Fuel heaters and energy conversion systems

Abstract

A system for energy conversion is provided, comprising a propulsion system, a fuel circuit, a combustion unit and a turbine separate from the propulsion system, and a load device. A fuel flow control device divides a fuel stream from a fuel tank into a first fuel portion and a second fuel portion. The fuel circuit is configured to supply the first fuel portion to a heat addition system at the propulsion system. The combustion unit receives an oxidant stream from the compressor section of the propulsion system via a fluid circuit. The fuel circuit supplies the second fuel portion to the combustion unit. The fluid circuit directs combustion gases to the turbine and the propulsion system. The load device is operatively coupled to the turbine to receive output torque via the expansion of the combustion gases through the turbine.

Description

[0001] This application is a divisional application of the invention patent application filed on February 7, 2022, with application number 202210115712.2 and title "Fuel Heater and Energy Conversion System".

[0002] Cross-references to related applications

[0003] This application is a non-provisional application that claims priority to the following U.S. provisional patent application filed on February 3, 2021, under 35 U.S. SC § 119(e): U.S. Non-Provisional Application No. 17 / 166,245, the entire contents of which are incorporated herein by reference. Technical Field

[0004] This topic broadly relates to energy conversion and thermal management systems used in propulsion systems and vehicles. Specifically, it covers thermal management and energy conversion systems used in aircraft and propulsion systems. Background Technology

[0005] Propulsion systems (such as gas turbine engines) face increasingly higher thermal loads and challenges in thermal management of energy conversion. These increasing thermal loads and energy demands are due to the increasing electrification of propulsion systems and vehicles (such as aircraft), larger electrical loads, and the need to improve the thermal efficiency of fuel systems, oil systems, and cooling fluids.

[0006] Traditional systems that generate auxiliary power to meet increased power demands are limited by the power output of the propulsion system. Low-power and partial-power conditions may be insufficient to generate heat for thermal management and energy conversion systems.

[0007] Therefore, an energy conversion system is needed to meet the increased power generation requirements of the launch vehicle and propulsion system without being limited by the power output from the propulsion system. Furthermore, an energy conversion and thermal management system is also required to meet the challenges associated with increased thermal loads. Summary of the Invention

[0008] Aspects and advantages of the invention will be set forth in part in the description which follows, or may be apparent from the description or may be learned by practice of the invention.

[0009] This disclosure relates to a system for energy conversion. The system includes a propulsion system comprising a compressor section, a heat-adding system, and an expansion section arranged in series flow. A fuel circuit is fluidly connected from a fuel tank to a fuel flow control device. The fuel flow control device divides a fuel flow from the fuel tank into a first fuel portion and a second fuel portion. The fuel circuit is configured to supply the first fuel portion to the heat-adding system. A combustion device is configured to receive an oxidant flow from the compressor section via a fluid circuit. The fuel circuit is fluidly connected to supply the second fuel portion to the combustion device. The combustion device is configured to generate combustion gases from the second fuel portion and the oxidant flow. The fluid circuit is fluidly connected to the combustion device to allow the combustion gases to flow to the propulsion system. A turbine is configured to receive the combustion gases from the combustion device via a fluid circuit. A load device is operatively coupled to the turbine via a drive shaft. The load device is configured to receive output torque from the drive shaft via the expansion of the combustion gases through the turbine.

[0010] These and other features, aspects, and advantages of the invention will be better understood by referring to the following description and the appended claims. The accompanying drawings, which are incorporated in and form part of this specification, illustrate embodiments of the invention and, together with the description, serve to explain the principles of the invention. Attached Figure Description

[0011] The complete and practical disclosure of the invention, including its preferred mode, is set forth in the description with reference to the accompanying drawings, for those skilled in the art, wherein:

[0012] Figure 1 These are exemplary embodiments of a launch vehicle including a propulsion system and an energy conversion system, according to aspects of this disclosure;

[0013] Figure 2 This is an exemplary schematic embodiment of an energy conversion system including a propulsion system according to aspects of this disclosure;

[0014] Figure 3 These are exemplary schematic embodiments of energy conversion systems according to aspects of this disclosure; and

[0015] Figure 4-6 This is an exemplary schematic embodiment of an energy conversion and thermal management system according to aspects of this disclosure.

[0016] Reference numerals used repeatedly in this specification and drawings are intended to indicate the same or similar features or elements of the invention. Detailed Implementation

[0017] Reference will now be made in detail to embodiments of the invention, one or more examples of which are illustrated in the accompanying drawings. Each example is provided to explain the invention and not to limit it. In fact, it will be apparent to those skilled in the art that various modifications and variations can be made to the invention without departing from the scope or spirit of the invention. For example, a feature shown or described as part of one embodiment may be used with another embodiment to produce yet another embodiment. Therefore, the invention is intended to cover such modifications and variations that fall within the scope of the appended claims and their equivalents.

[0018] As used herein, the terms “first,” “second,” and “third” are used interchangeably to distinguish one component from another and are not intended to indicate the location or importance of the individual components.

[0019] The terms "upstream" and "downstream" refer to the relative directions of fluid flow within a fluid path. For example, "upstream" refers to the direction from which the fluid flows, and "downstream" refers to the direction from which the fluid flows.

