fuel cells), etc. It is worth noting that when the power source 180 is one or more fuel cells, the hybrid electric propulsion system 150 can also include a separate fuel storage tank 181 containing fuel for the one or more fuel cells (e.g., liquid and / or gaseous hydrogen, methane, etc.).

[0034] The hybrid electric propulsion system 150 also includes a power management system having a controller 182 and a power bus 184. The electric machines 164, 174, the electrical energy storage unit 180, and the controller 182 may each be electrically connected to one another via one or more electrical wires 186 of the power bus 184. For example, the power bus 184 may include various switches or other power electronics that may be moved to selectively electrically connect various components of the hybrid electric propulsion system 150. Additionally, the power bus 184 may also include power electronics (such as inverters, converters, rectifiers, etc.) for regulating or converting power within the hybrid electric propulsion system 150.

[0035] The controller 182 is configured to distribute power between the various components of the hybrid electric propulsion system 150. For example, the controller 182 can control the power electronics of the power bus 184 to provide power to or draw power from the various components (such as the motors 164, 174) to operate the hybrid electric propulsion system 150 between various operating modes and to perform various functions. This is schematically depicted as wires 186 of the power bus 184 extending through the controller 182.

[0036] Still reference Figure 1 As schematically shown, it will be appreciated that the aircraft 100 also includes a fuel system having a fuel tank 190 located in each of the first wing 120 and the second wing 122, and a fuel delivery system (although not depicted) fluidly connecting the fuel tank 190 to the gas turbine engines 162, 172.

[0037] Reference now Figure 2 , Figure 2 is a schematic cross-sectional view of a propeller according to an exemplary embodiment of the present disclosure. In certain exemplary embodiments, the propeller may be Figure 1 One of the first or second thrusters 160, 170.

[0038] for Figure 2 In an embodiment of the present invention, the propulsor is configured as a hybrid electric high bypass turbofan jet engine, referred to herein as engine 200. Figure 2 As shown in , engine 200 defines an axial direction A (extending parallel to a longitudinal centerline 212 provided for reference), a radial direction R, and a circumferential direction C extending about longitudinal centerline 212. Generally, engine 200 includes a fan section 214 and a turbine 216 disposed downstream of fan section 214.

[0039] The depicted exemplary turbine 216 generally includes a substantially tubular casing 218 defining an annular inlet 220. The casing 218 encloses, in serial flow relationship: a compressor section including a supercharger or low pressure (LP) compressor 222 and a high pressure (HP) compressor 224; a combustion section 226; a turbine section including a high pressure (HP) turbine 228 and a low pressure (LP) turbine 230; and an ejection exhaust nozzle section 232. A high pressure (HP) shaft 234 (which may additionally or alternatively be a spool) drivingly connects the HP turbine 228 to the HP compressor 224. A low pressure (LP) shaft 236 (which may additionally or alternatively be a spool) drivingly connects the LP turbine 230 to the LP compressor 222. The compressor section, combustion section 226, turbine section, and ejection exhaust nozzle section 232 together define a working gas flow path 237.

[0040] For the depicted embodiment, the fan section 214 includes a fan 238 having a plurality of fan blades 240 coupled to a disk 242 in a spaced-apart manner. As depicted, the fan blades 240 extend outwardly from the disk 242 generally in a radial direction R. Each fan blade 240 can rotate relative to the disk 242 about a pitch axis P by virtue of the fan blades 240 being operably coupled to an appropriate pitch mechanism 244, which is configured to collectively change the pitch of the fan blades 240, for example, in unison. The engine 200 also includes a power gearbox 246, and the fan blades 240, the disk 242, and the pitch mechanism 244 can rotate together about the longitudinal centerline 212 across the power gearbox 246 via the LP shaft 236. The power gearbox 246 includes a plurality of gears for adjusting the rotational speed of the fan 238 relative to the rotational speed of the LP shaft 236 so that the fan 238 can rotate at a more efficient fan speed.

[0041] Still reference Figure 2 In the exemplary embodiment of the present invention, disk 242 is covered by a rotatable front hub 248 (sometimes also referred to as a “spinner”) of fan section 214 . Front hub 248 is aerodynamically shaped to facilitate airflow through a plurality of fan blades 240 .

[0042] Additionally, the exemplary fan section 214 includes an annular fan case or outer nacelle 250 that circumferentially surrounds at least a portion of the fan 238 and / or turbine 216. It should be appreciated that in the depicted embodiment, the nacelle 250 is supported relative to the turbine 216 by a plurality of circumferentially spaced outlet guide vanes 252. Furthermore, a downstream section 254 of the nacelle 250 extends over an outer portion of the turbine 216 to define a bypass airflow passage 256 therebetween.

