Gas turbine engine with heat exchanger assembly
Patent Information
- Application Number
- CN202580022932.2
- Authority / Receiving Office
- CN · China
- Patent Type
- Applications(China)
- Current Assignee / Owner
- Priority Date
- 2024-03-28
- Filing Date
- 2025-03-25
- Publication Date
- 2026-10-27
Smart Images
Figure CN122996256A_ABST
Abstract
Description
Priority information
[0001] This application claims priority to U.S. Patent Application Serial No. 18 / 619,703, filed March 28, 2024, the entire contents of which are incorporated herein by reference. Technical Field
[0002] This disclosure relates to a gas turbine engine and a heat exchanger assembly for the gas turbine engine. Background Technology
[0003] Gas turbine engines typically consist of 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 can be configured as a fan assembly, and the fan assembly can be enclosed by an external nacelle. The external nacelle can define a bypass passage between the turbine and the fan assembly.
[0004] Generally, improvements to turbofan engines in the areas of thermal management and aerodynamics are welcomed by the industry. Attached Figure Description
[0005] To those skilled in the art, this specification sets forth the complete and enabling disclosure of the present invention, including its preferred embodiments, and references are made to the accompanying drawings, wherein:
[0006] Figure 1 This is a cross-sectional view of a gas turbine engine according to an exemplary aspect of this disclosure.
[0007] Figure 2 yes Figure 1 A schematic diagram of a portion of a gas turbine engine, including the turbine and core cowling.
[0008] Figure 3 yes Figure 2 A detailed schematic diagram of a part of the turbine.
[0009] Figure 4 yes Figure 2 The front view of the core fairing.
[0010] Figure 5 This is a perspective view of a heat exchanger assembly according to an exemplary embodiment of the present disclosure.
[0011] Figure 6 yes Figure 5 Side view of the heat exchanger assembly.
[0012] Figure 7 yes Figure 5 The main view of the heat exchanger component.
[0013] Figure 8This is a perspective schematic diagram of another heat exchanger assembly according to an exemplary embodiment of the present disclosure.
[0014] Figure 9 This is a perspective view of a heat exchanger assembly according to an exemplary embodiment of the present disclosure. Detailed Implementation
[0015] Reference will now be made in detail to the present embodiments of this disclosure, one or more examples of which are illustrated in the accompanying drawings. Detailed description uses numerical and alphabetic designations to refer to features in the drawings. The same or similar designations have been used in the drawings and description to refer to the same or similar parts of this disclosure.
[0016] The word "exemplary" is used here to mean "serving as an example, instance, or illustration." Any implementation described herein as "exemplary" is not necessarily to be construed as being more preferred or advantageous than other implementations. Furthermore, unless specifically stated otherwise, all embodiments described herein should be considered exemplary.
[0017] Unless the context clearly specifies otherwise, the singular forms “a,” “one,” and “the” include plural references.
[0018] The phrases “from X to Y” and “between X and Y” each refer to a range of values including the endpoints (i.e., a range of values including both X and Y).
[0019] The term "gas turbine engine" refers to an engine that has a turbine as its power source, either entirely or partially. Exemplary gas turbine engines include turbofan engines, turboprop engines, turbojet engines, turboshaft engines, and hybrid electric versions of one or more of these engines.
[0020] The term "combustion section" refers to any heating system used in a turbine. For example, the term "combustion section" can refer to a section including one or more of a knock-off combustion assembly, a rotary detonation combustion assembly, a pulse detonation combustion assembly, or other suitable heating assemblies. In some exemplary embodiments, the combustion section may include an annular burner, a tubular burner, an annular-tube burner, a vortex burner (TVC), or other suitable combustion systems, or combinations thereof.
[0021] Unless otherwise specified, when used with compressors, turbines, shafts, or spool components, the terms “low” and “high” or their respective comparatives (e.g., the –er suffix where applicable) each refer to a relative speed within the engine. For example, “low turbine” or “low-speed turbine” defines a component configured to operate at a speed (e.g., maximum permissible speed) lower than that of the engine’s “high turbine” or “high-speed turbine”.
[0022] The terms "front" and "rear" refer to relative positions within a gas turbine engine or vehicle, and to the normal operating posture of the gas turbine engine or vehicle. For example, in the context of a gas turbine engine, "front" refers to the position near the engine inlet, while "rear" refers to the position near the engine nozzle or exhaust port.
[0023] 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 out, while “downstream” refers to the direction in which the fluid flows in.
[0024] This disclosure generally relates to heat exchanger assemblies for gas turbine engines. As design standards for turbofan engines drive operating conditions toward higher pressures and temperatures, it has been recognized in the art that increased cooling capacity will be desirable. Some gas turbine engines bleed air from, for example, a high-pressure compressor and use this bleed air to cool various accessory systems of the turbofan engine. However, such a configuration can lead to inefficiency because the air bleed from the high-pressure compressor undergoes a relatively large amount of work and is typically at a high temperature.