[0020] This document provides embodiments of propulsion systems, launch vehicles, and energy conversion systems that can improve overall system and launch vehicle efficiency, for example, by utilizing a relatively small amount of bleed air or other oxidizer from the propulsion system or auxiliary power unit to generate combustion gases from a combustion system separate from the combustion system at the propulsion system or auxiliary power unit. The combustion gases are then expanded by a power turbine to generate output torque for the load unit. A fuel-air heat exchanger is positioned in thermal communication with at least a portion of the fuel flow. The combustion gases generated via a dedicated combustion system separate from the propulsion system's combustion system allow for the generation and release of heat, which is substantially decoupled from the power output from the propulsion system or engine speed. Some embodiments include receiving an oxidizer flow from the propulsion system and generating and expanding combustion gases at a combustion device and turbine separate from the propulsion system. Other embodiments include thermal communication of the combustion gases with one or more heat exchangers.

[0021] The embodiments of the energy conversion and thermal management system provided herein allow for weight-advantageous systems and methods for fuel heating and fuel temperature control independent of engine speed from the propulsion system. System embodiments may include the unexpectedly beneficial results of fuel heating and power generation from a combustion system and turbine that are separate from the propulsion system to produce load devices (e.g., motors, accessory gear assemblies, pumps, general mechanical loads, or other systems typically powered mechanically, electrically, or pneumatically by the propulsion system).

[0022] Now referring to the attached diagram, in Figure 1The present disclosure provides exemplary embodiments of a vehicle 100 including a propulsion system 10 and an energy conversion system 200 according to aspects thereof. In embodiments, the vehicle 100 is an aircraft including an aircraft structure or airframe 105. The airframe 105 includes a fuselage 110, wings 120, and a tail 130 attached to the fuselage 110. The propulsion system 10 according to aspects thereof is attached to one or more portions of the airframe. In some embodiments, the vehicle 100 includes an auxiliary power unit (APU) 15. The APU 15 may form a gas turbine engine including, for example, a compressor section, a heat addition system, an expansion section, and an exhaust section as further described herein. In various embodiments, the energy conversion system 200 is a system configured to desirably distribute heat loads, for example, by adding or removing heat from one or more fluids or structures, such as, but not limited to, oxidants at the propulsion system, fuels for motors, electronic devices, computing systems, environmental control systems, gear assemblies, or other systems or structures, lubricants, hydraulic fluids, pneumatic fluids, or cooling fluids.

[0023] In some cases, the propulsion system 10 is attached to the rear of the fuselage 110. In other cases, the propulsion system 10 is attached below, above, or via a portion of the wing 120 and / or tail 130. In various embodiments, the propulsion system 10 is attached to the frame 105 via a pylon or other mounting structure. In other embodiments, the propulsion system 10 is housed within the frame, as exemplified in certain supersonic military or commercial aircraft.

[0024] Various embodiments of the launch vehicle 100 include a computing system 140, such as avionics or other electronic devices or computing devices configured to control the launch vehicle 100 or propulsion system 10. The launch vehicle 100 may also include an environmental control system (ECS) 150, for example, to provide thermoconditioned air to the launch vehicle compartment, computing system 140, launch vehicle surface anti-icing system 160, propulsion system anti-icing system, or other systems of the launch vehicle 100 or propulsion system 10. In various embodiments such as those described herein, the energy conversion system 200 may be configured to provide energy to one or more subsystems of the launch vehicle or propulsion system, as described above and further herein. Further embodiments of the energy conversion system may be configured to provide thermoconditioned fluid to one or more systems described herein.

[0025] Now for reference Figure 2 It provides a system for propulsion system 10 or APU 15 ( Figure 1An exemplary schematic embodiment of engine 13 is shown, which is operatively coupled to energy conversion system 200. Specific embodiments of propulsion system 10 may be configured as a turbine, ramjet engine, or scramjet engine. Other specific embodiments of propulsion system 10 may include a turbine configured as a turbofan engine, turboprop engine, turbojet engine, turboshaft engine, propeller fan engine, or open rotor engine. Figure 2 In this embodiment, the propulsion system 10 is configured as a three-flow engine comprising a fan bypass flow 14, a core flow path 70, and a core bypass or third flow 71. Certain embodiments of the propulsion system 10 include a fan section 12, a compressor section 20, a combustion section or heat additive system 26, an expansion section 30, and an exhaust section 36 arranged in series flow. In various embodiments, the heat additive system 26 may be configured as a detonation combustion system or a knock combustion system. The heat additive system 26 may include any suitable type of system for receiving liquid and / or gaseous fuel flows and generating hot gases, including but not limited to annular, can-annular, canister, trap, volute or vortex, rotating detonation, pulse detonation, subsonic or supersonic combustion systems. The fan section 12 includes one or more stages of rotors and blades 121. Some embodiments also include one or more stages of impeller blades stationary relative to the centerline axis of the propulsion system 10.