[0043] During operation of the engine 200, a quantity of air 258 enters the engine 200 through the nacelle 250 and the associated inlet 260 of the fan section 214. As the quantity of air 258 passes through the fan blades 240, a first portion of air 262 is directed or directed into the bypass airflow passage 256, and a second portion of air 264, as indicated by arrows 264, is directed or directed into the working gas flow path 237, or more specifically, into the LP compressor 222. The ratio between the first portion of air 262 and the second portion of air 264 is generally referred to as the bypass ratio. The pressure of the second portion of air 264 is then increased as the second portion of air 264 is directed through the HP compressor 224 and into the combustion section 226, where the second portion of air 264 is mixed with fuel and combusted to provide combustion gases 266. A fuel delivery system 290 of a fuel system of an aircraft incorporating the engine 200 is depicted, which provides a fuel flow to the combustors of the combustion section 226.

[0044] The combustion gases 266 are directed through the HP turbine 228, wherein a portion of the thermal and / or kinetic energy from the combustion gases 266 is extracted via sequential stages of HP turbine stator blades 268 coupled to the casing 218 and HP turbine rotor blades 270 coupled to the HP shaft 234, thereby rotating the HP shaft 234 to support operation of the HP compressor 224. The combustion gases 266 are then directed through the LP turbine 230, wherein a second portion of the thermal and kinetic energy is extracted from the combustion gases 266 via sequential stages of LP turbine stator blades 272 coupled to the casing 218 and LP turbine rotor blades 274 coupled to the LP shaft 236, thereby rotating the LP shaft 236 to support operation of the LP compressor 222 and / or rotation of the fan 238.

[0045] The combustion gases 266 are then directed through the jet exhaust nozzle section 232 of the turbine 216 to provide propulsive thrust. At the same time, the pressure of the first portion of air 262 is significantly increased, also providing propulsive thrust, as it is directed through the bypass airflow passage 256 before being discharged from the fan nozzle exhaust section 276 of the engine 200. The HP turbine 228, the LP turbine 230, and the jet exhaust nozzle section 232 at least partially define a hot gas path 278 for directing the combustion gases 266 through the turbine 216.

[0046] Figure 2200 , and more specifically, a motor 280. The motor 280 is operably coupled to the engine 200 for, for example, adding power to the engine 200, extracting power from the engine 200, or both. The motor 280 is depicted as being embedded within the engine 200 at a location inside the working gas flow path 237. However, in other exemplary embodiments, the electric drive assembly may additionally or alternatively include a motor located at an under-hood location (e.g., below the housing 218 and outside the working gas flow path 237) or any other suitable location.

[0047] However, it should be understood that Figure 2 The exemplary engine 200 depicted in FIG. 2 is merely an example, and in other exemplary embodiments, the engine 200 may have any other suitable configuration. For example, although the depicted engine 200 is configured as a ducted gas turbine engine (i.e., including an outer nacelle 250), in other embodiments, the engine 200 may be a non-ducted gas turbine engine (such that the fan 238 is a non-ducted fan and the outlet guide vanes 252 are cantilevered from the casing 218). Additionally or alternatively, although the depicted engine 200 is configured as a geared gas turbine engine (i.e., including a power gearbox 246) and a variable pitch gas turbine engine (i.e., including a fan 238 configured as a variable pitch fan), in other embodiments, the engine 200 may be additionally or alternatively configured as a direct drive gas turbine engine (such that the LP shaft 236 rotates at the same speed as the fan 238), a fixed pitch gas turbine engine (such that the fan 238 includes fan blades 240 that are not rotatable about the pitch axis P), or both. It should also be understood that in other exemplary embodiments, aspects of the present disclosure may be incorporated into any other suitable gas turbine engine. For example, in other exemplary embodiments, aspects of the present disclosure may be incorporated into, for example, a turbopropeller gas turbine engine, a turboshaft gas turbine engine, or a turbojet gas turbine engine, as the case may be.

[0048] Reference now Figure 3A , an aircraft system 300 according to an exemplary aspect of the present disclosure is provided. The system generally includes a hybrid electric propulsion system 302, a fuel system 304, an electric drive assembly 306, and a thermal management system 308.

[0049] In the depicted embodiment, the hybrid electric propulsion system 302 includes a gas turbine engine 310. In at least some exemplary embodiments, Figure 3A The gas turbine engine 310 shown in FIG. 3 may be configured similarly to Figure 2 An exemplary engine 200 of Figure 1One or more of the gas turbine engines 162 , 172 , or configured in any other suitable manner.

[0050] Fuel system 304 includes fuel tank 312 and fuel delivery system 314. More specifically, in the illustrated embodiment, fuel tank 312 is an aircraft fuel tank. For example, in at least some exemplary embodiments, fuel tank 312 may be configured similarly to Figure 1 A fuel delivery system 314 fluidly connects the fuel tank 312 to the gas turbine engine 310 , and more specifically, the fuel delivery system 314 fluidly connects the fuel tank 312 to a combustor 316 within a combustion section of the gas turbine engine 310 .