[0025] Therefore, in one exemplary aspect, a gas turbine engine is provided having a turbine defining a chamber having a chamber inlet in fluid communication with a working gas flow path of the turbine at a location upstream of a compressor section, and a chamber outlet in fluid communication with a bypass passage. A heat exchanger assembly is provided in the chamber, and a heat exchanger of the heat exchanger assembly receives airflow through the chamber. By transferring heat to the airflow, the heat exchanger can be used to cool one or more accessory systems of the gas turbine engine.
[0026] In such an exemplary aspect, the airflow supplied to the heat exchanger may be relatively cold and may be provided from an upstream location where significant compression occurs (significant compression leads to work loss). Utilizing cold air to cool accessory systems improves the operation of a gas turbine engine by requiring fewer additional components.
[0027] Referring now to the accompanying drawings, where the same numbers denote the same elements throughout the drawings, Figure 1 This is a schematic cross-sectional view of a gas turbine engine 10 according to an exemplary embodiment of the present disclosure. More specifically, for Figure 1 In one embodiment, the gas turbine engine 10 is a high bypass ratio turbofan jet engine, sometimes also referred to as a "turbofan engine". For example... Figure 1As shown, the gas turbine engine 10 defines an axial direction A (extending parallel to a longitudinal centerline 12 provided for reference), a radial direction R, and a circumferential direction C extending around the longitudinal centerline 12. Generally, the gas turbine engine 10 includes a fan section 14 and a turbine 16 disposed downstream of the fan section 14, the turbine 16 driving a fan 38 coupled to the fan section 14.
[0028] The depicted exemplary turbine 16 typically includes a generally tubular outer casing 18 defining an annular turbine inlet 20. The casing 18 encloses, in a series flow relationship, a compressor section comprising a supercharger or low-pressure (LP) compressor 22 and a high-pressure (HP) compressor 24; a combustion section 26; a turbine section comprising a high-pressure (HP) turbine 28 and a low-pressure (LP) turbine 30; and an exhaust nozzle section 32. A high-pressure (HP) shaft 34 (which may additionally or alternatively be a spool) drives the HP turbine 28 to the HP compressor 24. A low-pressure (LP) shaft 36 (which may additionally or alternatively be a spool) drives the LP turbine 30 to the LP compressor 22. The compressor section, combustion section 26, turbine section, and exhaust nozzle section 32 collectively define a working gas flow path 37. In this way, it can be understood that the turbine inlet 20 is the inlet to the working gas flow path 37.
[0029] In the illustrated embodiment, the turbine inlet 20 is positioned immediately downstream of the fan 38 (i.e., there are no intervening structures, such as blades, impellers, or struts). Furthermore, it should be understood that in the depicted embodiment, the LP compressor 22 is located downstream of the fan 38, and there is no intermediate compression stage between the fan 38 and the LP compressor 22.
[0030] In the depicted embodiment, fan section 14 includes a fan 38 having a plurality of fan blades 40 spaced apart and coupled to disk 42. As shown, fan 38 is a single-stage fan, and the fan blades 40 generally extend outward from disk 42 in a radial direction R. Each fan blade 40 is rotatable relative to disk 42 about a pitch axis P and is operatively coupled via the fan blades 40 to a suitable pitch-changing mechanism 44 configured to collectively (e.g., uniformly) change the pitch of the fan blades 40. Gas turbine engine 10 further includes a reduction gearbox 46, and the fan blades 40, disk 42, and pitch-changing mechanism 44 collectively rotate about a longitudinal centerline 12 across reduction gearbox 46 via LP shaft 36. Reduction gearbox 46 includes a plurality of gears for adjusting the rotational speed of fan 38 relative to LP shaft 36, enabling fan 38 to rotate at a more efficient fan speed.
[0031] Still referencing Figure 1In an exemplary embodiment, the disc 42 is covered by a rotatable front hub 48 (sometimes referred to as a "rotor") of the fan section 14. The front hub 48 has an aerodynamic profile to facilitate airflow through the multiple fan blades 40.
[0032] Furthermore, the exemplary fan section 14 includes an annular fan housing or outer nacelle 50 circumferentially surrounding at least a portion of the fan 38 and / or turbine 16. It should be understood that, in the depicted embodiment, the outer nacelle 50 is supported relative to the turbine 16 by a plurality of circumferentially spaced outlet guide blades 52. Additionally, a downstream section 54 of the outer nacelle 50 extends through an external portion of the turbine 16, thereby defining a bypass passage 56 therebetween. The bypass passage 56 is at least partially defined above the turbine 16.
