High Temperature Fuel Thermal Transfer Bus Architecture

Abstract

A gas turbine engine assembly includes a gas turbine engine having a combustion section, a fuel delivery system, and a thermal energy management system. The fuel delivery assembly supplies fuel to the combustion section of the gas turbine engine. The thermal energy management system includes a heat transfer bus, a heat source heat exchanger, and a radiator heat exchanger. The heat transfer bus has a portion of the fuel configured to flow therethrough. The fuel is provided as a heat exchange fluid of the thermal energy management system. The heat source heat exchanger is in thermal communication with the fuel flow through the transfer bus. The radiator heat exchanger is in thermal communication with the fuel flow through the transfer bus.

Description

Technical Field

[0001] Generally, the present disclosure relates to thermal energy management in an engine. In particular, the present disclosure relates to a heat transfer bus in a gas turbine engine. Background Art

[0002] A gas turbine engine generally includes a turbine and a rotor assembly. Gas turbine engines (such as turbofan engines) are used for aircraft propulsion. In the case of a turbofan engine, the rotor assembly may be configured as a fan assembly.

[0003] Various accessory systems are included to ensure that the fan and / or core operate as needed. For example, a main lubrication system provides lubrication for bearings and gear meshes within, for example, a compressor section, a turbine section, and a power gearbox (if provided). In addition to the lubricating properties provided to such components, the main lubrication system can also be used to remove heat from such components so that such components operate within a desired temperature range.

[0004] Other accessory systems of a gas turbine engine (such as an environmental control system) also require heat removal during operation. Accordingly, a gas turbine engine typically includes a plurality of heat exchangers, each dedicated to a separate accessory system of the gas turbine engine.

[0005] In existing thermal energy management systems, a heat transfer bus is used to manage the transfer of thermal energy between various heat exchangers. However, the heat exchange medium used in existing heat transfer buses adds an increased fluid volume to the engine. In other cases, existing heat exchange fluids can include relatively low pressure and temperature thresholds to prevent these media from operating at the higher temperatures and pressures that the engine may require. Summary of the Invention

[0006] Aspects and advantages of the present disclosure will be set forth in part in the following description, or may be apparent from the description, or may be learned by practice of the present disclosure.

[0007] In one exemplary embodiment of the present disclosure, a gas turbine engine assembly includes a gas turbine engine having a combustion section, a fuel delivery system, and a thermal energy management system. The fuel delivery assembly provides fuel to the combustion section of the gas turbine engine. The thermal energy management system includes a heat transfer bus, a heat source heat exchanger, and a radiator heat exchanger. The heat transfer bus has a portion of the fuel configured to flow therethrough. The fuel is provided as the heat exchange fluid of the thermal energy management system. The heat source heat exchanger is in thermal communication with the fuel flow through the transfer bus. The radiator heat exchanger is in thermal communication with the fuel flow through the transfer bus.

[0008] In an exemplary embodiment of the present disclosure, a thermal energy management system for a gas turbine engine assembly includes a heat transfer bus, a heat source heat exchanger, and a fuel-fuel heat exchanger. The heat transfer bus is configured to make a portion of the fuel flow flowing through the heat transfer bus serve as a heat exchange fluid of the thermal energy management system. The heat source heat exchanger is in thermal communication with the fuel flow passing through the transfer bus. The fuel-fuel heat exchanger is configured to thermally connect the fuel flow passing through the heat transfer bus to the fuel flow passing through a fuel delivery assembly of the gas turbine engine.

[0009] In an exemplary embodiment of the present disclosure, a gas turbine engine includes a fan section, a compressor section, a combustion section, a turbine section, an exhaust section, a fuel delivery system, and a thermal management system. The fan section, the compressor section, the combustion section, the turbine section, and the exhaust section are arranged in a serial flow order and together define a core air flow path. The fuel delivery system supplies a fuel flow to the combustion section. The thermal management system includes a first heat exchanger, a second heat exchanger, a third heat exchanger, a first radiator exchanger, and a heat transfer bus. The first heat exchanger is positioned to be in thermal communication with the flow passing through the core air flow path within or downstream of the turbine section, the exhaust section, or both. The second heat exchanger is in thermal communication with the fuel delivery system for transferring heat to the fuel flow supplied to the combustion section. The third heat exchanger is connected to and in fluid communication with the heat transfer bus at a location upstream of the first heat exchanger. The first radiator is connected to and in fluid communication with the heat transfer bus at a location downstream of the second heat exchanger. The heat transfer bus has a heat exchange fluid flowing through it. The first heat exchanger, the second heat exchanger, the third heat exchanger, and the first radiator are each fluidly coupled to the heat transfer bus. The heat exchange fluid includes deoxygenated fuel such that the combustion of the deoxygenated fuel provides the driving force of the gas turbine engine.

[0010] These and other features, aspects, and advantages of the present disclosure will be better understood with reference to the following description and the appended claims. The accompanying drawings, which are incorporated in and constitute a part of this specification, illustrate embodiments of the present disclosure and, together with the description, serve to explain the principles of the present disclosure. BRIEF DESCRIPTION OF THE DRAWINGS

[0011] A complete and enabling description of the present disclosure for a person of ordinary skill in the art is set forth in the specification, including its best mode, in which:

[0012] Figure 1 is a cross-sectional view of a gas turbine engine in accordance with an exemplary aspect of the present disclosure.

[0013] Figure 2 is a simplified schematic diagram of a first thermal management system in accordance with an exemplary embodiment of the present disclosure.

[0014] Figure 3 is a simplified schematic diagram of a second thermal management system according to an exemplary embodiment of the present disclosure. Detailed Description

[0015] Reference will now be made in detail to the present embodiments of the disclosure, one or more examples of which are illustrated in the accompanying drawings. The detailed description uses numerical and letter designations to refer to features in the drawings. The same or similar designations in the drawings and description have been used to refer to the same or similar parts of the disclosure.

[0016] The term "exemplary" is used herein to mean "serving as an example, instance, or illustration". Any embodiment described herein as "exemplary" is not necessarily to be construed as superior or advantageous to other embodiments. Further, unless specifically stated otherwise, all embodiments described herein should be considered exemplary.

[0017] As used herein, the terms "first", "second", and "third" may be used interchangeably to distinguish one component from another and are not intended to denote the position or importance of the individual components.

[0018] The terms "forward" and "backward" refer to relative positions within a gas turbine engine or vehicle and refer to the normal operating attitude of the gas turbine engine or vehicle. For example, for a gas turbine engine, forward refers to a position closer to the engine inlet, and backward refers to a position closer to the engine nozzle or exhaust.

[0019] The terms "upstream" and "downstream" refer to the relative direction with respect to the fluid flow in a fluid path. For example, "upstream" refers to the direction from which the fluid flows, and "downstream" refers to the direction towards which the fluid flows.

[0020] Unless otherwise specified herein, the terms "coupled", "fixed", "attached to", etc. refer to direct coupling, fixing, or attachment, as well as indirect coupling, fixing, or attachment through one or more intermediate components or features.

[0021] Unless the context clearly dictates otherwise, the singular forms "a", "an", and "the" include plural referents.