[0051] Electric drive assembly 306 includes component module 318, which may be any portion of the electric section of electric drive assembly 306. As will be appreciated, the term "electric drive assembly" refers to a component configured to assist the aircraft in generating thrust for the aircraft. For example, in certain exemplary embodiments, component module 318 may be a battery (see, e.g., Figure 1 ), a fuel cell (e.g., a proton electron membrane fuel cell or a solid oxide fuel cell integrated with a gas turbine engine 310 and configured to generate electricity), an electric motor (see, e.g., Figure 2 an electric machine), power electronics associated with one or more of the above components, or a combination thereof. Figure 3A Only a single component module 318 is depicted in FIG. 3 , but other embodiments may include multiple component modules 318 arranged in series or parallel to form the electric drive assembly 306 .

[0052] The thermal management system 308 generally includes a thermal fluid loop 320, a heat source heat exchanger 322, an airflow radiator heat exchanger 324, and a fuel radiator heat exchanger 326. Notably, to facilitate the desired flow through the thermal fluid loop 320 in the illustrated embodiment, the thermal management system 308 includes a thermal management system (TMS) pump 328 in fluid communication with the thermal fluid loop 320 for providing a flow of thermal fluid through the thermal fluid loop 320, and an expansion tank 330 for helping to facilitate changes in fluid volume and temperature during operating conditions of the aircraft system 300.

[0053] The heat source heat exchanger 322 is in thermal communication with the component modules 318 of the electric drive assembly 306 and the thermal fluid loop 320. The heat source heat exchanger 322 is designed to transfer heat from the component modules 318 to the thermal fluid via the thermal fluid loop 320. In doing so, the heat source heat exchanger 322 effectively cools the component modules 318 of the electric drive assembly 306.

[0054] The airflow radiator heat exchanger 324 is in thermal communication with the hot fluid loop 320 and is designed to be in thermal communication with the cooling airflow 332 during operation of the gas turbine engine 310. In the illustrated embodiment, the airflow radiator heat exchanger 324 is intended to receive an airflow passing through the gas turbine engine 310 (such as through the working gas flow path of the gas turbine engine 310 (see, e.g., Figure 2 The working gas flow path 237) in the embodiment of the present invention may be a bypass passage of the gas turbine engine 310 (see, for example, Figure 2 320). More specifically, for the depicted embodiment, the airflow radiator heat exchanger 324 is configured to receive a bleed airflow from a compressor 334 of the gas turbine engine 310. The airflow from the airflow radiator heat exchanger 324 may be discharged to any suitable location (e.g., the environment). Therefore, it should be understood that the airflow radiator heat exchanger 324 is designed to transfer heat from the hot fluid through the hot fluid loop 320 to the cooling airflow 332.

[0055] Brief reference Figure 3B , an aircraft system 300 according to another exemplary aspect of the present disclosure is provided. It should be understood that in other exemplary embodiments, the cooling airflow 332 may not be provided from the gas turbine engine 310, but may be received from a dedicated airflow supply system 333 having a dedicated air intake 333-1, a dedicated air pressurization system 333-2 (e.g., a fan, an ejector, or a combination thereof), and a dedicated hot air exhaust 333-3.

[0056] Now return to reference Figure 3A It should be appreciated that during certain operations of an aircraft including aircraft system 300, airflow radiator heat exchanger 324 may provide sufficient heat removal to thermal fluid and thermal management system 308 to maintain the temperature of component module 318 within a desired range. For example, under operating conditions where sufficient airflow is available to airflow radiator heat exchanger 324 and the amount of heat generated by component module 318 is relatively low, thermal management system 308 may use airflow radiator heat exchanger 324 as the sole heat sink for thermal management system 308 to maintain the temperature of component module 318 within a desired temperature range.

[0057] However, under certain operating conditions, such as the first operating condition, the airflow available to airflow radiator heat exchanger 324 may be insufficient to meet the amount of heat generated by component module 318, thereby requiring the thermal fluid within thermal fluid loop 320 to supplement or assist with the heat exchanger.

[0058] Thus, the thermal management system 308 includes a fuel radiator heat exchanger 326 that is selectively thermally connected to the thermal fluid loop 320, selectively thermally connected to the fuel tank 312, or both. For example, in the illustrated embodiment, the thermal management system 308 includes an actively controlled switch valve 336 that selectively fluidly connects the fuel radiator heat exchanger 326 to the thermal fluid loop 320. In addition, the fuel system 304 includes an auxiliary loop 338 extending between an inlet 340 and an outlet 342. Each of the inlet 340 and the outlet 342 is in fluid communication with the fuel tank 312 during at least the first operating condition to provide a fuel flow through the auxiliary loop 338. The auxiliary loop 338 is in fluid communication with the fuel radiator heat exchanger 326 for providing fuel to the fuel radiator heat exchanger 326 during at least the first operating condition. Notably, in the illustrated embodiment, the auxiliary loop 338 is fluidly coupled to the fuel tank 312 and fluidly coupled to the fuel radiator heat exchanger 326. Additionally, the fuel system 304 includes an auxiliary fuel pump 344 in fluid communication with the auxiliary loop 338 for providing a fuel flow through the auxiliary loop 338 during at least the first operating condition.