[0026] Compressor section 20, heat addition system 26, and expansion section 30 are positioned in a series aerodynamic flow arrangement. Compressor section 20, heat addition system 26, and expansion section 30 may together define the core engine or gas generator of propulsion system 10. For example, as described herein, in some embodiments, compressor section 20 includes a high-pressure compressor 24 arranged in direct series flow with heat addition system 26 and high-pressure turbine 32 of expansion section 30. Low-pressure turbine 34 of expansion section 30 may be operatively coupled to fan section 12 to drive one or more stages of fan section 12. In some embodiments, propulsion system 10 may include a low-pressure compressor or intermediate-pressure compressor 24 aerodynamically positioned between fan section 12 and high-pressure compressor 24. In a further embodiment, intermediate-pressure turbine may be aerodynamically positioned between high-pressure turbine 32 and low-pressure turbine 34.

[0027] The core flow path 70 extends at least through the high-pressure compressor 24, the heat addition system 26, and the high-pressure turbine 32. A core bypass or third flow path 71 extends downstream of the intermediate- or low-pressure compressor 22 and bypasses the core flow path 70 at the high-pressure compressor 24. In some embodiments, the third flow path 71 is in fluid communication with a fan bypass flow 14 downstream of the impeller 122.

[0028] The third flow path 71 is an airflow configured to recover fluid energy to generate a portion of the total thrust of the propulsion system 10. In one embodiment, a portion of the total thrust generated through the third flow path 71 may include a dedicated exhaust nozzle at the outlet end. In another embodiment, a portion of the total thrust generated through the third flow path 71 may be mixed with the fan bypass flow 14. In yet another embodiment, a portion of the total thrust generated through the third flow path 71 may be mixed with the core flow path 70 downstream of the heat addition system 26 and exit through the exhaust section 36. Various embodiments of the third flow path 71 are configured to generate less than 50% of the total thrust of the propulsion system 10.

[0029] Those skilled in the art will understand that the third flow path 71 extends from the core flow path 70 upstream of the heat addition system 26 and downstream of the fan section 12, and is further configured to allow airflow to exit the propulsion system 10 as part of the total thrust of the propulsion system 10. The operating temperature of the air passing through the third flow path 71 can typically correspond to the temperature range of the air exiting the medium- or low-pressure compressor 22.

[0030] Now for reference Figure 3-6 The energy conversion system 200 includes a conduit, manifold, or other wall-mounted conduit forming a fluid circuit 210 that provides fluid communication with an oxidant (or specifically air) flow 201 extracted from the compressor section 20 of the engine 13. In a particular embodiment, the oxidant flow 201 is drawn from or otherwise directed from a portion of the oxidant compressed by the compressor assembly 220. In a particular embodiment, the oxidant flow 201 is drawn from the engine 13 ( Figure 2 The system 200 can be extracted from a single-stage or multi-stage high-pressure compressor 24. However, it should be understood that some embodiments of the system 200 may receive power from a medium-pressure compressor or a low-pressure compressor 22. Figure 2 The oxidant flow is described above. In a further embodiment, the oxidant flow 201 can be received from a dedicated compressor separate from the compressor section 20 of the propulsion system 10. A flow control device 205 (e.g., a valve or other suitable control mechanism) can desiccate or modulate the amount or magnitude of the oxidant flow extracted from the compressor section 20.

[0031] Fluid circuit 210 is configured to provide fluid communication of oxidant flow 201 from compressor section 20 to combustion device 230. Combustion device 230 can be any suitable type of detonation or knock combustion device configuration. Embodiments may include, but are not limited to, annular burners, canister burners, canister-annular burners, trap burners (TVC), involute or vortex burners, fuel-rich burners, lean burners, pulse detonation burners, rotary detonation burners, or combinations thereof, or other suitable types of detonation or knock combustion systems.

[0032] In a particular embodiment, the flow control device 205 may limit, regulate, or control the oxidant flow 201 received by the combustion device 230 such that the flow is stabilized or otherwise within a specific or desired parameter range. The desired parameter range may be a specific flow rate or pressure of the oxidant 201 supplied to the combustion device 230, or a combination thereof. The flow control device 205 may generally allow combustion within desired operating limits. Such limits may include emissions or greenhouse gases (e.g., nitrogen oxides, soot, unburned hydrocarbons, carbon dioxide, carbon monoxide, etc.). Limits may additionally or alternatively include those associated with lean burn, rich burn, reignition, combustion stability, pressure oscillations, acoustics, or other performance or operability parameters of the combustion system.

[0033] It should also be understood that the flow control device 205 can allow for stable operation of the combustion device 230 configured as a knock burner. The flow control device 205 can provide an oxidizer flow 201 within a pressure and / or flow rate range specific to the operation of the combustion device 230 as a knock burner. The combustion device 230 configured as a knock burner can further provide improvements in energy and thermal efficiency compared to a knock burner configuration. Furthermore, the combustion device 230 can be configured within an operating range suitable for driving the turbine 240, different from the operation of the heat addition system 26 and expansion section 30 at the propulsion system 10. For example, the system 200 can be configured to operate the combustion device 230 within a narrower operating range or steady state than the operating range or steady state of the propulsion system 10. In another example, the propulsion system 10, such as that for an aircraft, is typically configured for operating conditions corresponding to the landing-takeoff cycle of the aircraft. In contrast, the flow control device 205 can be configured to provide an oxidizer flow 201 to the combustion device 230 within an operating range that avoids problems associated with transient operation or differences from relatively low power output (e.g., ignition, idle) to relatively high power output (e.g., takeoff). Additionally or alternatively, the compressor assembly 220 may be of the APU 15. The APU 15 can typically be configured to operate in a relatively stable operating state, as described above. Other embodiments may receive the oxidizer flow 201 from a compressor or pump that is separate from the aerodynamic or thermodynamic flow path of the propulsion system 10 or the APU 15. Such embodiments include a compressor section forming an electric motor-driven air compressor.