[0022] As used throughout this specification and the claims, approximating language is used to modify any quantitative representation that can vary without resulting in a change in the basic function associated therewith. Thus, values modified by one or more terms such as "about," "approximately," and "substantially" are not limited to the specified exact values. In at least some instances, the approximating language can 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, the approximating language can refer to within a margin of 1%, 2%, 4%, 10%, 15%, or 20%. These approximating margins can be applied to a single value, one or both of the endpoints defining a numerical range, and / or the margins of the range between the endpoints.

[0023] Herein and throughout the specification and claims, range limitations are combined and interchanged, and unless the context or language indicates otherwise, such ranges are identified and include all subranges subsumed therein. For example, all ranges disclosed herein include the endpoints, and the endpoints can be combined independently of one another.

[0024] The disclosure presented herein describes a heat transfer bus architecture that utilizes the combustible fuel of an engine as a heat exchange medium (e.g., a fluid). The deoxygenated fuel is recycled through a heat transfer bus in a loop that is pressurized and continuously charged by a reference pressure set by the fuel system of the engine propulsion system. Utilizing fuel as a heat exchange fluid does not require an additional working fluid such as supercritical carbon dioxide. Additionally, certain embodiments of the disclosed fuel-based heat transfer bus and associated components include a pressure-controlled recirculation loop within the engine fuel system that provides the reference pressure, thereby eliminating the need for discrete accumulators or reservoirs to accommodate system expansion, contraction, and / or charging.

[0025] Furthermore, the disclosed thermal management system allows for leak isolation along the heat transfer bus. For example, in the case of a leak associated with one of the heat exchangers on the heat transfer bus circuit, the isolated heat transfer bus circuit also allows for leak detection and isolation, potentially eliminating the need for in-flight shutdowns. The proposed thermal management system also discloses a bus control system that is capable of purging and priming the heat transfer bus. The bus control system includes a commissioning valve that is set to control the recycling of fuel into and out of the heat transfer bus and to purge air out of the heat transfer bus. This purging and priming function allows all necessary surfaces of the heat exchangers along the heat transfer bus to be properly wetted to provide the desired performance, thereby preventing excessive air from entering the system and purging air from the system.

[0026] Now referring to the drawings, in which like numerals represent like elements throughout the drawings, Figure 1 is a schematic cross-sectional view of a gas turbine engine according to an exemplary embodiment of the present disclosure. More specifically, for Figure 1In an embodiment, the gas turbine engine is a high bypass turbofan jet engine 10, referred to herein as "turbofan engine 10". As Figure 1 shown, the turbofan engine 10 defines an axial direction A (extending parallel to a longitudinal centerline 12 provided for reference) and a radial direction R. Generally, the turbofan engine 10 includes a fan section 14 and a core turbine engine 16 disposed downstream of the fan section 14.

[0027] The exemplary core turbine engine 16 shown generally includes a substantially tubular outer casing 18 defining an annular inlet 20. The outer casing 18 surrounds in serial flow relationship: a compressor section including a booster or low pressure (LP) compressor 22 and a high pressure (HP) compressor 24; a turbine section including a high pressure (HP) turbine 28 and a low pressure (LP) turbine 30; and a combustion section 26. A high pressure (HP) shaft or spool 34 drivingly connects the HP turbine 28 to the HP compressor 24. A low pressure (LP) shaft or spool 36 drivingly connects the LP turbine 30 to the LP compressor 22.

[0028] For the depicted embodiment, the fan section 14 includes a variable pitch fan 38 having a plurality of fan blades 40 coupled to a disk 42 in a spaced apart manner. As shown, the fan blades 40 generally extend radially outwardly from the disk 42 in the radial direction R. Each fan blade 40 is rotatable relative to the disk 42 about a pitch axis P by being operatively coupled to a suitable actuating member 44, and the actuating member 44 is configured to commonly and uniformly change the pitch of the fan blades 40. The fan blades 40, the disk 42, and the actuating member 44 together are rotatable about the longitudinal axis 12 by the LP shaft 36 through a cross power gearbox 46. The power gearbox 46 includes a plurality of gears for reducing the rotational speed of the LP shaft 36 to a more efficient rotational fan speed.

[0029] Still referring to Figure 1 the exemplary embodiment, the disk 42 is covered by a rotatable front hub 48, and the aerodynamic profile of the front hub 48 facilitates airflow through the plurality of fan blades 40. Additionally, the fan section 14 includes an annular fan casing or nacelle 50 that circumferentially surrounds the fan 38 and / or at least a portion of the core turbine engine 16. It should be understood that in some embodiments, the nacelle 50 may be configured to be supported relative to the core turbine engine 16 by a plurality of circumferentially spaced outlet guide vanes 52. Further, a downstream section 54 of the nacelle 50 may extend over an outer portion of the core turbine engine 16 to define a bypass airflow passage 56 therebetween.

[0030] During operation of the turbofan engine 10, a volume of air 58 enters the turbofan engine 10 through the nacelle 50 and / or the associated inlet 60 of the fan section 14. As the volume of air 58 passes through the fan blades 40, a first portion of the air 58, as indicated by arrow 62, is directed or routed into the bypass airflow passage 56, and a second portion of the air 58, as indicated by arrow 64, is directed or routed into the LP compressor 22. The ratio between the first portion of air 62 and the second portion of air 64 is commonly referred to as the bypass ratio. When the second portion of air 64 is routed through the high-pressure (HP) compressor 24 and into the combustion section 26, the pressure of the second portion of air 64 then increases, and in the combustion section 26 the second portion of air 64 is mixed with fuel and burned to provide combustion gases 66. Subsequently, the combustion gases 66 are routed through the HP turbine 28 and the LP turbine 30, where a portion of the thermal energy and / or kinetic energy from the combustion gases 66 is extracted.

[0031] The combustion gases 66 are then routed through the combustion section 26 of the core turbofan engine 16 to provide propulsion thrust. At the same time, as the first portion of air 62 is routed through the bypass airflow passage 56 before being exhausted from the fan nozzle exhaust section 76 of the turbofan engine 10, the pressure of the first portion of air 62 significantly increases, also providing propulsion thrust.

[0032] In addition, as schematically depicted, the turbofan engine 10 also includes various auxiliary systems to assist in the operation of the turbofan engine 10 and / or an aircraft including the turbofan engine 10. For example, the turbofan engine 10 also includes a main lubrication system 78 that is configured to provide lubrication to various bearings and gear meshes in, for example, the compressor section (including the LP compressor 22 and the HP compressor 24), the turbine section (including the HP turbine 28 and the LP turbine 30), the HP spool 34, the LP spool 36, and the power gearbox 46. The lubricant provided by the main lubrication system 78 increases the service life of these components and removes a certain amount of heat from these components. In addition, the turbofan engine 10 includes a compressor cooling air ("CCA") system 80 for providing air from one or both of the HP compressor 24 or the LP compressor 22 to one or both of the HP turbine 28 or the LP turbine 30. In addition, the turbofan engine 10 includes an active thermal clearance control ("ACC") system 82 for cooling the housing of the turbine section to maintain the clearance between various turbine rotor blades and the turbine housing within a desired range under various engine operating conditions. In addition, the turbofan engine 10 includes a generator lubrication system 84 for providing lubrication to an electric generator. In one example, the electric generator provides power to a starting motor of the turbofan engine 10, various other electrical components of the turbofan engine 10, and / or an aircraft including the turbofan engine 10.