[0059] Still reference Figure 3A It should be understood that during the first operating condition, the fuel radiator heat exchanger 326 is in fluid communication with the hot fluid loop 320, receiving at least a portion (e.g., at least 25%, such as at least 50%, such as at least 75%, such as up to 100%) of the hot fluid flow through the hot fluid loop 320, and is also in fluid communication with the fuel tank 312 via the auxiliary loop 338, receiving the fuel flow from the auxiliary loop 338. Notably, the fuel radiator heat exchanger 326 is in fluid communication with the fuel tank 312 independently of the fuel delivery system 314 during at least the first operating condition, and more specifically, during all operating conditions of the depicted embodiment. In this way, the ability of the fuel radiator heat exchanger 326 to reject heat from the hot fluid to the fuel does not depend on the amount of fuel provided from the fuel tank 312 to, for example, the gas turbine engine 310 through the fuel delivery system 314.

[0060] The first operating condition may be any operating condition in which the mass flow rate of the cooling airflow 332 provided to the airflow radiator heat exchanger 324 is not high enough to remove a sufficient amount of heat to maintain the component module 318 within a desired operating temperature range. For example, in certain exemplary embodiments, the first operating condition may be a takeoff operating condition. During the takeoff operating condition, the component module 318 may be operated in a manner that generates a relatively large amount of heat. For example, in an embodiment in which the component module 318 is a battery, the battery may transmit a large amount of power to an electric motor. In an embodiment in which the component module 318 is an electric motor, the electric motor may be operated to drive one or more shafts of the gas turbine engine 310, etc. In addition, during the takeoff operating condition, the aircraft may not be traveling at a very high speed, so that the mass flow rate of the cooling airflow 332 provided to the airflow radiator heat exchanger 324 may be relatively low. Therefore, operating the thermal management system 308 to provide at least a portion of the thermal fluid to the fuel radiator heat exchanger 326 through the thermal fluid loop 320 to transfer heat from the thermal fluid to the fuel flow from the fuel tank 312 can allow the necessary amount of heat removal to maintain the component module 318 within a desired operating temperature range.

[0061] However, it should be understood that during other operating conditions, supplemental heat rejection as described above with respect to the first operating condition may not be required. For example, during at least the second operating condition, the fuel radiator heat exchanger 326 is thermally disconnected from the hot fluid loop 320, the fuel tank 312, or both. More specifically, in the depicted embodiment, during at least the second operating condition, the fuel radiator heat exchanger 326 is thermally disconnected from the hot fluid loop 320 using an actively controlled switch valve 336 (e.g., the fuel radiator heat exchanger 326 receives less than 10% of the hot fluid passing through the hot fluid loop 320, such as less than 5%, such as less than 1%). As previously described, the fuel radiator heat exchanger 326 is fluidly connected to the fuel tank 312 via an auxiliary loop 338, which, in the illustrated embodiment, is fluidly coupled to the fuel tank 312. However, it is noteworthy that the auxiliary fuel pump 344 may not operate during the second operating condition.

[0062] In certain exemplary embodiments, the second operating condition may be a cruise operating condition. During the cruise operating condition, a relatively high mass flow rate of cooling airflow 332 may be provided to airflow radiator heat exchanger 324 so that a sufficient amount of heat may be removed from the hot fluid through airflow radiator heat exchanger 324 to maintain component module 318 within a desired operating temperature range.

[0063] However, it should be understood that Figure 3A The exemplary embodiments of the present invention are provided by way of example only.

[0064] For example, now refer to Figure 4A and Figure 4B , a schematic diagram of an aircraft system 300 is provided according to another exemplary embodiment of the present disclosure. Figure 4A and Figure 4B The embodiments can each be compared with the above references, for example Figure 3A The embodiments described are constructed substantially similarly. Figure 4A and Figure 4B In the embodiment, the fuel system 304 does not include an auxiliary fuel pump (see Figure 3A In contrast, the thermal management system 308 includes a power flow line 346 ( Figure 4A )、346'( Figure 4B ).

[0065] In particular, refer to Figure 4A , a motive flow line 346 is depicted extending from the fuel tank 312 of the fuel system 304 to the auxiliary loop 338 of the fuel system 304 to induce fuel flow through the auxiliary loop 338 during at least the first operating condition. The motive flow line 346 may include a venturi pump or other similar device to induce fuel flow through the auxiliary loop 338. In this manner, the motive flow line 346 may be designed to provide a pressurized fuel flow to the auxiliary loop 338 when fuel flow through the auxiliary loop 338 is desired.