[0034] System 200 also includes a conduit, manifold, or wall-mounted conduit forming a fuel circuit 310 extending fluidly from fuel tank 300. Fuel tank 300 contains liquid and / or gaseous fuel for mixing and combustion / knock at combustion device 230. Fuel circuit 310 is configured to provide a fuel flow to combustion device 230, schematically depicted via arrow 302. In a particular embodiment, fuel tank 300 is also in fluid communication with engine 13's heat addition system 26 to provide a liquid and / or gaseous fuel flow (schematically depicted via arrow 305) for generating combustion gases to expand at expansion section 30.

[0035] In some embodiments, system 200 includes a fuel flow control device 307 configured to provide a first portion of fuel (schematically depicted via line 303) toward the heat addition system 26 of engine 13 and a second portion of fuel (schematically depicted via line 304) toward combustion device 230. The fuel flow control device 307 may form a valve, diverter, or other suitable mechanism for dividing the fuel flow 301 into a first portion 303 and a second portion 304. In some embodiments, the fuel control device 307 is configured to control the amount or quantity of the second portion of fuel 304 supplied to combustion device 230. In some embodiments, the fuel control device 307 determines the output energy of combustion gas 202 to turbine 240 at least in part, for example, by regulating or modulating the amount of fuel supplied to combustion device 230 and burned / knocked at combustion device 230. In a particular embodiment, flow control device 205 ( Figure 2 Furthermore, for example, the output energy of the combustion gas 202 to the turbine 240 is determined by adjusting or modulating the oxidizing dose supplied to and mixed with the fuel 302 for combustion / knock at the combustion device 230.

[0036] During operation, combustion device 230 generates and supplies combustion gas 202 to drive turbine 240. Turbine 240 is operatively coupled to load device 270 via drive shaft 241. During operation, as combustion gas 202 expands through turbine 240, turbine generates output torque and supplies power to load device 270 via transmission through drive shaft 241. Load device 270 may include one or more fuel pumps, electric motors (e.g., motors and / or generators, constant frequency or variable frequency machines, hybrid power systems, etc.), lubricant pumps, hydraulic pumps, air compressors, engine starters, sensor drivers (e.g., one or more sensor devices, instrument sensors, or telemetry, including but not limited to transducers, capacitors, slip rings, thermocouples, electronic measuring devices, or computing systems), and auxiliary gearbox drivers, or combinations thereof.

[0037] The expanded combustion gases from turbine 240 (schematically depicted via line 203) are supplied to engine 13. Reference Figure 2-3 In some embodiments, the combustion gases supplied to engine 13 are specifically supplied to core flow path 70. In one embodiment, the combustion gases are supplied from turbine 240 to core flow path 70 at exhaust section 36 of engine 13. In another embodiment, the combustion gases are supplied from turbine 240 to core flow path 70 at expansion section 30 of engine 13. In yet another specific embodiment, the combustion gases are supplied from turbine 240 to core flow path at intermediate-pressure turbine or low-pressure turbine 34 at expansion section 30 of engine 13. In other embodiments, combustion gas flow 204 is supplied to one or more of fan flow 14 or third flow path 71.

[0038] In a particular embodiment, a second portion of fuel (schematically depicted via line 304) is provided in thermal communication with the combustion gas stream 203 exiting the turbine 240 via a fuel-air heat exchanger 250. The fluid is typically an oxidant, such as the air stream passing through the propulsion system 10. In one embodiment, the fuel-air heat exchanger 250 is configured to transfer heat, or thermal energy, from the combustion gas stream 203 downstream of the turbine 240 to the fuel stream 304 upstream of the heat addition system 26. The cooled combustion gas stream (schematically depicted via line 204) is provided from the fuel-air heat exchanger 250 to the engine 13, as described above. The heated fuel stream (schematically depicted via line 305) is provided from the fuel-air heat exchanger 250 to the heat addition system 26 of the engine 13.

[0039] In some embodiments, the deaerator 320 is positioned in a flow arrangement along the fuel circuit 310. The deaerator 320 is configured to remove oxygen from the second portion of fuel 303. In embodiments, the deaerator 320 receives energy or power from a load device 270. The energy may be provided from the load device 270 via shaft, electrical power, or other suitable energy transfer methods. The deaerated fuel flow is provided downstream of the fuel-air heat exchanger 250.