[0033] During operation of the gas turbine engine 10, a volume of air 58 enters the gas turbine engine 10 through the associated inlet 60 of the outer nacelle 50 and the fan section 14. As the volume of air 58 passes through the fan blades 40, a first portion 62 of the air 58 is directed or routed to the bypass passage 56, and a second portion 64 of the air is directed or routed to the working gas flow path 37, or more specifically, into the LP compressor 22. The ratio between the first portion 62 and the second portion 64 of the air 58 is commonly referred to as the bypass ratio. Subsequently, the second portion 64 of the air 58 is routed through the HP compressor 24 and enters the combustion section 26 at increased pressure, where it mixes with fuel and is burned to provide combustion gases 66.
[0034] Combustion gas 66 is routed through HP turbine 28, where a portion of the thermal and / or kinetic energy from combustion gas 66 is extracted via a sequential stage of HP turbine stator blades 68 connected to housing 18 and HP turbine rotor blades 70 connected to HP shaft 34, thereby causing HP shaft 34 to rotate to support the operation of HP compressor 24. Combustion gas 66 is then routed through LP turbine 30, where a second portion of the thermal and kinetic energy from combustion gas 66 is extracted via a sequential stage of LP turbine stator blades 72 connected to housing 18 and LP turbine rotor blades 74 connected to LP shaft 36, thereby causing LP shaft 36 to rotate to support the operation of LP compressor 22 and / or fan 38.
[0035] Combustion gas 66 is then routed through the injector nozzle section 32 of turbine 16 to provide propulsive thrust. Simultaneously, the pressure of the first portion 62 of air 58 increases significantly as it is routed through bypass passage 56 before exiting from fan nozzle exhaust section 76 of gas turbine engine 10, which also provides propulsive thrust. HP turbine 28, LP turbine 30, and injector nozzle section 32 at least partially define a hot gas path 78 for routing combustion gas 66 through turbine 16.
[0036] However, it should be understood that Figure 1 The exemplary gas turbine engine 10 depicted is merely an example, and in other exemplary embodiments, the gas turbine engine 10 may have any other suitable configuration. For example, while the depicted gas turbine engine 10 is configured as a ducted gas turbine engine (i.e., including the outer nacelle 50), in other embodiments, the gas turbine engine 10 may be a non-ducted gas turbine engine (such that the fan 38 is a non-ducted fan, and the outlet guide vane 52 cantilevered out from the housing 18). Additionally or alternatively, while the depicted gas turbine engine 10 is configured as a geared gas turbine engine (i.e., including a reduction gearbox 46) and a variable-pitch gas turbine engine (i.e., including a fan 38 configured as a variable-pitch fan), in other embodiments, the gas turbine engine 10 may additionally or alternatively be configured as a direct-drive gas turbine engine (such that the LP shaft 36 rotates at the same speed as the fan 38), a fixed-pitch gas turbine engine (such that the fan 38 includes fan blades 40 that cannot rotate about the pitch axis P), or both. It should also be understood that, in other exemplary embodiments, aspects of this disclosure may be incorporated into any other suitable gas turbine engine. For example, in other exemplary embodiments, aspects of this disclosure may (as appropriate) be incorporated into a turboprop gas turbine engine, a turboshaft gas turbine engine, or a turbojet gas turbine engine.
[0037] Now for reference Figure 2 , showed Figure 1 A close-up schematic diagram of a portion of an exemplary gas turbine engine 10. In particular, Figure 2 The view is Figure 1 A close-up view of the turbine 16 depicts the turbine inlet 20, the compressor section containing the LP compressor 22 and the HP compressor 24, the LP shaft 36, and the reduction gearbox 46. Furthermore, the outer casing 18 of the turbine 16 is depicted extending around at least a portion of the compressor section, with a bypass passage 56 partially defined thereby.
[0038] from Figure 2As will be understood from the view, turbine 16 also includes a frame 100. Frame 100 includes a strut 102 extending through the working gas flow path 37 upstream of the LP compressor 22 and downstream of the turbine inlet 20, and a strut 104 extending through the working gas flow path 37 downstream of the LP compressor 22 and upstream of the HP compressor 24. Frame 100 provides structural support for the various components of the gas turbine engine 10.
[0039] The frame 100 defines a chamber 106 extending radially outward from the LP compressor 22. The chamber 106 includes a flow path inlet 108 in fluid communication with a bypass passage 56, and a flow path outlet 110 also in fluid communication with the bypass passage 56. Cool air from the bypass passage 56 can flow into the chamber 106 through the flow path inlet 108 and out through the flow path outlet 110 back to the bypass passage 56, thus defining a cooling air flow path.
[0040] The gas turbine engine 10 includes a core cowling 112. The core cowling 112 is part of the gas turbine engine 10 that surrounds the core airflow. The core airflow flows through the HP compressor 24 to the HP turbine 28. The frame 100 includes a connector 114 that engages the core cowling 112. The connector 114 may define a circumferential recess 116, and the core cowling 112 may include a wedge 118 mounted in the circumferential recess 116 to secure the core cowling 112 to the turbine 16. Figure 2 In an exemplary embodiment, the circumferential groove 116 is V-shaped, and it should be understood that the circumferential groove 116 may have any suitable shape.