[0066] In addition, reference Figure 4B , depicted motive flow lines 346 extend from the fuel delivery system 314 to the auxiliary loop 338 to facilitate fuel flow through the auxiliary loop 338 during at least the first operating condition. As will be appreciated, with this configuration, fuel flow from the fuel delivery system 314 can travel to and through the auxiliary loop 338 and then return to the fuel tank 312.

[0067] In addition, now refer to Figure 5 , a schematic diagram of an aircraft system 300 is provided according to yet another exemplary embodiment of the present disclosure. Figure 5 The embodiments can be compared with the above reference Figure 3A The embodiments described are constructed substantially similarly. Figure 5 In the embodiment of the present invention, the thermal management system 308 does not include the auxiliary loop 338. In contrast, in Figure 5 In the exemplary embodiment, fuel radiator heat exchanger 326 of thermal management system 308 of aircraft system 300 is positioned in direct fluid communication with fuel tank 312 of fuel system 304 of aircraft system 300. For example, in certain exemplary embodiments, fuel radiator heat exchanger 326 may be formed integrally with fuel tank 312 or may be formed separately from fuel tank 312 and coupled to fuel tank 312.

[0068] In addition, now refer to Figure 6, a schematic diagram of an aircraft system 300 is provided according to yet another exemplary embodiment of the present disclosure. Figure 6 The embodiments can also be compared with the previous reference Figure 3A The described exemplary embodiments are constructed essentially similarly.

[0069] For example, the exemplary aircraft system 300 includes a thermal management system 308 having a thermal fluid loop 320, a heat source heat exchanger 322 that is thermally connected to the component module 318 and the thermal fluid loop 320, an airflow radiator heat exchanger 324 that is thermally connected to the thermal fluid loop 320 (and is configured to be thermally connected to the cooling airflow 332 during operation of the gas turbine engine 310), and a fuel radiator heat exchanger 326 that is selectively thermally connected to the thermal fluid loop 320, the fuel tank 312, or both.

[0070] However, in the depicted embodiment, the airflow radiator heat exchanger 324 is a first airflow radiator heat exchanger 324-1, and the thermal management system 308 also includes a second airflow radiator heat exchanger 324-2, and so on to an Nth airflow radiator heat exchanger 324-N. The thermal fluid loop 320 also includes a respective first branch 350-1, a second branch 350-2, and so on to an Nth branch 350-N, each branch being for a respective airflow radiator heat exchanger 324-1, 324-2, 324-N. This configuration may allow for one or more cooling airflows 332 (see, e.g., Figure 3A ) to dissipate the desired amount of heat.

[0071] Similarly, fuel radiator heat exchanger 326 is a first fuel radiator heat exchanger 326-1, and fuel tank 312 is a first fuel tank 312-1 of aircraft system 300. Thermal management system 308 also includes a second fuel radiator heat exchanger 326-2, and up to an Nth fuel radiator heat exchanger 326-N. Aircraft system 300 also includes a second fuel tank 312-2, and up to an Nth fuel tank 312-N. In certain exemplary embodiments, first fuel tank 312-1 may be a liquid fuel tank (including, for example, jet fuel, such as Jet A), and second fuel tank 312-2 may be a cryogenic fuel tank (including, for example, cryogenic hydrogen fuel).

[0072] The second fuel radiator heat exchanger 326-2 is selectively thermally connected to the thermal fluid loop 320, the second fuel tank 312-2, or both. In particular, in the depicted embodiment, the thermal management system 308 includes an actively controlled switching valve 336 that selectively fluidly connects the first fuel radiator heat exchanger 326-1 to the thermal fluid loop 320, selectively fluidly connects the second fuel radiator heat exchanger 326-2 to the thermal fluid loop 320, and selectively thermally connects the Nth fuel radiator heat exchanger 326-N to the thermal fluid loop 320.

[0073] Reference now Figure 7 , a method 400 for operating a thermal management system of an aircraft according to an exemplary aspect of the present disclosure is provided. The method 400 may be used to operate the thermal management system of the aircraft as described above with reference to Figures 3A to 6 One or more of the thermal management systems of the aircraft systems described.

[0074] The method 400 includes, at ( 402 ), adding heat to a thermal fluid flow through a main channel of a thermal fluid loop using a thermal source heat exchanger thermally coupled to a component module of an electric drive assembly.

[0075] The method 400 also includes, at ( 404 ), removing heat from the hot fluid flow through the main channel of the hot fluid loop using an airflow radiator heat exchanger thermally coupled to the cooling airflow.

[0076] Still refer to Figure 7 The exemplary method 400 depicted also includes directing at least a portion of the hot fluid flow through the main passage of the hot fluid loop through a supplemental portion of the hot fluid loop at (406), and exchanging heat from the hot fluid directed through the supplemental portion of the hot fluid loop to a fuel flow through an auxiliary loop in fluid communication with a fuel tank of the aircraft at (408). The auxiliary loop is separate from a fuel delivery system fluidly coupled to the fuel tank. In this manner, any fuel passing through the supplemental portion is returned to the main cavity of the tank before being provided through the fuel delivery system, and any fuel provided to the supplemental portion does not travel through the fuel delivery system between the main cavity and the supplemental portion.