[0033] It is also schematically depicted that the turbofan engine 10 drives or enables various other accessory systems of an aircraft including the turbofan engine 10. For example, the turbofan engine 10 supplies compressed air from the compressor section to an environmental control system (“ECS”) 86. In one example, the ECS 86 provides an air supply to the aircraft's cockpit for pressurization and thermal control. Air can be supplied from the turbofan engine 10 to an electronics cooling system 88 to maintain the temperature of the turbofan engine 10 and / or certain electronic components of the aircraft within a desired range. Additionally, a surface cooler 90 can be included in the fan section 14. In this example, the surface cooler 90 is a radiator exchanger. The surface cooler 90 is disposed along the surface of the fan section 14 (e.g., the annular fan housing or outer nacelle 50 circumferentially surrounding the fan 38). In this example, the surface cooler 90 is exposed to a first portion of the air 62 flowing through the turbofan engine 10 (e.g., in thermal communication therewith). The surface cooler 90 operates by transferring thermal energy from the surface cooler 90 directly via conduction (and / or convection) or indirectly via the sidewall of the fan section 14 to the first portion of the air 62. As will be discussed with respect to Figure 2 which will be further discussed, the surface cooler 90 is connected to a thermal management system (e.g., see Figure 2 the thermal management system 100 in

[0034] However, it should be understood that Figure 1 the turbofan engine 10 depicted in

[0035] Figure 2 is only an example, and in other exemplary embodiments, aspects of the present disclosure can additionally or alternatively be applied to any other suitable gas turbine engine. For example, in other exemplary embodiments, the turbofan engine 10 can alternatively be any other suitable aero gas turbine engine, such as a turbojet engine, a turboshaft engine, a turboprop engine, etc. Additionally, in other exemplary embodiments, the turbofan engine 10 can include or be operably connected to any other suitable accessory system. Additionally or alternatively, the turbofan engine 10 can not include or can not be operably connected to one or more of the above-mentioned accessory systems.is a simplified schematic diagram of a thermal management system 100 and shows a heat transfer bus 102, a pump 104, a heat source exchanger 106 (including a first heat source exchanger 106A and a second heat source exchanger 106B), a radiator exchanger 108 (including a first radiator exchanger 108A and a second radiator exchanger 108B), a bypass line 110 (including a first bypass line 110A, a second bypass line 110B, a third bypass line 110C, and a fourth bypass line 110D), upstream junctions 112A - 112D, downstream junctions 114A - 114D, check valves 116A - 116D, a flow control valve 118, a leak isolation system 120, a fuel delivery assembly 122 (including a BP 124, an MFP 126, a fuel metering unit 130, and a restrictor 132), and a fuel delivery line 134. Figure 2 Also shown is a simplified schematic representation of a combustion section 26 of a turbofan engine 10.

[0036] As shown, the thermal management system 100 is a thermal energy management system and generally includes a heat transfer bus 102. The heat transfer bus 102 includes an intermediate heat exchange fluid flowing therethrough and can be formed by one or more suitable fluid conduits. In this example, the heat exchange fluid is the fuel of the aircraft. For example, the heat exchange fluid can be a deoxygenated fuel, whereby the combustion of the deoxygenated fuel provides the driving force for the turbofan engine 10 via the combustion section 26. In certain exemplary embodiments, the deoxygenated fuel can have an oxygen content of less than or equal to about 5 parts per million to 1 part per million.

[0037] The thermal management system 100 also includes a pump 104 in fluid communication with the heat exchange fluid in the heat transfer bus 102 for generating a heat exchange fluid flow in and / or through the heat transfer bus 102. As Figure 2 shown, the pump 104 generates a heat exchange fluid flow through the heat transfer bus 102. In one example, the pump 104 can be a centrifugal pump, a rotary pump including an impeller, or alternatively can be any other suitable fluid pump. In one example, the pump 104 can be mechanically coupled to and / or driven by an accessory gearbox of the turbofan engine 10. In another example, the pump 104 can be powered by an electric motor, or alternatively can be mechanically in communication with and powered by the HP shaft 34 or the LP shaft 36 of the turbofan engine 10.

[0038] The thermal management system 100 also includes a plurality of heat source exchangers 106 (including a first heat source exchanger 106A and a second heat source exchanger 106B), each heat source exchanger being in thermal communication with a heat exchange fluid in the heat transfer bus 102. Each of the plurality of heat source exchangers 106 is configured to transfer or absorb thermal energy from one or more accessory systems of the turbofan engine 10 (or operably connected to the turbofan engine 10) to the heat exchange fluid in the heat transfer bus 102.

[0039] For example, the first heat source exchanger 106A may be configured as an "air - heat exchange fluid" heat exchanger, which is integrated into the turbofan engine 10 or one or both of the aircraft including the turbofan engine 10. During operation, the first heat source exchanger 106A transfers thermal energy to a portion of the heat exchange fluid in the heat transfer bus 102 by passing a quantity of air (e.g., engine bleed air or ECS bleed air) through the first heat source exchanger 106A. In one example, the first heat source exchanger 106A may be a bus cooling air cooler, which transfers heat from an air stream (e.g., the air stream discharged from the HP compressor 24) to the fuel flowing through the heat transfer bus 102. From there, the cooled air can then be used for cooling (e.g., for cooling the HP turbine 28).

[0040] Similarly, the second heat source exchanger 106B can generally be an "air - heat exchange fluid" heat exchanger, which passes air (e.g., a waste heat air stream) through the heat exchange fluid to transfer thermal energy to the heat exchange fluid. In one example, the second heat source exchanger 106B may be a waste heat recovery heat exchanger, which extracts thermal energy (e.g., from the exhaust of the turbofan engine 10) and supplies the thermal energy to the fuel flowing through the heat transfer bus 102.

[0041] In one example, the first radiator exchanger 108A may be configured as a "fluid - heat exchange fluid" radiator, where thermal energy from the heat exchange fluid in the heat transfer bus 102 is transferred to a liquid fuel stream for the turbofan engine 10. In particular, the first radiator exchanger 108A is configured as a "fuel - fuel" radiator, where thermal energy from the fuel in the heat transfer bus 102 is transferred to the fuel stream flowing through the fuel delivery line 134, thereby increasing the temperature of the fuel to be combusted (e.g., by the combustion section 26).

[0042] In other exemplary embodiments, the plurality of heat source exchangers 106 may include two or more of the following: a main lubrication system heat exchanger for transferring heat from the main lubrication system 78; a CCA system heat source exchanger for transferring heat from the CCA system 80; an ACC system heat source exchanger for transferring heat from the ACC system 82; a generator lubrication system heat source exchanger for transferring heat from the generator lubrication system 84; an ECS heat exchanger for transferring heat from the ECS 86; and an electronic cooling system heat exchanger for transferring heat from the electronic cooling system 88. Thus, according to Figure 2 the exemplary embodiment of the thermal management system 100 transfers thermal energy from various independent systems to the heat exchange fluid in the heat transfer bus 102 for removal.