[0040] Now for reference Figure 4 The construction of System 200 and Figure 2-3 The basic similarities between the descriptions and depictions are as follows. Figure 4 In the depicted embodiment, the energy conversion system 200 also includes a thermal management system 400, which includes a wall-mounted duct forming a heat transfer fluid loop 410 configured to provide a heat transfer fluid in thermal communication with the combustion gas flow 203 exiting the turbine 240. Figure 4In this embodiment, the fuel-air heat exchanger 250 provides combustion gas 203 that is received from the turbine 240 and is in thermal communication with a heat transfer fluid flow (schematically depicted via line 401) received from the heat transfer fluid flow device 405. In various embodiments, the flow device 405 is operatively coupled to and driven by a load device 270, such as those described with respect to the deaerator 320.

[0041] The thermal management system 400 is also configured to provide a heat transfer fluid, schematically depicted via line 402, that is in thermal communication with the second portion of fuel 304 through a first heat bus heat exchanger 410. The heat exchanger 410 is configured to heat the fuel stream 304 by receiving heat or thermal energy from the heat transfer fluid stream 402. The heat transfer fluid stream 402 receives heat or thermal energy from the combustion gas stream 203 exiting the turbine 240.

[0042] In various embodiments, the heat transfer fluid is a lubricant (e.g., oil, oil-based fluid, synthetic oil, polyalphaolefin, polyalphaolefin-based fluid, etc., or combinations thereof), a liquid and / or gaseous fuel (e.g., hydrocarbon fuel, fuel oil, aviation turbine fuel, or other suitable propulsion system fuel), a supercritical fluid (e.g., supercritical carbon dioxide, water, methane, ethane, propane, ethylene, propylene, methanol, ethanol, acetone, nitrous oxide, or other suitable substances at temperatures and pressures above the endpoints of their phase equilibrium curves), or a silicone or silicone-based heat transfer fluid (e.g., a polydimethylsiloxane fluid, such as Syltherm). TM (or similar fluids), or other suitable heat transfer fluids.

[0043] In some embodiments, system 200 further includes a second heat exchanger 420 positioned in thermal communication with a heat transfer fluid (e.g., schematically depicted via line 403). The second heat exchanger 420 provides thermal communication between the heat transfer fluid flow 403 and a cooling fluid flow 501 provided by an engine cooling flow 500. The engine cooling flow 500 is a relatively cold oxidant flow from engine 13, such as fan flow 14 or a third flow path 71 (…). Figure 2 The oxidizer flow is located at [location missing]. The second heat exchanger 420 is configured to remove heat energy or heat from the heat transfer fluid flow 403 and transfer the heat energy or heat to the cooling fluid flow 501. The cooled heat transfer fluid flow (schematically depicted via line 404) is provided to the fuel-air heat exchanger 250, for example, to cool the combustion gas flow 203. The cooled combustion gas flow 204 is provided to the engine 13, for example, to cool one or more components of the intermediate-pressure turbine or low-pressure turbine 34, or as part of an active clearance control system.

[0044] Now for reference Figure 5 The construction and related aspects of System 200 Figure 4 The descriptions are basically similar. In Figure 5In addition, system 200 also includes a combustion gas flow control device 207 configured to divide the combustion gas flow 202 into a first portion 202a supplied to turbine 240 and a second portion 202b bypassing turbine 240. The first portion of combustion gas 202a expands through turbine 240. When energy is released through turbine 240, the expanded combustion gas 203 is cooler relative to combustion gas 202. The relatively cool, expanded combustion gas 203 is supplied to one or more suitable modules, components, or subsystems at engine 13, for example, using a relatively low-pressure cooling fluid.

[0045] Compared to the combustion gas flow 203 exiting the turbine 240, the bypassed combustion gas flow 202b provides a relatively higher pressure flow. The bypassed combustion gas flow 202b is provided in thermal communication with, for example, a fuel-air heat exchanger 250 as described above. The cooled combustion gas flow 204 is provided to one or more portions of, for example, the expansion section 30 as described above.

[0046] In a particular embodiment, the flow control device 207 and the turbine 240 are integrated as a variable zone turbine nozzle (VATN). The flow control device 207 and the turbine 240, together defining the VATN, are configured to control the mass flow entering or passing through the turbine 240. In one embodiment, the flow control device 207 is a variable zone structure at the inlet of the turbine 240, such as a variable zone nozzle. However, in other embodiments, the flow control device 207 is a separate flow control structure configured to regulate or modulate the mass flow through the turbine 240. In some embodiments, the turbine 240 is downstream of the flow control device 207.

[0047] Now for reference Figure 6 The construction of the provided system 200 embodiments and related Figure 1-4 The descriptions are basically similar. In Figure 6 In this embodiment, turbine 240 is further connected to compressor 245 via drive shaft 242. Oxidant stream 201 is received from compressor section 220, for example, as described above. Compressor 245 further compresses or agitates oxidant stream 201 before supplying compressed oxidant stream (schematically depicted via arrow 201a) to combustion device 230. In such an embodiment, compressor 245, operatively coupled to turbine 240, allows a relatively low-pressure oxidant stream to be drawn from compressor section 20 of engine 13 (e.g., from intermediate or low-pressure compressor 22, or from one or more pre- or upstream stages of high-pressure compressor 24 (generally opposite to high-pressure compressor 24) or one or more downstream stages thereof). In other embodiments, compressor 245 allows a relatively low-pressure oxidant stream to be received from a relatively low-pressure compressor or pump assembly.