[0077] More specifically, at least in the exemplary aspect depicted, method 400 also includes operating the thermal management system at a first operating condition at (410). Operating the thermal management system at the first operating condition at (410) includes the steps of directing at least a portion of the cooling fluid flow through the primary passage through the supplemental portion using an actively controlled switching valve at (406), and exchanging heat from the hot fluid directed through the supplemental portion of the hot fluid loop to the fuel flow through the auxiliary loop at (408).

[0078] Figure 7The method 400 also includes operating the thermal management system at a second operating condition at (412). Operating the thermal management system at the second operating condition at (412) includes thermally isolating the cooling fluid flow through the main passage from the fuel tank at (414).

[0079] Still reference Figure 7 , method 400 also includes operating the thermal management system under a third operating condition at (416). Operating the thermal management system under the third operating condition at (416) includes directing all thermal fluid flow through the main passage of the thermal fluid loop through a supplemental portion of the thermal fluid loop at (418), and exchanging heat from the thermal fluid directed through the supplemental portion of the thermal fluid loop to a fuel flow through an auxiliary loop in fluid communication with a fuel tank of the aircraft at (420). As described above, the auxiliary loop is separate from a fuel delivery system that is fluidly coupled to the fuel tank. In this manner, any fuel that passes through the supplemental portion is returned to the main chamber of the tank before being provided through the fuel delivery system, and any fuel provided to the supplemental portion does not travel through the fuel delivery system between the main chamber and the supplemental portion.

[0080] It should be understood that operating the thermal management system at the third operating condition at (416) may include operating the thermal management system during ground operating conditions (e.g., when the cooling airflow is insufficient to provide meaningful cooling), during a fault condition of the airflow cooling system, or both.

[0081] It should be understood that the exemplary systems and methods described above may be incorporated into an aircraft and / or propulsion system utilizing any suitable gas turbine engine and electric drive assembly. For example, although the above system is described as being incorporated into a propulsion system having a ducted turbofan engine (see, e.g., Figure 2 ), but in other exemplary embodiments, one or more systems or methods of the present disclosure may be incorporated into an aircraft and / or propulsion system utilizing any other suitable gas turbine engine and electric drive assembly. For example, in other exemplary embodiments, one or more systems or methods of the present disclosure may be incorporated into an aircraft and / or propulsion system utilizing a turboprop engine.

[0082] Reference now Figure 8 and Fig. 9 , two gas turbine engines according to exemplary embodiments of the present disclosure are provided, which may be used with one or more systems or methods of the present disclosure. Figure 8 and Fig. 9Each of the gas turbine engines is configured as a turboprop 500. The turboprops 500 each include a propeller 502, a turbine 504, and, for the depicted embodiment, a gearbox 506. The turbine 504 includes a compressor 508, a combustor 510, and a high-pressure turbine 512. The compressor 508 and the high-pressure turbine 512 are connected by a high-pressure shaft 514.

[0083] Turbine 504 of each of turboprop engines 500 also includes a power turbine 516 coupled to a drive shaft 518 rotatably coupled to propeller 502 across gearbox 506 .

[0084] Each of the turboprop engines 500 may also be operable with an electric drive assembly having an electric motor 520 that is operable to add power to the turboprop engine 500 , extract power from the turboprop engine 500 , or both.

[0085] Special References Figure 8 , the motor 520 is rotatably coupled to the turboprop engine 500 via the gearbox 506. With particular reference to Fig. 9 , the motor 520 is rotatably connected to the drive shaft 518.

[0086] exist Figure 8 and Fig. 9 In one or both of the embodiments of the present invention, it should be understood that a thermal management system (such as the one described above with reference to Figures 3A to 6 One or more of the described exemplary thermal management systems) can allow for the inclusion of an electric machine 520 having a desired amount of power to achieve desired efficient operation by ensuring that the electric machine 520 can be maintained within a desired operating temperature range during various operating conditions (e.g., takeoff, cruise, ground idle, etc.).

[0087] Further aspects are provided by the subject matter of the following clauses:

[0088] 1. An aircraft system, comprising: a gas turbine engine; a fuel system having a fuel tank and a fuel delivery system fluidly connecting the fuel tank to the gas turbine engine; an electric drive assembly, the electric drive assembly including a component module; and a thermal management system, the thermal management system including: a thermal fluid loop; a heat source heat exchanger, the heat source heat exchanger being thermally connected to the component module and the thermal fluid loop; an airflow radiator heat exchanger, the airflow radiator heat exchanger being thermally connected to the thermal fluid loop and further configured to be thermally connected to a cooling airflow during operation of the gas turbine engine; and a fuel radiator heat exchanger, the fuel radiator heat exchanger being selectively thermally connected to the thermal fluid loop, the fuel tank, or both, the fuel radiator heat exchanger being fluidly connected to the fuel tank independently of the fuel delivery system during at least a first operating condition.