[0041] The gas turbine engine 10 includes a supercharger 120. The supercharger 120 provides additional power to the air in the working gas flow path 37. Figure 2 In this configuration, the booster 120 includes an LP compressor 22. Generally, the booster 120 increases one or more of the velocity, pressure, or temperature of the air in the working gas flow path 37. Thus, the booster 120 "pressurizes" the air toward the HP compressor 24.
[0042] The turbine 16 includes a turbocharger cowl 122. The turbocharger cowl 122 is disposed in the chamber 106 and extends around the turbocharger 120 in the circumferential direction C. Generally, the turbocharger cowl 122 may be shaped according to the profile of the frame 100 adjacent to the LP compressor 22.
[0043] The gas turbine engine 10 includes a heat exchanger assembly 124 disposed in a cooling airflow path. The heat exchanger assembly 124 includes a heat exchanger 126, a diffuser section 128, and an exhaust section 130. The heat exchanger assembly 124 utilizes cooling air from a bypass passage 56 to extract heat from other components of the gas turbine engine 10, including one or more accessory systems such as an oil lubrication system, a fuel delivery system, a cooled air conditioning (CCA) system, or an engine controller cooling system. The heat exchanger assembly 124 then exhausts the heated air into the bypass passage 56.
[0044] Heat exchanger assembly 124 is positioned within gas turbine engine 10 to address space constraints in the axial direction A. Heat exchanger assembly 124 is positioned axially in the A direction rear of fan 38 and axially in the A direction forward of core cowling 112. Specifically, heat exchanger assembly 124 is positioned axially in the A direction rear of outlet guide vane 52 and axially in the A direction forward of connector 114. By positioning heat exchanger assembly 124 between outlet guide vane 52 and core cowling 112, heat exchanger assembly 124 can utilize the space within chamber 106 to collect cool air that has not been compressed by turbocharger 120 (thus increasing air temperature) from working gas flow path 37.
[0045] The heat exchanger assembly 124 is disposed in the gas turbine engine 10 to address space constraints in the radial direction R. For example... Figure 2 As shown, the heat exchanger assembly 124 is disposed on the outside of the turbocharger 120 in the radial direction R. Specifically, the heat exchanger assembly 124 is disposed in the chamber 106, and the turbocharger shroud 122 is disposed between the LP compressor 22 and the heat exchanger assembly 124. By positioning the heat exchanger assembly 124 on the outside of the turbocharger shroud 122 adjacent to the bypass passage 56, the heat exchanger assembly 124 can directly discharge heated air through the flow path outlet 110 into the bypass passage 56, thereby improving the cooling of the accessory system by the heat exchanger assembly 124.
[0046] Heat exchanger assembly 124 includes heat exchanger 126. Heat exchanger 126 includes inlet 132 and outlet 134. Inlet 132 receives cold air from a cooling air flow path, and outlet 134 discharges heated air into a chamber. Heat exchanger 126 can be of any suitable type, such as an air-cooled oil cooler, a variable frequency generator air-cooled oil cooler, an air-to-air heat exchanger, or an air-cooled fluid cooler. That is, heat exchanger 126 can utilize the air as the working fluid for absorbing heat from the hot oil. The cooled oil then leaves the heat exchanger and returns to other components.
[0047] The heat exchanger assembly 124 includes a diffuser section 128. The diffuser section 128 is in fluid communication with the inlet 132 of the heat exchanger 126. The diffuser section 128 is configured to reduce the air velocity flowing towards the inlet 132 within the chamber 106, thereby reducing the pressure drop across the heat exchanger 126 and improving the efficiency of the heat exchanger 126. The diffuser section 128 is in airflow communication with the bypass passage 56 at the flow path inlet 108.
[0048] The heat exchanger assembly 124 includes an exhaust section 130. The exhaust section 130 is fluidly connected to an outlet 134 and a bypass passage 56. That is, the exhaust section 130 directs air heated by the heat exchanger 126 to the bypass passage 56. The exhaust section 130 is configured to accelerate (e.g., by compression) the air flowing from the outlet 134 to the bypass passage 56, further increasing the temperature of the air. A booster 120 may be in fluid communication with the exhaust section 130, and the exhaust section 130 may be configured to accelerate the air from the outlet 134 using an airflow from the booster 120. More specifically, the booster 120 may provide an airflow 136 that, when compressed together with the air from the heat exchanger 126, can assist in driving the air through the flow path outlet 110 into the bypass passage 56. Therefore, the pressure in the exhaust section 130 may decrease, which may drive more air into the inlet 132 of the heat exchanger 126, thereby improving the operation of the heat exchanger 126.