[0048] The embodiments of the energy conversion system 200, propulsion system 10, or vehicle 100 provided herein can provide improved overall propulsion system and vehicle efficiency through improved systems, structures, or methods for energy conversion or thermal management, such as those provided herein. The embodiments provided herein include specific positioning, placement, and serial flow of fluids configured to improve overall system performance. One or more components of the system (e.g., but not limited to, heat exchangers, loops, ducts, flow devices, combustion devices, turbines, or compressors provided herein) can be manufactured via one or more additive manufacturing methods described below. Furthermore, such a system would be impossible without flow paths, ducts, loops, structures, or other details permitted by additive manufacturing methods. Additionally, certain arrangements provided herein can produce beneficial and unintended results via various fluids or serial flows, such as those provided herein, that transfer heat or thermal energy to specific junctions.

[0049] One or more components of the propulsion system 10 and energy conversion system 200 described herein can be manufactured or formed using any suitable process (e.g., additive manufacturing processes, such as 3D printing). Using such a process allows such components to be formed integrally, as a single monolithic component, or as any suitable number of sub-components, or at a scale and complexity previously not permitted or conceived in the art. In particular, additive manufacturing processes can allow such components to be formed integrally and include a variety of features that are not possible using existing manufacturing methods. For example, the additive manufacturing methods described herein can allow the combustion device 230 or one or more heat exchangers to be manufactured to a size, scale, and complexity previously unconcerned in the art. As another example, the additive manufacturing methods described herein can allow the turbine 240 and the flow control device to be manufactured as a single monolithic component. In further embodiments, the additive manufacturing methods described herein allow the manufacture of turbines, flow control devices, combustion devices, or circuits with unique features, constructions, thicknesses, materials, densities, fluid channels, manifolds, and mounting structures that are not possible or impractical using prior manufacturing methods. Despite the addition of additional combustion systems, turbines, compressors, or heat exchangers, additive manufacturing can allow for combinations of such structures and their specific flow and thermal arrangements, which can improve thermal efficiency, enhance energy conversion, and improve the overall efficiency or operability of the propulsion system or vehicle.

[0050] Suitable additive manufacturing techniques according to this disclosure include, for example, fused deposition modeling (FDM), selective laser sintering (SLS), 3D printing such as by inkjet, laser jetting, and binder jetting, stereolithography (SLA), direct selective laser sintering (DSLS), electron beam sintering (EBS), electron beam melting (EBM), laser engineered net-shape forming (LENS), laser net-shape manufacturing (LNSM), direct metal deposition (DMD), digital light processing (DLP), direct selective laser melting (DSLM), selective laser melting (SLM), direct metal laser melting (DMLM), and other known processes. Suitable powder materials for manufacturing the structures provided herein as a single, integral structure or structures of the scale and complexity provided herein include metal alloys, polymers, or ceramic powders. Exemplary metal powder materials are stainless steel alloys, cobalt-chromium alloys, aluminum alloys, titanium alloys, nickel-based superalloys, and cobalt-based superalloys. Furthermore, suitable alloys may include those designed to have good oxidation resistance, referred to as “superalloys,” which possess acceptable strength at elevated operating temperatures in gas turbine engines, such as Hastelloy, Inconel alloys (e.g., IN 738, IN 792, IN 939), Rene alloys (e.g., Rene N4, Rene N5, Rene 80, Rene 142, Rene 195), Haynes alloys, Mar M, CM 247, CM 247LC, C263, 718, X-850, ECY 768, 282, X45, PWA 1483, and CMSX (e.g., CMSX-4) single-crystal alloys. The manufactured objects of this disclosure may be formed with one or more selected crystalline microstructures, such as directional solidification (“DS”) or single crystals (“SX”).

[0051] This written description uses examples to disclose the invention, including the best mode, and also enables those skilled in the art to practice the invention, including making and using any device or system and performing any combined methods. The patentable scope of the invention is defined by the claims, but may include other examples that would occur to those skilled in the art. These other examples are intended to fall within the scope of the claims if they include structural elements that are not indistinguishable 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.

[0052] Further aspects of the invention are provided by the subject matter of the following clauses:

[0053] 1. A system for energy conversion, the system comprising: a propulsion system including a compressor section, a heat addition system, and an expansion section arranged in series flow; a fuel circuit fluidly connected from a fuel tank to a fuel flow control device, wherein the fuel flow control device divides a fuel flow from the fuel tank into a first portion of fuel and a second portion of fuel, wherein the fuel circuit is configured to supply the first portion of fuel to the heat addition system; a combustion device configured to receive an oxidant flow from the compressor section via the fluid circuit, wherein the fuel circuit extends fluidly to supply the second portion of fuel to the combustion device, and wherein the combustion device is configured to generate combustion gases from the second portion of fuel and the oxidant flow, and further wherein the fluid circuit is fluidly connected to the combustion device to allow the combustion gases to flow to the propulsion system; a turbine configured to receive the combustion gases from the combustion device via the fluid circuit; and a load device operably coupled to the turbine via a drive shaft, wherein the load device is configured to receive output torque from the drive shaft via the combustion gases through the expansion of the turbine.