[0089] An aircraft system as recited in any preceding clause, wherein the thermal management system includes an actively controlled switching valve that selectively fluidly connects the fuel radiator heat exchanger to the thermal fluid loop.

[0090] An aircraft system as recited in any preceding clause, wherein the fuel system further comprises an auxiliary loop extending between an inlet and an outlet, wherein the inlet and the outlet are each in fluid communication with the fuel tank during at least the first operating condition.

[0091] An aircraft system as claimed in any preceding clause, wherein the fuel system further comprises an auxiliary fuel pump in fluid communication with the auxiliary loop for providing a fuel flow through the auxiliary loop during at least the first operating condition.

[0092] An aircraft system as recited in any preceding clause, wherein the thermal management system further comprises a power flow line extending from the fuel tank to the auxiliary loop to facilitate fuel flow through the auxiliary loop during at least the first operating condition.

[0093] An aircraft system as recited in any preceding clause, wherein the component modules include batteries, fuel cells, electric machines, power electronics, or combinations thereof.

[0094] An aircraft system as claimed in any preceding clause, wherein the thermal management system comprises a TMS pump in fluid communication with the thermal fluid loop for providing a flow of thermal fluid through the thermal fluid loop.

[0095] An aircraft system as described in any preceding clause, wherein the fuel tank is a liquid fuel tank.

[0096] An aircraft system as described in any preceding clause, wherein the fuel tank is a cryogenic fuel tank.

[0097] An aircraft system as described in any preceding clause, wherein the first operating condition is a takeoff operating condition.

[0098] An aircraft system as recited in any preceding clause, wherein the fuel radiator heat exchanger is thermally disconnected from the thermal fluid loop, the fuel tank, or both during at least a second operating condition.

[0099] An aircraft system as described in any preceding clause, wherein the second operating condition is a cruise operating condition.

[0100] An aircraft system according to any preceding clause, wherein the fuel tank is a first fuel tank, wherein the fuel radiator heat exchanger is a first fuel radiator heat exchanger, wherein the fuel system further includes a second fuel tank, and wherein the thermal management system further includes a second fuel radiator heat exchanger, the second fuel radiator heat exchanger selectively thermally connected to the hot fluid loop, the second fuel tank, or both.

[0101] An aircraft system according to any preceding clause, wherein the thermal management system includes an actively controlled switching valve that selectively connects the first fuel radiator heat exchanger fluid to the thermal fluid loop and selectively connects the second fuel radiator heat exchanger fluid to the thermal fluid loop.

[0102] An aircraft system as recited in any preceding clause, wherein the fuel radiator heat exchanger is integrated into the fuel tank.

[0103] An aircraft system as recited in any preceding clause, wherein the gas turbine engine is a turboprop engine, and wherein the component module of the electric drive assembly is an electric machine capable of rotating with the turboprop engine.

[0104] An aircraft system as described in any preceding clause, wherein the thermal management system further comprises a power flow line extending from the fuel tank or from the fuel delivery system to the auxiliary loop to promote fuel flow through the auxiliary loop during at least the first operating condition.

[0105] A method of operating a thermal management system for an aircraft, the method comprising: adding heat to a thermal fluid flow through a main channel of a thermal fluid loop by utilizing a heat source heat exchanger thermally coupled to a component module of an electric drive assembly; removing heat from the thermal fluid flow through the main channel of the thermal fluid loop by utilizing an airflow radiator heat exchanger thermally coupled to a cooling airflow; directing at least a portion of the thermal fluid flow through the main channel of the thermal fluid loop through a supplemental portion of the thermal fluid loop; and exchanging heat from the thermal fluid directed through the supplemental portion of the thermal fluid loop to a fuel flow through an auxiliary loop fluidly coupled to a fuel tank of the aircraft, wherein the auxiliary loop is separate from a fuel delivery system fluidly coupled to the fuel tank.

[0106] The method according to any of the preceding clauses further includes: operating the thermal management system under a first operating condition, and wherein operating the thermal management system under the first operating condition includes the following steps: using the actively controlled switching valve to direct at least a portion of the cooling fluid flow through the main channel through the supplementary portion; and exchanging heat from the hot fluid directed through the supplementary portion of the hot fluid loop to the fuel flow through the auxiliary loop.

[0107] A method as in any preceding clause, wherein the first operating condition is a take-off operating condition.

[0108] The method of any preceding clause, further comprising operating the thermal management system under a second operating condition, and wherein operating the thermal management system under the second operating condition comprises thermally isolating the cooling fluid flow through the primary passage from the fuel tank.