[0049] The heat exchanger assembly 124 includes a heat exchanger shroud 138 disposed radially in the direction R outside the heat exchanger 126, the diffuser section 128, and the exhaust section 130. The heat exchanger shroud 138 provides an outward exposure to the bypass passage 56, thereby defining a cooling airflow path below it. The heat exchanger shroud 138 has a length L1 in the axial direction A, while the turbine 16 defines a length L2 in the axial direction A from the turbine inlet 20 to the core shroud 112. The length L1 of the heat exchanger shroud 138 can be determined as a specific percentage of the length L2, for example, from 10% to 60%. That is, in one form, the length L1 of the heat exchanger shroud can be... .
[0050] The gas turbine engine 10 may include a variable bleed assembly 140. The variable bleed assembly 140 is disposed axially in the direction A downstream of the exhaust section 130. The variable bleed assembly 140 may be in fluid communication with the exhaust section 130, thereby directing air in the downstream working gas flow path 37 of the LP compressor 22 to a bypass passage 56. More specifically, the variable bleed assembly 140 includes a variable bleed inlet 142 for receiving air from the working gas flow path 37 and a variable bleed outlet 144 for supplying that air to the exhaust section 130. The variable bleed assembly 140 may also further accelerate the air in the exhaust section 130 before it exits into the bypass passage 56.
[0051] Exhaust section 130 may include injector 146. Injector 146 propels cooling air in exhaust section 130 through flow path outlet 110 into bypass passage 56. More specifically, injector 146 increases the pressure of cooling air to drive it out of heat exchanger 126 and into bypass passage 56.
[0052] Now refer to Figure 3 The image shows an enlarged view of the injector 146. The variable vent outlet 144 includes a fluid nozzle 148 configured to provide a relatively high-pressure fluid flow. Furthermore, the injector 146 includes a nozzle portion 150 comprising a converging inlet nozzle 152, a diffuser throat 154, and a diverging outlet diffuser 156 arranged in series, with the fluid nozzle 148 of the variable vent outlet 144 oriented to provide relatively high-pressure fluid to the converging inlet nozzle 152. It should be understood that the variable vent assembly 140 acts as a high-pressure air source providing a high-pressure fluid flow through the fluid nozzle 148, with the nozzle portion 150 located downstream. Such a high-pressure fluid flow can induce a relatively low-pressure fluid flow through it, thereby increasing the airflow through the exhaust section 130 to the bypass passage 56.
[0053] Reference Figure 4 A schematic diagram of a core fairing 112 is provided. The core fairing 112 includes a first member 160 and a second member 162. The first member 160 and the second member 162 extend approximately 360° together when joined. For example, to mount the core fairing 112 into a circumferential recess 116 of the frame 100 (see...),... Figure 2 The first component 160 and the second component 162 of the core fairing 112 can be installed separately (e.g., Figure 4 (As shown by the dashed line in the image).
[0054] In this context, it can be understood that the first member 160 and the second member 162 may jointly define a hinge connection 164 on one side (e.g., any mechanical connection that allows the first member to pivot relative to the second member, such as a pin connection) and define a releasable mechanical connection 166 on the opposite side (i.e., any mechanical connection that allows release in normal operation without damaging the connection, such as a bolt connection, ratchet connection, latch connection, etc.).
[0055] However, it should be understood that in other exemplary embodiments, the core fairing 112 may be configured in any other suitable manner. For example, in other embodiments, the core fairing 112 may utilize any other suitable mechanical connection, such as bolts, screws, rivets, welding, complementary geometries (e.g., dovetail grooves), etc.
[0056] Now refer to Figure 5-7 The image shown is an enlarged view of the heat exchanger assembly 124. Figure 5 This is a perspective view of heat exchanger assembly 124. Figure 6 This is a side view of the heat exchanger assembly 124. Figure 7 This is a front view of the heat exchanger assembly 124.
[0057] As shown most clearly Figure 5 In the heat exchanger assembly 124, there are multiple heat exchangers 126. Each of the multiple heat exchangers 126 includes an inlet 132 and an outlet 134 (for clarity, in...) Figure 3-5 (Only one is marked in the text). Each heat exchanger 126 receives cold air at inlet 132 and discharges heated air from outlet 134. Figure 3 Six heat exchangers 126A, 126B, 126C, 126D, 126E, and 126F (collectively, “heat exchangers 126”) are shown, but it should be understood that the heat exchanger assembly 124 may include a different number of heat exchangers 126, ranging from at least two to up to twenty.
[0058] As shown most clearly Figure 6-7 In the chamber 106, multiple heat exchangers 126 are arranged around the turbine 16 along the circumferential direction C. Figure 2 Specifically, multiple heat exchangers 126 are arranged around the turbocharger shroud 122. The arrangement of multiple heat exchangers 126 around the turbocharger shroud 122 allows for cooling of the accessory system at different locations in the circumferential direction C.