[0043] For the depicted embodiment, there are two heat source exchangers 106 (e.g., a first heat source exchanger 106A and a second heat source exchanger 106B), where each heat source exchanger 106 is arranged in a serial flow arrangement along the heat transfer bus 102. In other examples, a different number of heat source exchangers 106 may be included, and one or more of the heat source exchangers 106 may be arranged in a parallel flow arrangement along the heat transfer bus 102. For example, in other embodiments, there may be one heat source exchanger 106 in thermal communication with the heat exchange fluid in the heat transfer bus 102, or alternatively, there may be at least three heat source exchangers 106, at least four heat source exchangers 106, or at least five heat source exchangers 106 in thermal communication with the heat exchange fluid in the heat transfer bus 102.

[0044] Figure 2 the thermal management system 100 also includes a first radiator exchanger 108A and a second radiator exchanger 108B that are permanently or selectively in thermal communication with the heat exchange fluid in the heat transfer bus 102. The first radiator exchanger is located downstream of the plurality of heat source exchangers 106. The second radiator exchanger 108B is located downstream of the first radiator exchanger 108A and is configured to transfer heat from the heat exchange fluid in the heat transfer bus 102 to, for example, the atmosphere, fuel, fan flow, etc. In this example, the second radiator exchanger 108B is the same component as Figure 1 the surface cooler 90 shown (e.g., a fan air surface cooler exposed to air flowing through the fan section 14).

[0045] The heat exchangers 106A-106B and the radiator exchangers 108A-108B are described as being selectively in thermal communication with the heat exchange fluid in the heat transfer bus 102. More specifically, the thermal management system 100 includes bypass lines 110A, 110B, 110C, and 110D for selectively bypassing each of the first heat source exchanger 106A, the second heat source exchanger 106B, the first radiator exchanger 108A, and the second radiator exchanger 108B, respectively. In one example, the bypass lines 110A, 110B, 110C, and 110D can be used to isolate the volumes of the first heat source exchanger 106A, the second heat source exchanger 106B, the first radiator exchanger 108A, and the second radiator exchanger 108B. In this case, by utilizing the now isolated volumes of the first heat source exchanger 106A, the second heat source exchanger 106B, the first radiator exchanger 108A, and the second radiator exchanger 108B, leak identification can be achieved by monitoring the pressure of a fixed volume in any one of the first heat source exchanger 106A, the second heat source exchanger 106B, the first radiator exchanger 108A, and the second radiator exchanger 108B.

[0046] Each bypass line 110 extends between a respective upstream junction 112 and a respective downstream junction 114. For example, the upstream junction 112 is located immediately upstream of the respective heat source exchanger 106 or radiator exchanger 108, while the downstream junction 114 is located immediately downstream of the respective heat source exchanger 106 or radiator exchanger 108. Each of the bypass lines 110A and 110B meets the heat transfer bus 102 at their respective upstream junctions 112A and 112B via three-way valves 136A and 136B, respectively. Each of the bypass lines 110C and 110D meets the heat transfer bus 102 at their respective downstream junctions 114A and 114B via three-way valves 136C and 136D, respectively.

[0047] For example, the three-way valve 136A includes an inlet fluidly connected to the heat transfer bus 102, a first outlet fluidly connected to a portion of the heat transfer bus 102 extending to the first heat source exchanger 106A, and a second outlet fluidly connected to the bypass line 110A, and the same is true for the three-way valve 136B. For the three-way valve 136C, the outlet of the three-way valve 136C is fluidly connected to the heat transfer bus 102, the first inlet is fluidly connected to a portion of the heat transfer bus 102 extending from the first radiator exchanger 108A, and the second inlet is fluidly connected to the bypass line 110C, and the same is true for the three-way valve 136D.

[0048] In this example, the three-way valves 136A - 136D ("three-way valve 136") can each be variable throughput three-way valves such that the three-way valve 136 can vary the throughput from the inlet to the first outlet and / or the second outlet. For example, the three-way valve 136 can be configured to provide anywhere from zero percent (0%) to one hundred percent (100%) of the heat exchange fluid from the inlet to the first outlet, and similarly, the three-way valve 136 can be configured to provide anywhere from zero percent (0%) to one hundred percent (100%) of the heat exchange fluid from the inlet to the second outlet.

[0049] In addition, each of the bypass lines 110A - 110D also meets or joins the heat transfer bus 102 at a respective one of the downstream junctions 114A - 114D. Between each of the heat source exchangers 106 or radiator exchangers 108 and the downstream junction 114, the heat transfer bus 102 includes check valves 116 (see, for example, check valves 116A - 116D) for ensuring the proper flow direction of the heat exchange fluid. More specifically, each of the check valves 116 prevents the flow of heat exchange fluid from the downstream junction 114 towards the respective heat source exchanger 106 or radiator exchanger 108.

[0050] The leak isolation system 120 is a system for detecting and isolating fuel leaks in the thermal management system 100. The fuel delivery system 122 includes a BP 124, a main fuel pump 126, an actuation device 128, a fuel metering unit 130, and a restrictor 132. The BP 124 is a boost pump for providing initial pressurization of the fuel prior to the gear stage. The main fuel pump 126 is a pump for pressurizing and distributing fuel to and from the fuel delivery system 122. The actuation device 128 is a component configured to use high-pressure fuel as a power fluid for moving variable geometry components (such as one or more variable geometry components of an engine (e.g., guide vanes)). The fuel metering unit 130 is a device for regulating or metering the fluid flow therethrough. In this example, the fuel metering unit 130 regulates the fuel flow therethrough. The restrictor 132 is a fixed or variable orifice for restricting or slowing the fuel flow therethrough.

[0051] In this example, the fuel delivery system 122 functions by controlling and providing a pressurized fuel flow to and from the leak isolation system 120 and the first radiator exchanger 108A. In one example, the fuel delivery system 122 may be referred to as a pressure maintenance system. The fuel delivery line 134 is a tube or conduit configured to transmit fluid (e.g., fuel) therethrough. The fuel delivery line 134 is fluidly connected to the fuel metering unit 130 of the fuel delivery system 122, extends to and through the first radiator exchanger 108A, and continues to be fluidly connected to the combustion section 26 of the turbofan engine 10. The three-way valves 136A - 136D (collectively referred to as the three-way valve 136) are each variable throughput three-way valves such that each of the three-way valves 136 can vary the throughput from an inlet to a first outlet and / or a second outlet.

[0052] Since fuel is used as the heat exchange fluid flowing through the heat transfer bus 102, there is no longer a need to use other more specialized fluids (e.g., supercritical carbon dioxide) as the heat exchanger fluid. This reduction or elimination of an additional fluid coolant source minimizes the fuel volume amount of the engine fuel system, thereby reducing the heat load and pressure load on the thermal management system 100 and on the turbofan engine 10. In this way, the thermal management system 100 removes heat from the turbofan engine 10 and / or various accessory systems of the aircraft more effectively than existing thermal management systems that do not utilize fuel as the heat exchange fluid.