[0109] The method according to any preceding clause further comprises: operating the thermal management system under a third operating condition, and wherein operating the thermal management system under the third operating condition comprises directing all of the thermal fluid flow through the main channel of the thermal fluid loop through the supplementary portion of the thermal fluid loop; and exchanging heat from the thermal fluid directed through the supplementary portion of the thermal fluid loop to the fuel flow through the auxiliary loop fluidly connected to the fuel tank of the aircraft.

[0110] A method as in any preceding clause, wherein the third operating condition is a ground operating condition, a failure condition of an air flow cooling system, or both.

[0111] A method as in any preceding clause, wherein the component modules include batteries, fuel cells, electric machines, power electronics, or combinations thereof.

[0112] 1. An aircraft system, comprising: a gas turbine engine; a fuel system, the fuel system having a fuel tank, a fuel delivery system fluidly connecting the fuel tank to the gas turbine engine, and an auxiliary loop extending between an inlet and an outlet, the inlet and the outlet each being in fluid communication with the fuel tank; an electric drive assembly, the electric drive assembly including a component module; and a thermal management system, the thermal management system comprising: a thermal fluid loop; a heat source heat exchanger, the heat source heat exchanger being in thermal communication with the component module and the thermal fluid loop; an airflow radiator heat exchanger, the airflow radiator heat exchanger being in thermal communication with the thermal fluid loop and further configured to be in thermal communication with a cooling airflow during operation of the gas turbine engine; and a fuel radiator heat exchanger, the fuel radiator heat exchanger being in thermal communication with the thermal fluid loop and the auxiliary loop of the fuel system during at least a first operating condition.

[0113] An aircraft as recited in any preceding clause, wherein the thermal management system further comprises a power flow line extending from the fuel delivery system to the auxiliary loop to facilitate fuel flow through the auxiliary loop during at least the first operating condition.

[0114] An aircraft as claimed in any preceding clause, wherein the thermal management system further comprises a power flow line extending from the fuel tank or from the fuel delivery system to the auxiliary loop to promote fuel flow through the auxiliary loop during at least the first operating condition.

[0115] This written description uses examples to disclose the present disclosure, including the best mode, and also enables any person skilled in the art to practice the present disclosure, including making and using any device or system and performing any combined methods. The patent scope of the present disclosure is defined by the claims, and may include other examples that occur to those skilled in the art. If these other examples include structural elements that do not differ from the literal language of the claims, or if they include equivalent structural elements that do not differ substantially from the literal language of the claims, these other examples are intended to fall within the scope of the claims.

Claims

1. An aircraft system, characterized in that: include: Gas turbine engines; a fuel system having a fuel tank and a fuel delivery system fluidly connecting the fuel tank to the gas turbine engine; An electric drive assembly, the electric drive assembly comprising a component module; as well as A thermal management system, the thermal management system comprising: Thermal fluid loop; a heat source heat exchanger in thermal communication with the component module and the thermal fluid loop; an airflow radiator heat exchanger in thermal communication with the thermal fluid loop and further configured to be in thermal communication with a cooling airflow during operation of the gas turbine engine; and A fuel radiator heat exchanger is in selective thermal communication with the hot fluid loop, the fuel tank, or both, the fuel radiator heat exchanger being in fluid communication with the fuel tank independent of the fuel delivery system during at least a first operating condition.

2. The aircraft system according to claim 1, characterized in that: in, The thermal management system includes an actively controlled on-off valve that selectively fluidly connects the fuel radiator heat exchanger to the thermal fluid loop.

3. The aircraft system according to claim 1, characterized in that: in, The fuel system further includes an auxiliary loop extending between an inlet and an outlet, wherein the inlet and the outlet are each in fluid communication with the fuel tank during at least the first operating condition.

4. The aircraft system according to claim 3, characterized in that: in, The fuel system further includes an auxiliary fuel pump in fluid communication with the auxiliary loop for providing fuel flow through the auxiliary loop during at least the first operating condition.

5. The aircraft system according to claim 3, characterized in that: in, The thermal management system further includes a motive flow line extending from the fuel tank or from the fuel delivery system to the auxiliary loop to facilitate fuel flow through the auxiliary loop during at least the first operating condition.

6. The aircraft system according to claim 1, characterized in that: in, The component modules include batteries, fuel cells, motors, power electronics, or a combination thereof.

7. The aircraft system according to claim 1, characterized in that: in, The thermal management system includes a thermal management system (TMS) pump in fluid communication with the thermal fluid loop for providing a thermal fluid flow through the thermal fluid loop.

8. The aircraft system according to claim 1, characterized in that: in, The fuel tank is a liquid fuel tank.

9. The aircraft system according to claim 1, characterized in that: in, The first operating condition is a takeoff operating condition.

10. The aircraft system according to claim 1, characterized in that: in, The fuel radiator heat exchanger is thermally disconnected from the thermal fluid loop, the fuel tank, or both during at least a second operating condition.

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