[0059] Back Figure 5The heat exchanger assembly 124 includes a plurality of diffuser sections 128, of which only one is labeled for clarity. Each of the plurality of heat exchangers 126 is in fluid communication with at least one of the plurality of diffuser sections 128. More specifically, the inlet 132 of each of the plurality of heat exchangers 126 receives air diffused by at least one of the plurality of diffuser sections 128. The diffuser sections 128 may be arranged in the circumferential direction C such that each diffuser section 128 is in fluid connection to one of the plurality of heat exchangers 126. In such a configuration, the diffuser section 128 receives air flowing through a cooling air flow path.
[0060] The heat exchanger assembly 124 includes a plurality of exhaust sections 130, only one of which is labeled for clarity. Each of the plurality of heat exchangers 126 is in fluid communication with at least one of the plurality of exhaust sections 130. More specifically, the outlet 134 of each of the plurality of heat exchangers 126 transmits air heated by the heat exchangers 126 to the exhaust section 130. The exhaust sections 130 may be arranged in the circumferential direction C such that each exhaust section 130 is in fluid connection to one of the plurality of heat exchangers 126. Each exhaust section 130 is in fluid communication with a bypass passage 56 to discharge the heated air.
[0061] like Figure 7 As shown, heat exchanger assembly 124 includes at least one heat source 170 in thermal communication with heat exchanger 126. The heat source 170 supplies a heating fluid (such as oil) to heat exchanger 126 to transfer heat from the heating fluid to the air flowing through heat exchanger 126. Heat source 170 may be one of the accessory systems described above, or an intermediate system that directs unheated working fluid to an accessory system to receive heat, and then directs heated working fluid to heat exchanger 126 to dissipate heat. Heat exchanger assembly 124 may include more than one heat source 170 to accommodate space constraints and location of accessory systems. For example, as... Figure 7 As shown, the heat exchanger assembly 124 may include three heat sources 170A, 170B, and 170C.
[0062] Figure 5-7 The heat exchanger assembly shown includes multiple heat exchangers 126, each with its corresponding individual diffuser section 128. However, it should be understood that heat exchangers with diffuser sections of different configurations and shapes can be implemented. As an example, Figure 8 A perspective view of a heat exchanger assembly 180 is shown, which has a single diffuser section 182 extending in a ring shape in the circumferential direction. This diffuser section 182 surrounds the gas turbine engine 10 (… Figure 1 The longitudinal centerline 12 extends 360 degrees. The single diffuser section 182 supplies air to each heat exchanger 126, distributing the air separately to each discrete heat exchanger 126.
[0063] As another example, Figure 9 A perspective view of a heat exchanger assembly 190 is shown, which has a surrounding gas turbine engine 10 ( Figure 1 The longitudinal centerline 12 extends 360 degrees in the circumferential direction to form a single heat exchanger 192. Discrete diffuser sections 128 supply air to the heat exchanger 192, and the air flows throughout the heat exchanger 192 to transfer heat to the accessory system.
[0064] Further aspects are provided by the following topics:
[0065] A gas turbine engine, defined in a radial and axial direction, includes a fan assembly comprising a fan; a turbine driven coupled to the fan and including a compressor section, a combustion section, and a turbine section arranged in a series flow sequence and partially defining a working gas flow path; a bypass passage defined above the turbine; a core cowling; and a heat exchanger assembly disposed axially behind the fan and in front of the core cowling, the heat exchanger assembly including a heat exchanger and a heat exchanger cowling, the heat exchanger cowling defining a cooling air flow path extending between a flow path inlet and a flow path outlet, the flow path inlet communicating with the bypass passage airflow to receive cooling airflow from the bypass passage, the flow path outlet communicating with the bypass passage airflow to discharge the cooling airflow back to the bypass passage, the heat exchanger positioned within the cooling air flow path, the cooling air flow path including a diffuser section located between the flow path inlet and the heat exchanger.
[0066] According to any of the preceding clauses, the gas turbine engine wherein the cooling air flow path further defines an exhaust section located between the heat exchanger and the flow path outlet.
[0067] The gas turbine engine according to any of the preceding clauses, wherein the turbine includes an injector positioned in communication with the exhaust gas flow, wherein the injector is in communication with a high-pressure air source flow.
[0068] According to any of the preceding clauses, the gas turbine engine, wherein the compressor section includes a low-pressure compressor and a high-pressure compressor, wherein the turbine further includes a variable bleed assembly disposed downstream of the low-pressure compressor and upstream of the high-pressure compressor, wherein the high-pressure air source is the variable bleed assembly.
[0069] According to any of the preceding clauses, in a gas turbine engine, the compressor section of the turbine includes a turbocharger, wherein the heat exchanger is disposed outside the turbocharger in the radial direction.