[0053] In addition, the fuel delivery system 122 is used to provide a reference pressure to maintain and / or control the pressure of the heat exchange fluid (e.g., fuel) flowing through the heat transfer bus 102. The fuel is recirculated in a loop through the heat transfer bus 102, which is pressurized and continuously charged by the reference pressure set by the fuel delivery system 122. In this way, the thermal management system 100 does not need to include different reservoir or accumulator components to accommodate the expansion, contraction, and charging of the components of the thermal management system 100. In one example, the fuel from the fuel metering unit 130 effectively acts as an accumulator or reservoir to control the pressure in the loop of the thermal management system 100. As the heat transfer bus 102 gets hotter (e.g., gains thermal energy), a small amount of fuel will migrate from the heat transfer bus 102 into the fuel delivery assembly 122. Conversely, as the heat transfer bus 102 gets colder (e.g., loses thermal energy), a small amount of fuel will migrate from the fuel delivery assembly 122 into the heat transfer bus 102.

[0054] However, it should be understood that the thermal management system 100 is provided only as an example, and in other exemplary embodiments, the thermal management system 100 may be constructed in any other suitable manner (see, for example Figure 3 the thermal management system 200).

[0055] Figure 3is a simplified schematic diagram of the thermal management system 200 and shows a heat transfer bus 202, a pump 204, a heat source exchanger 206 (including a first heat source exchanger 206A and a second heat source exchanger 206B), a radiator exchanger 208 (including a first radiator exchanger 208A and a second radiator exchanger 208B), a bypass line 210 (including a first bypass line 210A, a second bypass line 210B, a third bypass line 210C, and a fourth bypass line 210D), upstream joints 212A - 212D, downstream joints 214A - 214D, a flow control valve 218, a fuel metering unit 230, a fuel delivery line 234, a filter 238, a pressure relief valve 240, a bus control assembly 242 (having an overcurrent valve 244, a filling line 246, a first commissioning valve 248, a second commissioning valve 250, a control valve 252, and a bypass line 254), an injector outlet pressure source 256, an engine boost suction source 258, and a diverter valve 260. Figure 3 Also shown in Figure 3 is a simplified schematic diagram of the combustion section 26 of the turbofan engine 10. It should be noted that Figure 2 includes components that are the same as or similar to the components shown and discussed with respect to Figure 3 Here in Figure 2 components corresponding to the same or similar components in Figure 2 are assigned character labels that are 100 greater than the corresponding character labels assigned in Figure 3 For example, the thermal management system 200 shown in Figure 1 corresponds to the thermal management system 100 shown in Figure 2 and so on. It should be understood that the components of Figure 3 corresponding to the same or similar elements in Figure 2 are constructed in substantially the same manner as the corresponding components shown and described with respect to

[0056] The filter 238 is an element for removing specific particles or air from the liquid fuel passing through the filter 238. The pressure relief valve 240 is a valve for releasing or reducing the fluid pressure in a fluid circuit in fluid communication with the pressure relief valve 240. The bus control assembly 242 is a system for controlling the pressurization and commissioning of the heat transfer bus 202, and includes an overcurrent valve 244, a filling line 246, a first commissioning valve 248, a second commissioning valve 250, a control valve 252, and a bypass line 254. In one example, the bus control assembly 242 may be referred to as a pressure maintenance system. The overcurrent valve 244 is a valve for controlling the fluid flow by closing when there is a large change in the pressure of the fluid. The filling line 246 and the bypass line 254 are pipes or conduits for transporting fluid (such as fuel). The first commissioning valve 248, the second commissioning valve 250, and the control valve 252 are valves for controlling the fluid flow therethrough. In one example, the control valve 252 may be a calibrated orifice or a flow meter. The injector outlet pressure source 256 is a fuel source from the fuel injector outlet of the turbofan engine 10. The engine boost suction source 258 uses fuel to provide boost for the turbofan engine 10. The diverter valve 260 is a valve that divides the fluid flow into two or more parts.

[0057] The filter 238 is fluidly connected to the heat transfer bus 202 on a portion of the heat transfer bus 202 that extends between the second radiator exchanger 208B and the flow control valve 218. The pressure relief valve 240 is fluidly connected to a line or conduit that bypasses the filter 238. The bus control assembly 242 is fluidly connected to the heat transfer bus 202 at a point between the filter 238 and the flow control valve 218 and at a point between the flow control valve 218 and the pump 204. The overcurrent valve 244 is fluidly connected to the filling line 246 and is arranged in a parallel flow relationship with the first commissioning valve 248. The overcurrent valve 244 is disposed on the filling line 246 between the injector outlet pressure source 256 and the control valve 252. The filling line 246 is fluidly connected to the injector outlet pressure source 256 and a point along the heat transfer bus 202 that is between the flow control valve 218 and the pump 204 and extends therebetween.

[0058] The first commissioning valve 248 is positioned between the injector outlet pressure source 256 and the control valve 252 and is fluidly connected to the injector outlet pressure source 256 and the control valve 252. The second commissioning valve 250 is positioned between the engine supercharging suction source 258 and a point along the heat transfer bus 202 between the filter 238 and the flow control valve 218 and is in fluid communication therewith. The control valve 252 is fluidly connected to the charging line 246 between the overcurrent valve 244 and a point along the heat transfer bus between the flow control valve 218 and the pump 204 and is disposed on the charging line 246. The bypass line 254 is fluidly connected to the charging line 246 and extends through or around the control valve 252. The injector outlet pressure source 256 is fluidly connected to the heat transfer bus 202 via the charging line 246. The engine supercharging suction source 258 is fluidly connected to the heat transfer bus 202 at a point along the heat transfer bus 202 between the filter 238 and the flow control valve 218. The diverter valve 260 is fluidly connected to the flow metering unit 230 and the heat source exchanger 206C. The diverter valve 260 is disposed downstream of the flow metering unit 230 and upstream of the heat source exchanger 206C.

[0059] The filter 238 removes particles from the fuel flow through the heat transfer bus 202 at the filter 238. The pressure relief valve 240 controls the amount and pressure of the fuel passing through the filter 238. The bus control assembly 242 controls the amount and pressure of the fuel flowing into and out of the heat transfer bus 202.

[0060] The overcurrent valve 244 includes a diaphragm attached to an orifice such that the diaphragm does not move in response to a small equalizing flow or pressure equalization between the heat transfer bus 202 and a reference pressure set by the bus control assembly 242. In response to a large and / or sudden equalizing flow or pressure equalization between the heat transfer bus 202 and the reference pressure set by the bus control assembly 242, the overcurrent valve 244 closes to prevent fuel from flowing through the overcurrent valve 244. In one example, the overcurrent valve 244 provides a means for leak detection of the heat transfer bus 202. In the event of a leak in one of the heat source exchanger 206 or the radiator exchanger 208 (e.g., due to a crack or catastrophic failure), the pressure in the heat transfer bus 202 will suddenly or inexplicably drop due to a certain amount of fuel leaking from the heat transfer bus 202. In the case of such a leak, where there is a continuous fuel flow into the heat transfer bus 202 and the pressure drops, the overcurrent valve 244 will close, thereby isolating the leak in the thermal management system 200 from the rest of the fuel system of the turbofan engine 10.