[0054] 2. The system according to any one or more clauses herein, wherein the fluid circuit is in fluid communication with the combustion device to allow the combustion gases to flow to the expansion section of the propulsion system.

[0055] 3. The system according to any one or more clauses herein, wherein the fluid circuit is in fluid communication with the combustion device to allow the combustion gases to flow to the low-pressure turbine of the expansion section of the propulsion system.

[0056] 4. The system according to any one or more clauses herein, the system comprising: a fuel-air heat exchanger positioned along the fuel circuit, the fuel-air heat exchanger being in thermal communication with the first portion of the fuel.

[0057] 5. The system according to any one or more clauses herein, wherein the fuel-air heat exchanger is positioned along the fluid loop, wherein the fuel-air heat exchanger is configured to provide heat transfer between the combustion gases and the first portion of fuel.

[0058] 6. The system according to any one or more clauses herein, the system comprising: a heat transfer fluid loop configured to provide a heat transfer fluid flow in thermal communication with the fuel-air heat exchanger, wherein the fuel-air heat exchanger is configured to provide heat transfer between the first portion of fuel and the heat transfer fluid.

[0059] 7. The system according to any one or more clauses herein, the system comprising: a first heat bus heat exchanger configured to provide heat transfer between the heat transfer fluid and the combustion gas.

[0060] 8. The system according to any one or more clauses herein, the system comprising: a second heat bus heat exchanger configured to provide heat transfer between the heat transfer fluid and the cooling fluid flow.

[0061] 9. The system according to any one or more clauses herein, wherein the cooling fluid flow is a compressed oxidant flow generated by the compressor section of the propulsion system.

[0062] 10. The system according to any one or more clauses herein, the system comprising: a deaerator positioned along the fuel loop to receive the first portion of fuel.

[0063] 11. The system according to any one or more clauses herein, the system comprising: an oxidant flow control device located in the fluid circuit downstream of the compressor section and upstream of the combustion device, wherein the oxidant flow control device is configured to modulate the amount of oxidant flow supplied to the combustion device.

[0064] 12. The system according to any one or more clauses herein, wherein the load device is one or more of a fuel pump, an electric motor, a lubricant pump, a hydraulic pump, an air compressor, an engine starter, a sensor driver, an auxiliary gearbox driver, or a combination thereof.

[0065] 13. The system according to any one or more clauses herein, wherein the combustion device is configured as a deflagration combustion device.

[0066] 14. The system according to any one or more clauses herein, wherein the combustion device is configured as a detonation combustion device.

[0067] 15. The system according to any one or more clauses herein, the system comprising a fuel flow control device at the fuel circuit, wherein the fuel flow control device is configured to modulate the second portion of fuel supplied to the combustion apparatus.

[0068] 16. The system according to any one or more clauses herein, the system comprising: a combustion gas flow control device located in the fluid loop downstream of the combustion device, wherein the combustion gas flow control device divides the combustion gas flow into a first portion of combustion gas in fluid communication with the turbine and a second portion of combustion gas in fluid communication with the heat bus heat exchanger.

[0069] 17. The system according to any one or more clauses herein, wherein the heat bus heat exchanger provides the second portion of combustion gas in thermal communication with the heat transfer fluid at the heat transfer fluid loop.

[0070] 18. The system according to any one or more clauses herein, wherein the second portion of the combustion gas bypasses the turbine.

[0071] 19. The system according to any one or more clauses herein, wherein the combustion gas flow control device is integrated into the turbine as a variable zone turbine nozzle.

[0072] 20. The system according to any one or more clauses herein, the system comprising: a compressor operatively coupled to the turbine via the drive shaft, wherein the compressor is configured to receive the oxidant flow from the compressor section upstream of the combustion device.

[0073] 21. A system for energy conversion, the system comprising: a propulsion system including a compressor section, a heat addition system, and an expansion section arranged in series flow; a fuel circuit fluidly connected from a fuel tank to a fuel flow control device, wherein the fuel flow control device divides a fuel flow from the fuel tank into a first portion of fuel and a second portion of fuel, wherein the fuel circuit is configured to supply the first portion of fuel to the heat addition system; a combustion device configured to receive an oxidant flow directly from the compressor section via the fluid circuit, wherein the fuel circuit is fluidly connected to supply the second portion of fuel to the combustion device, and wherein the combustion device is configured to generate combustion gases from the second portion of fuel and the oxidant flow, and further wherein the fluid circuit is fluidly connected to the combustion device to allow the combustion gases to flow to the propulsion system; a turbine configured to receive the combustion gases from the combustion device via the fluid circuit; and a load device operably coupled to the turbine via a drive shaft, wherein the load device is configured to receive output torque from the drive shaft via the combustion gases through the expansion of the turbine.