[0070] The gas turbine engine according to any of the preceding clauses further includes a turbocharger cowling, wherein the turbocharger cowling is disposed in the radial direction between the turbocharger and the heat exchanger assembly.
[0071] According to any of the preceding clauses, in a gas turbine engine, the fan further includes an outlet guide vane, and the heat exchanger assembly is disposed behind the outlet guide vane in the axial direction.
[0072] According to any of the preceding clauses, the gas turbine engine includes a circumferential groove, the core cowling includes a first member, a second member and a hinged connection therebetween, wherein the first member and the second member are coupled to the circumferential groove, and wherein the heat exchanger assembly is disposed in front of the circumferential groove in the axial direction.
[0073] The gas turbine engine according to any of the preceding clauses, wherein the heat exchanger is the first heat exchanger among a plurality of heat exchangers in the heat exchanger assembly.
[0074] According to any of the preceding clauses, in the gas turbine engine, the heat exchanger shroud is a first heat exchanger shroud among a plurality of heat exchanger shrouds of the heat exchanger assembly, the cooling air flow path is a first cooling air flow path among a plurality of cooling air flow paths of the heat exchanger assembly, and each of the plurality of cooling air flow paths is defined by a corresponding one of the plurality of heat exchanger shrouds.
[0075] The gas turbine engine according to any of the preceding clauses further defines a circumferential direction, and wherein the plurality of heat exchanger shrouds and the plurality of heat exchangers are arranged in the circumferential direction.
[0076] The gas turbine engine according to any of the preceding clauses further includes a heat source in thermal communication with the plurality of heat exchangers.
[0077] The gas turbine engine according to any of the preceding clauses, wherein the plurality of heat exchangers comprises at least two heat exchangers and less than or equal to 20 heat exchangers.
[0078] The gas turbine engine according to any of the preceding clauses, wherein the heat exchanger is one of an air-cooled oil cooler, a variable frequency generator air-cooled oil cooler, an air-to-air heat exchanger, or an air-cooled fluid cooler.
[0079] The gas turbine engine according to any of the preceding clauses, wherein the heat exchanger assembly further includes a heat source in thermal communication with the heat exchanger.
[0080] According to any of the preceding clauses, in a gas turbine engine, wherein the diffuser section is configured to reduce the velocity of the air flowing toward the inlet.
[0081] According to any of the preceding clauses, in a gas turbine engine, wherein the cooling air flow path further defines an exhaust section located between the heat exchanger and the flow path outlet, wherein the exhaust section is configured to accelerate the cooling airflow from the heat exchanger provided to the bypass passage.
[0082] The gas turbine engine according to any of the preceding clauses, wherein the fan is a single-stage fan.
[0083] According to any of the preceding clauses, the gas turbine engine defines an inlet to the working gas flow path and includes a circumferential recess for mounting at least a portion of the core shroud, wherein the heat exchanger shroud defines a first length along the axial direction from the flow path inlet to the flow path outlet, wherein the turbine defines a second length along the axial direction from the inlet to the circumferential recess, and wherein L1 is greater than or equal to 10% of L2 and less than or equal to 60% of L2.
[0084] The gas turbine engine according to any of the preceding clauses, wherein L1 is greater than or equal to 20% of L2 and less than or equal to 50% of L2.
[0085] According to any of the preceding clauses, the gas turbine engine wherein the diffuser section extends 360 degrees around the longitudinal centerline of the gas turbine engine.
[0086] According to any of the preceding clauses, the gas turbine engine wherein the heat exchanger extends 360 degrees around the longitudinal centerline of the gas turbine engine.
[0087] A heat exchanger assembly is disposed behind a fan in the axial direction and in front of a core shroud in the same axial direction. The heat exchanger assembly includes a heat exchanger and a heat exchanger shroud. The heat exchanger shroud defines a cooling airflow path extending between a flow path inlet and a flow path outlet. The flow path inlet is in airflow communication with a bypass channel to receive cooling airflow from the bypass channel. The flow path outlet is in airflow communication with the bypass channel to discharge the cooling airflow back to the bypass channel. The heat exchanger is positioned within the cooling airflow path, which includes a diffuser section located between the flow path inlet and the heat exchanger.
[0088] This written description uses examples to disclose this disclosure, including best practices, and also enables any person skilled in the art to practice this disclosure, including making and using any device or system and performing any combined methods. The patent scope of this disclosure is defined by the claims and may include other examples that would occur to a person skilled in the art. Such other examples are intended to fall within the scope of the claims if they contain structural elements that are not different from the literal language of the claims, or if they contain equivalent structural elements that differ only slightly from the literal language of the claims.