[0061] The fill line 246 conveys or transfers fuel from the injector outlet pressure source 256, through the bus control assembly 242, and to the heat transfer bus 202. The first commissioning valve 248 and the second commissioning valve 250 are configured to control at least one of the purging and filling of the heat transfer bus. For example, both the first commissioning valve 248 and the second commissioning valve 250 are configured to assume one of a first closed position and a second open position. When the first commissioning valve 248 and / or the second commissioning valve 250 assumes the first closed position, the fuel flow or air flow into and out of the heat transfer bus 202 is blocked by the first commissioning valve 248 and / or the second commissioning valve 250. Similarly, when the first commissioning valve 248 and the second commissioning valve 250 assume the second open position, the fuel flow or air flow into and out of the heat transfer bus 202 is enabled.

[0062] In one example, the heat transfer bus 202 can be purged or filled (with air and / or fuel) by opening the first commissioning valve 248 and the second commissioning valve 250. For example, with the first commissioning valve 248 and the second commissioning valve 250 in the open position, fuel will be allowed to circulate through the heat transfer bus 202 to ensure that all surfaces within the heat source exchanger 206 and the radiator exchanger 208 are wetted. In another example, with the first commissioning valve 248 and the second commissioning valve 250 in the open position, air can be purged from the heat transfer bus 202 to prevent an excessive amount of air in the heat transfer bus 202.

[0063] The control valve 252 controls the amount of fuel delivered from the bus control assembly 242 to the heat transfer bus 202. In one example, the control valve is used to detect a situation where the overcurrent valve fails to close during a leak. In this way, the control valve 252 acts as a backup for a leak detection system or a leak mitigation system (e.g., the overcurrent valve 244). When in use, the bypass line 254 conveys a certain amount of fuel around or through the control valve 252. The injector outlet pressure source 256 supplies a pressurized fuel source to the heat transfer bus 202 via the bus control assembly 242. The engine boost suction source 258 receives fuel from the heat transfer bus 202 via the bus control assembly 242.

[0064] In one example, the change in the fuel quantity in the filling line 246 is monitored relative to a predetermined baseline fuel quantity of the filling line 246. The filling line 246 ensures that the pressure at the pump 204 is at the minimum level or a sufficient level of the fuel. In one example, the suction of the pump 204 can be pressurized to meet the minimum positive suction head requirement of the pump 204 to prevent cavitation. In another example, if there is a sudden or unexplained decrease in the volume and / or pressure in the heat transfer bus 202 due to a pipe rupture or a major failure in one of the heat source exchanger 206 or the radiator exchanger 208, an excessive amount of fuel in the filling line 246 can be detected. Once an excessive amount of fuel is detected in the filling line 246, the leak can be isolated from the rest of the thermal management system 200 by closing the overcurrent valve 244. In this example, the overcurrent valve 244 is set to selectively allow the fuel flow from the injector outlet pressure source 256 to the heat transfer bus 202.

[0065] Here, since the thermal management system 200 with the fuel-based heat transfer bus 202 includes a leak isolation capability, if a leak occurs, the operator (e.g., the pilot) will not need to shut down the turbofan engine 10 during use (e.g., during flight), thus allowing the continuous and safe use of the turbofan engine 10.

[0066] The benefits of the embodiments disclosed herein can include at least improved fuel combustion and thrust of the turbofan engine 10, leak identification and leak isolation occurring in the heat source exchanger 106 / 206 and the radiator elements 108 / 208 along the heat transfer bus 102 / 202, minimization of the overall engine fuel system volume, and utilization of the compressor cooling air and the waste heat recovery heat exchanger in the thermal management system 100 / 200.

[0067] This written description uses examples to describe aspects of the present disclosure and also enables those skilled in the art to practice the present disclosure, including manufacturing and using any device or system and performing any incorporated methods. The patentable 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 are not different from the literal language of the claims or if they include equivalent structural elements that are not substantially different from the literal language of the claims, then these other examples are intended to fall within the scope of the claims.

[0068] Further aspects are provided by the subject matter of the following clauses:

[0069] A gas turbine engine assembly, the gas turbine engine assembly including a gas turbine engine having a combustion section, a fuel delivery system, and a thermal energy management system. The fuel delivery assembly supplies fuel to the combustion section of the gas turbine engine. The thermal energy management system includes a heat transfer bus, a heat source heat exchanger, and a radiator heat exchanger. The heat transfer bus has a portion of the fuel configured to flow therethrough. The fuel is provided as the heat exchange fluid of the thermal energy management system. The heat source heat exchanger is in thermal communication with the fuel flow through the transfer bus. The radiator heat exchanger is in thermal communication with the fuel flow through the transfer bus.

[0070] The gas turbine engine assembly according to one or more of these clauses, further comprising: a pressure maintenance system, wherein the pressure maintenance system fluidly couples the fuel delivery assembly to the heat transfer bus, and wherein the pressure maintenance system is configured to pressurize and charge the portion of the fuel in the heat transfer bus.

[0071] The gas turbine engine assembly according to one or more of these clauses, wherein the heat source heat exchanger includes: a bus cooling air cooler heat exchanger in thermal communication with the fuel flow through the heat transfer bus, wherein the bus cooling air cooler heat exchanger is disposed along the heat transfer bus and upstream of the radiator heat exchanger; and a waste heat recovery heat exchanger in thermal communication with the fuel flow through the heat transfer bus, wherein the waste heat recovery heat exchanger is disposed along the heat transfer bus upstream of the radiator heat exchanger and downstream of the bus cooling air cooler heat exchanger.

[0072] The gas turbine engine assembly according to one or more of these clauses, wherein the heat transfer bus further includes a plurality of bypass lines for selectively bypassing any one of the radiator heat exchanger, the bus cooling air cooler heat exchanger, and the waste heat recovery heat exchanger.

[0073] A gas turbine engine component according to one or more of these clauses, wherein the plurality of bypass lines includes: a first bypass line having an upstream end and a downstream end, wherein the upstream end of the first bypass line is connected to the heat transfer bus at a location upstream of the bus cooling air cooler heat exchanger, wherein the downstream end of the first bypass line is connected to the heat transfer bus at a location downstream of the bus cooling air cooler heat exchanger, wherein the first bypass line is configured to convey a first portion of the fuel from upstream of the bus cooling air cooler heat exchanger to downstream of the bus cooling air cooler heat exchanger; a second bypass line having an upstream end and a downstream end, wherein the upstream end of the second bypass line is connected to the heat transfer bus at a location upstream of the heat source heat exchanger, wherein the downstream end of the second bypass line is connected to the heat transfer bus at a location downstream of the heat source heat exchanger, wherein the second bypass line is configured to convey a second portion of the fuel from upstream of the heat source heat exchanger to downstream of the heat source heat exchanger; and a third bypass line having an upstream end and a downstream end, wherein the upstream end of the third bypass line is connected to the heat transfer bus at a location upstream of the fuel-fuel heat exchanger, wherein the downstream end of the third bypass line is connected to the heat transfer bus at a location downstream of the fuel-fuel heat exchanger, wherein the third bypass line is configured to convey a third portion of the fuel from upstream of the fuel-fuel heat exchanger to downstream of the fuel-fuel heat exchanger.