[0074] 22. A system for energy conversion, the system comprising: a propulsion system including a compressor section, a heat addition system, and an expansion section arranged in series flow; a fuel circuit fluidly connected from a fuel tank to a fuel flow control device, wherein the fuel flow control device divides a fuel flow from the fuel tank into a first portion of fuel and a second portion of fuel, wherein the fuel circuit is configured to supply the first portion of fuel to the heat addition system; and a combustion device configured to receive an oxidant flow from the compressor section via the fluid circuit, wherein the fuel circuit is fluidly connected to supply the second portion of fuel to the combustion device. The combustion device is configured to generate combustion gases from the second portion of fuel and the oxidant stream, and further wherein the fluid circuit is in fluid communication with the combustion device to allow the combustion gases to flow to the propulsion system; a turbine configured to receive the combustion gases from the combustion device via the fluid circuit; a load device operably coupled to the turbine via a drive shaft, wherein the load device is configured to receive output torque from the drive shaft via the expansion of the turbine through the combustion gases; and a fuel-air heat exchanger positioned along the fuel circuit, the fuel-air heat exchanger being in thermal communication with the first portion of fuel.

Claims

1. A system for energy conversion, characterized in that, The system for energy conversion includes: A propulsion system comprising a compressor section, a heat addition system, and an expansion section arranged in a series flow configuration; A fuel circuit is in fluid communication with a fuel tank and a fuel flow control device, wherein the fuel flow control device divides a fuel flow from the fuel tank into a first portion of fuel and a second portion of fuel, and wherein the fuel circuit is configured to supply the first portion of fuel to the heat addition system. A combustion device configured to receive an oxidant stream from the compressor section via a fluid circuit, wherein a fuel circuit is in fluid communication with the combustion device to provide a second portion of fuel, and wherein the combustion device is configured to generate combustion gases from the second portion of fuel and the oxidant stream, and further wherein the fluid circuit is in fluid communication with the combustion device to allow the combustion gases to flow to the propulsion system; A turbine configured to receive the combustion gases from the combustion device via the fluid circuit; A load device operably coupled to the turbine via a drive shaft, wherein the load device is configured to receive output torque from the drive shaft via the expansion of the combustion gases through the turbine; and A combustion gas flow control device is located in the fluid circuit downstream of the combustion device, wherein the combustion gas flow control device divides the combustion gas into a first portion of combustion gas in fluid communication with the turbine and a second portion of combustion gas in fluid communication with the heat exchanger of the heat bus.

2. The system for energy conversion according to claim 1, characterized in that, The fluid circuit is in fluid communication with the expansion section of the combustion device and the propulsion system, and the fluid circuit provides the combustion gas from the combustion device to the expansion section.

3. The system for energy conversion according to claim 2, characterized in that, The fluid circuit is in fluid communication with the low-pressure turbine in the expansion section of the propulsion system, from the combustion device to the expansion section of the propulsion system.

4. The system for energy conversion according to claim 1, characterized in that, The system for energy conversion includes: A fuel-air heat exchanger, which is positioned along the fuel circuit and is in thermal communication with the first portion of the fuel.

5. The system for energy conversion according to claim 4, characterized in that, The fuel-air heat exchanger is positioned along the fluid loop and is configured to provide heat transfer between the combustion gases and the first portion of fuel.

6. The system for energy conversion according to claim 5, characterized in that, The fuel-air heat exchanger is located downstream of the turbine along the fuel circuit.

7. The system for energy conversion according to claim 1, characterized in that, The combustion device is further configured to operate differently from the heat addition system in order to drive the turbine.

8. The system for energy conversion according to claim 1, characterized in that, The fuel flow control device is configured to control the amount of the second portion of fuel supplied to the combustion device.

9. The system for energy conversion according to claim 1, characterized in that, The system for energy conversion further includes a deaerator positioned along the fuel loop to receive the first portion of fuel.

10. The system for energy conversion according to claim 9, characterized in that, The deaerator receives power from the load device.

11. The system for energy conversion according to claim 1, characterized in that, The system for energy conversion includes: An oxidant flow control device is located in the fluid circuit downstream of the compressor section and upstream of the combustion device, wherein the oxidant flow control device is configured to modulate the amount of oxidant flow supplied to the combustion device.

12. The system for energy conversion according to claim 1, characterized in that, The load device is one or more of a fuel pump, an electric motor, a lubricant pump, a hydraulic pump, an air compressor, an engine starter, a sensor driver, an auxiliary gearbox driver, or a combination thereof.

13. The system for energy conversion according to claim 1, characterized in that, The combustion device is configured as a deflagration combustion device.

14. The system for energy conversion according to claim 1, characterized in that, The combustion device is configured as a detonation combustion device.

15. The system for energy conversion according to claim 1, characterized in that, The fuel flow control device is configured to modulate the second portion of fuel supplied to the combustion device.

16. The system for energy conversion according to claim 1, characterized in that, The heat bus heat exchanger provides the second portion of combustion gas in thermal communication with the heat transfer fluid at the heat transfer fluid loop.

17. The system for energy conversion according to claim 16, characterized in that, The second portion of the combustion gas bypasses the turbine.

18. The system for energy conversion according to claim 1, characterized in that, The combustion gas flow control device is integrated into the turbine as a variable area turbine nozzle.

19. The system for energy conversion according to claim 1, characterized in that, The system for energy conversion includes: A compressor operatively coupled to the turbine via the drive shaft, wherein the compressor is configured to receive the oxidant stream from the compressor section upstream of the combustion device.

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