Claims
1. A gas turbine engine (10), defining a radial direction and an axial direction, characterized in that, The gas turbine engine (10) includes: A fan assembly (140) comprising a fan; A turbine (16) is driven to the fan and includes a compressor section, a combustion section and a turbine section arranged in series flow sequence and partially defining a working gas flow path (37), and the gas turbine engine (10) defines a bypass passage (56) above the turbine (16). Core fairing (112); and A heat exchanger assembly (124) is disposed behind the fan in the axial direction and in front of the core shroud (112) in the axial direction. The heat exchanger assembly (124) includes: Heat exchanger (126); and A heat exchanger shroud (138) defines a cooling airflow path extending between a flow path inlet (108) and a flow path outlet (110), the flow path inlet (108) being in airflow communication with the bypass channel (56) to receive cooling airflow from the bypass channel (56), and the flow path outlet (110) being in airflow communication with the bypass channel (56) to discharge the cooling airflow back to the bypass channel (56), the heat exchanger (126) being positioned within the cooling airflow path, the cooling airflow path including a diffuser section (128) located between the flow path inlet (108) and the heat exchanger (126).
2. The gas turbine engine (10) according to claim 1, characterized in that, The cooling air flow path further defines an exhaust section (130) located between the heat exchanger (126) and the flow path outlet (110).
3. The gas turbine engine (10) according to claim 2, characterized in that, The turbine (16) includes an injector (146) positioned in communication with the airflow of the exhaust section (130), wherein the injector (146) is in communication with the airflow of a high-pressure air source.
4. The gas turbine engine (10) according to claim 3, characterized in that, The compressor section includes a low-pressure compressor and a high-pressure compressor, and the turbine (16) further includes a variable venting assembly (140) disposed downstream of the low-pressure compressor and upstream of the high-pressure compressor, wherein the high-pressure air source is the variable venting assembly (140).
5. The gas turbine engine (10) according to claim 1, characterized in that, The compressor section of the turbine (16) includes a turbocharger (120), wherein the heat exchanger (126) is disposed on the outside of the turbocharger (120) in the radial direction.
6. The gas turbine engine (10) according to claim 5, characterized in that, It further includes a turbocharger (120) fairing, wherein the turbocharger (120) fairing is disposed in the radial direction between the turbocharger (120) and the heat exchanger assembly (124).
7. The gas turbine engine (10) according to claim 1, characterized in that, The fan further includes an outlet guide vane (52), and the heat exchanger assembly (124) is disposed behind the outlet guide vane (52) in the axial direction.
8. The gas turbine engine (10) according to claim 1, characterized in that, The turbine (16) includes a circumferential groove (116), the core fairing (112) includes a first member (160), a second member (162) and a hinge connection (164) therebetween, wherein the first member (160) and the second member (162) are connected to the circumferential groove (116), and wherein the heat exchanger assembly (124) is disposed in front of the circumferential groove (116) in the axial direction.
9. The gas turbine engine (10) according to claim 1, characterized in that, The heat exchanger (126) therein is the first heat exchanger (126) of a plurality of heat exchangers (126) of the heat exchanger assembly (124).
10. The gas turbine engine (10) according to claim 9, characterized in that, The heat exchanger shroud (138) is the first heat exchanger shroud (138) of a plurality of heat exchanger shrouds (138) of the heat exchanger assembly (124), the cooling airflow path is the first cooling airflow path of a plurality of cooling airflow paths of the heat exchanger assembly (124), and each of the plurality of cooling airflow paths is defined by a corresponding one of the plurality of heat exchanger shrouds (138).
11. The gas turbine engine (10) according to claim 10, characterized in that, The circumferential direction is further defined, and the plurality of heat exchanger shrouds (138) and the plurality of heat exchangers (126) are arranged in the circumferential direction.
12. The gas turbine engine (10) according to claim 9, characterized in that, It further includes a heat source (170) in thermal communication with the plurality of heat exchangers (126).
13. The gas turbine engine (10) according to claim 9, characterized in that, The plurality of heat exchangers (126) includes at least two heat exchangers (126) and less than or equal to 20 heat exchangers (126).
14. The gas turbine engine (10) according to claim 1, characterized in that, The cooling air flow path further defines an exhaust section (130) between the heat exchanger (126) and the flow path outlet (110), wherein the exhaust section (130) is configured to compress the cooling airflow from the heat exchanger (126) provided to the bypass passage (56).
15. The gas turbine engine (10) according to claim 1, characterized in that, The turbine (16) defines an inlet (60) leading to the working gas flow path (37) and includes a circumferential groove (116) for mounting at least a portion of the core shroud (112), wherein the heat exchanger shroud (138) defines a first length along the axial direction from the flow path inlet (108) to the flow path outlet (110), wherein the turbine (16) defines a second length along the axial direction from the inlet (60) to the circumferential groove (116), and wherein L1 is greater than or equal to 10% of L2 and less than or equal to 60% of L2.