[0074] A gas turbine engine component according to one or more of these clauses, further comprising: a radiator exchanger in thermal communication with the fuel flow through the transfer bus, wherein the radiator exchanger is disposed along the heat transfer bus at a location downstream of the radiator heat exchanger.

[0075] A gas turbine engine component according to one or more of these clauses, wherein the radiator exchanger is a surface cooler disposed along the surface of the fan section of the gas turbine engine.

[0076] A gas turbine engine component according to one or more of these clauses, wherein the pump is disposed along the heat transfer bus downstream of the radiator exchanger.

[0077] A gas turbine engine component according to one or more of these clauses, wherein the radiator exchanger, the radiator heat exchanger, the bus cooling air cooler heat exchanger, and the waste heat recovery heat exchanger are arranged in a serial flow arrangement along the heat transfer bus.

[0078] A gas turbine engine component according to one or more of these clauses, wherein the fuel delivery assembly further includes a commissioning valve configured to control at least one of purging and priming of the heat transfer bus.

[0079] A gas turbine engine component according to one or more of these clauses, wherein the commissioning valve is configured to occupy one of a first closed position and a second open position; wherein when the commissioning valve occupies the first closed position, the commissioning valve blocks the flow of fuel or air into and out of the heat transfer bus; and wherein when the commissioning valve occupies the second open position, purging air out of the heat bus or priming the heat transfer bus with fuel can be achieved.

[0080] A gas turbine engine component according to one or more of these clauses, wherein the pump includes a centrifugal pump configured to be mechanically coupled to and driven by an accessory gearbox of the gas turbine engine.

[0081] A gas turbine engine component according to one or more of these clauses, wherein the heat transfer bus is configured to isolate a portion of the fuel in the heat transfer bus from a portion of the fuel located downstream of the fuel metering unit of the fuel delivery assembly.

[0082] A gas turbine engine component according to one or more of these clauses, wherein the fuel delivery assembly further includes: a fuel metering unit; and a fuel delivery line that provides a fuel flow from the fuel metering unit to a combustion section of the gas turbine engine, wherein combustion of the fuel provides a driving force for the gas turbine engine.

[0083] A gas turbine engine component according to one or more of these clauses, wherein the thermal energy management system further includes a pump fluidly connected to the heat transfer bus, wherein the pump is configured to generate a flow of a portion of the fuel in the heat transfer bus.

[0084] A gas turbine engine component according to one or more of these clauses, wherein the radiator exchanger thermally connects the fuel flow through the heat transfer bus to the fuel flow through the fuel delivery line.

[0085] A gas turbine engine according to one or more of these clauses, or a gas turbine engine assembly according to one or more of these clauses, further comprises: a pressure control recirculation loop within the engine fuel system, the pressure control recirculation loop providing a reference pressure, and the gas turbine engine assembly having no discrete accumulator or reservoir to accommodate expansion, contraction, and / or filling of the thermal management system.

[0086] A thermal management system for a gas turbine engine assembly, the thermal management system comprising a heat transfer bus, a heat source heat exchanger, and a fuel-fuel heat exchanger. The heat transfer bus is configured to have a portion of the fuel flow passing through the heat transfer bus serve as the heat exchange fluid of the thermal management system. The heat source heat exchanger is in thermal communication with the fuel flow passing through the transfer bus. The fuel-fuel heat exchanger is configured to thermally connect the fuel flow passing through the heat transfer bus to the fuel flow passing through the fuel delivery assembly of the gas turbine engine.

[0087] The thermal management system according to one or more of these clauses, further comprising a pump fluidly connected to the heat transfer bus, wherein the pump is configured to generate flow of the portion of the fuel in the heat transfer bus.

[0088] A gas turbine engine, the gas turbine engine comprising a fan section, a compressor section, a combustion section, a turbine section, an exhaust section, a fuel delivery system, and a thermal management system. The fan section, the compressor section, the combustion section, the turbine section, and the exhaust section are arranged in a serial flow order and together define a core air flow path. The fuel delivery system supplies a fuel flow to the combustion section. The thermal management system comprises a first heat exchanger, a second heat exchanger, a third heat exchanger, a first radiator exchanger, and a heat transfer bus. The first heat exchanger is positioned in or downstream of the turbine section, the exhaust section, or both, in thermal communication with the flow passing through the core air flow path. The second heat exchanger is in thermal communication with the fuel delivery system for transferring heat to the fuel flow supplied to the combustion section. The third heat exchanger is connected to and in fluid communication with the heat transfer bus at a location upstream of the first heat exchanger. The first radiator is connected to and in fluid communication with the heat transfer bus at a location downstream of the second heat exchanger. The heat transfer bus has a heat exchange fluid flowing therethrough. The first heat exchanger, the second heat exchanger, the third heat exchanger, and the first radiator are each fluidly coupled to the heat transfer bus. The heat exchange fluid comprises deoxygenated fuel such that combustion of the deoxygenated fuel provides the driving force of the gas turbine engine.

[0089] A gas turbine engine according to one or more of these clauses, further comprising: wherein the first heat exchanger is a waste heat recovery heat exchanger; wherein the second heat exchanger is a fuel-fuel heat exchanger thermally coupled to the heat transfer bus and fluidly connected to the fuel delivery system; wherein the third heat exchanger is a bus cooling air cooler heat exchanger in fluid communication with the turbine section; and wherein the first radiator is a surface cooler mounted to the surface of the fan section of the gas turbine engine.

Claims

1. A gas turbine engine assembly, characterized in that, comprising: a gas turbine engine having a combustion section; a fuel delivery assembly that supplies fuel to the combustion section of the gas turbine engine; and a thermal management system, the thermal management system comprising: a heat transfer bus having a portion of the fuel configured to flow therethrough, wherein the fuel is provided as the heat exchange fluid of the thermal management system; a heat source heat exchanger in thermal communication with the fuel flow through the heat transfer bus; and a radiator heat exchanger in thermal communication with the fuel flow through the heat transfer bus; a pressure maintenance system, wherein the pressure maintenance system fluidly couples the fuel delivery assembly to the heat transfer bus, and wherein the pressure maintenance system is configured to pressurize, charge, and condition the portion of the fuel in the heat transfer bus.

2. The gas turbine engine assembly according to claim 1, characterized in that, wherein the heat source heat exchanger comprises: a bus cooling air cooler heat exchanger in thermal communication with the fuel flow through the heat transfer bus, wherein the bus cooling air cooler heat exchanger is disposed along the heat transfer bus and upstream of the radiator heat exchanger; and a waste heat recovery heat exchanger in thermal communication with the fuel flow through the heat transfer bus, wherein the waste heat recovery heat exchanger is disposed along the heat transfer bus upstream of the radiator heat exchanger and downstream of the bus cooling air cooler heat exchanger.

3. The gas turbine engine assembly according to claim 2, characterized in that, wherein the heat transfer bus further comprises a plurality of bypass lines for selectively bypassing any one of the radiator heat exchanger, the bus cooling air cooler heat exchanger, and the waste heat recovery heat exchanger.

4. The gas turbine engine assembly according to claim 3, characterized in that, wherein the plurality of bypass lines comprises: a first bypass line having an upstream end and a downstream end, wherein the upstream end of the first bypass line is connected to the heat transfer bus at a location upstream of the bus cooling air cooler heat exchanger, and wherein the downstream end of the first bypass line is connected to the heat transfer bus at a location downstream of the bus cooling air cooler heat exchanger, and wherein the first bypass line is configured to convey a first portion of the fuel from upstream of the bus cooling air cooler heat exchanger to downstream of the bus cooling air cooler heat exchanger; A second bypass line having an upstream end and a downstream end, wherein the upstream end of the second bypass line is connected to the heat transfer bus at a location upstream of the heat source heat exchanger, wherein the downstream end of the second bypass line is connected to the heat transfer bus at a location downstream of the heat source heat exchanger, and wherein the second bypass line is configured to convey a second portion of the fuel from upstream of the heat source heat exchanger to downstream of the heat source heat exchanger; and A third bypass line having an upstream end and a downstream end, wherein the upstream end of the third bypass line is connected to the heat transfer bus at a location upstream of the fuel - fuel heat exchanger, wherein the downstream end of the third bypass line is connected to the heat transfer bus at a location downstream of the fuel - fuel heat exchanger, and wherein the third bypass line is configured to convey a third portion of the fuel from upstream of the fuel - fuel heat exchanger to downstream of the fuel - fuel heat exchanger.

5. The gas turbine engine assembly according to claim 2, characterized in that, further comprising: A radiator exchanger in thermal communication with the fuel flow through the heat transfer bus, wherein the radiator exchanger is disposed along the heat transfer bus at a location downstream of the radiator heat exchanger.

6. The gas turbine engine assembly according to claim 5, characterized in that, wherein the radiator exchanger is a surface cooler disposed along the surface of the fan section of the gas turbine engine.

7. The gas turbine engine assembly according to claim 5, characterized in that, wherein a pump is disposed along the heat transfer bus downstream of the radiator exchanger.

8. The gas turbine engine assembly according to claim 5, characterized in that, wherein the radiator exchanger, the radiator heat exchanger, the bus cooling air cooler heat exchanger, and the waste heat recovery heat exchanger are arranged in a serial flow arrangement along the heat transfer bus.

9. The gas turbine engine assembly according to claim 1, characterized in that, wherein the fuel delivery assembly further comprises a commissioning valve configured to control at least one of purging and priming of the heat transfer bus.

10. The gas turbine engine assembly according to claim 9 ， characterized in that, wherein the commissioning valve is configured to assume one of a first closed position and a second open position; wherein when the commissioning valve assumes the first closed position, the commissioning valve blocks the flow of fuel or air into and out of the heat transfer bus; and wherein when the commissioning valve assumes the second open position, purging of air from the heat transfer bus or priming of the heat transfer bus with fuel can be achieved.

11. The gas turbine engine assembly according to claim 7, characterized in that, wherein the pump comprises a centrifugal pump configured to be mechanically coupled to and driven by an accessory gearbox of the gas turbine engine.

12. The gas turbine engine assembly according to claim 1, characterized in that, wherein the heat transfer bus is configured to isolate a portion of the fuel in the heat transfer bus from a portion of the fuel located downstream of the fuel metering unit of the fuel delivery assembly.

13. The gas turbine engine assembly according to claim 1, wherein, wherein the fuel delivery assembly further comprises: a fuel metering unit; and a fuel delivery line that provides a fuel flow from the fuel metering unit to the combustion section of the gas turbine engine, wherein combustion of the fuel provides a driving force for the gas turbine engine.

14. The gas turbine engine assembly according to claim 1, wherein, wherein the thermal energy management system further comprises a pump fluidly connected to the heat transfer bus, wherein the pump is configured to generate a flow of a portion of the fuel in the heat transfer bus.

15. The gas turbine engine assembly according to claim 5, wherein, wherein the radiator exchanger thermally connects the fuel flow through the heat transfer bus to the fuel flow through the fuel delivery line.

16. The gas turbine engine assembly according to claim 1, wherein, further comprising: a pressure control recirculation loop within the engine fuel system that provides a reference pressure, and the gas turbine engine assembly has no discrete accumulator or reservoir to accommodate expansion, contraction, and / or filling of the thermal energy management system.

17. A thermal energy management system for a gas turbine engine assembly, wherein, the gas turbine engine assembly includes a gas turbine engine having a combustion section and a fuel delivery assembly that supplies fuel to the combustion section of the gas turbine engine, and the thermal energy management system includes: a heat transfer bus configured to enable a portion of the fuel flow passing therethrough to serve as a heat exchange fluid for the thermal energy management system; a heat source heat exchanger in thermal communication with the fuel flow through the heat transfer bus; and a fuel-fuel heat exchanger configured to thermally connect the fuel flow through the heat transfer bus to the fuel flow through the fuel delivery assembly of the gas turbine engine; a pressure maintenance system, wherein the pressure maintenance system fluidly couples the fuel delivery assembly to the heat transfer bus, and wherein the pressure maintenance system is configured to pressurize, fill, and condition a portion of the fuel flow in the heat transfer bus.

18. The thermal energy management system according to claim 17, wherein, further comprising a pump fluidly connected to the heat transfer bus, wherein the pump is configured to generate a flow of a portion of the fuel in the heat transfer bus.

19. A gas turbine engine, wherein, comprising: A fan section, a compressor section, a combustion section, a turbine section, and an exhaust section, the fan section, the compressor section, the combustion section, the turbine section, and the exhaust section being arranged in a serial flow order and together defining a core air flow path; A fuel delivery system for providing a fuel flow to the combustion section; And A thermal management system, the thermal management system including: A first heat exchanger positioned in heat communication with a flow through the core air flow path within or downstream of the turbine section, the exhaust section, or both; A second heat exchanger in thermal communication with the fuel delivery system for transferring heat to the fuel flow provided to the combustion section; A third heat exchanger connected to and in fluid communication with a heat transfer bus at a location upstream of the first heat exchanger; A first radiator connected to and in fluid communication with the heat transfer bus at a location downstream of the second heat exchanger; and A heat transfer bus having a heat exchange fluid flowing therethrough, wherein the first heat exchanger, the second heat exchanger, the third heat exchanger, and the first radiator are each fluidly coupled to the heat transfer bus, wherein the heat exchange fluid includes deoxygenated fuel, and wherein combustion of the deoxygenated fuel provides a driving force for the gas turbine engine; A pressure maintenance system, wherein the pressure maintenance system fluidly couples the fuel delivery assembly to the heat transfer bus, and wherein the pressure maintenance system is configured to pressurize, charge, and commission the heat exchange fluid in the heat transfer bus.

20. The gas turbine engine according to claim 19, wherein, further comprising: wherein the first heat exchanger is a waste heat recovery heat exchanger; wherein the second heat exchanger is a fuel-fuel heat exchanger thermally coupled to the heat transfer bus and fluidly connected to the fuel delivery system; wherein the third heat exchanger is a bus cooling air cooler heat exchanger in fluid communication with the turbine section; and wherein the first radiator is a surface cooler mounted to a surface of the fan section of the gas turbine engine.

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