Gas turbine engine with fluid circuit and ejector

Abstract

A gas turbine engine is provided having a static structure including a flow path wall. A fluid circuit extends through the flow path wall and includes a first inlet opening in fluid communication with a first cavity to receive a first fluid flow through the fluid circuit. The static structure includes an injector positioned at the fluid circuit, wherein the injector includes a second inlet opening in fluid communication with a second cavity to receive a second fluid flow passing through the injector and entering the fluid circuit.

Description

[0001] Government-funded research

[0002] The project for which this application was submitted has been funded by the EU Clean Sky 2 Research and Innovation Program, with funding agreement number CS2-ENG-GAM-2014-2015-01. Technical Field

[0003] This topic generally concerns cooling structures used in gas turbine engines. Background Technology

[0004] Gas turbine engines produce high-temperature gases that come into thermal contact with components through the gas flow path. These high-temperature gases wear down and degrade gas turbine engine components, and sometimes the high temperature may exceed the melting point or other critical temperature of some components along the gas flow path. Gas turbine engines generally include cooling circuits and structures to reduce component temperatures, thereby mitigating the wear and degradation caused by the high-temperature gases.

[0005] This type of cooling circuit typically removes relatively cool air from the compressor and directs it to other components, such as the combustor and turbine section, to provide the desired cooling. It utilizes compressed air, especially high-pressure, high-energy compressed air from the compressor section, to remove and bypass input energy that would otherwise flow to the combustion process, and uses the compressed air for cooling purposes. Therefore, this method and structure for cooling compromises the engine's thermodynamic performance and efficiency in terms of structural durability and component lifespan.

[0006] Therefore, improved cooling structures are needed for gas turbine engines. Furthermore, improved cooling structures are required to reduce losses associated with utilizing relatively high-pressure air. Summary of the Invention

[0007] The aspects and advantages of this disclosure 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 this disclosure.

[0008] One aspect of the invention relates to a gas turbine engine having a blade assembly including a flow path wall. A fluid circuit extends through the flow path wall. The fluid circuit defines a first inlet opening in fluid communication with a first cavity to receive a first fluid flow through the fluid circuit. The blade assembly includes an injector positioned at the fluid circuit. The injector defines a second inlet opening in fluid communication with a second cavity to receive a second fluid flow through the injector and into the fluid circuit.

[0009] Another aspect of this disclosure relates to a static structure for a gas turbine engine. The static structure includes a flow path wall having a fluid loop extending through the flow path wall. The fluid loop includes a first inlet opening in fluid communication with a first cavity to receive a first fluid flow through the fluid loop. The static structure includes an injector positioned at the fluid loop. The injector includes a second inlet opening in fluid communication with a second cavity to receive a second fluid flow passing through the injector and entering the fluid loop.

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

[0011] The specification with reference to the accompanying drawings sets forth a complete and feasible disclosure for those skilled in the art, including its best mode, in which:

[0012] Figure 1 These are illustrative embodiments of exemplary engines based on various aspects of this disclosure;

[0013] Figure 2 These are illustrative embodiments of exemplary combustion sections for an engine based on various aspects of this disclosure;

[0014] Figure 3 This is a perspective view of a portion of a static structure having a fluid circuit according to various aspects of this disclosure;

[0015] Figure 4 This is a perspective view of an embodiment of a static structure having a fluid circuit according to various aspects of this disclosure; and

[0016] Figure 5 This is a perspective view of an embodiment of a static structure having a fluid circuit according to various aspects of this disclosure; and

[0017] Figure 6 This is a cross-sectional view of an embodiment of an ejector in a fluid circuit according to various aspects of this disclosure;

[0018] Figure 7 This is a cross-sectional view of an embodiment of an ejector in a fluid circuit according to various aspects of this disclosure;

[0019] Figure 8 This is a cross-sectional view of an embodiment of a static structure having a fluid circuit according to various aspects of this disclosure;

[0020] Figure 9 This is a cross-sectional view of an embodiment of a static structure having a fluid circuit according to various aspects of this disclosure;

[0021] Figure 10 This is a cross-sectional view of an embodiment of a static structure having a fluid circuit according to various aspects of this disclosure;

[0022] Figure 11 This is a cross-sectional view of an embodiment of a static structure having a fluid circuit according to various aspects of this disclosure;

[0023] Figure 12 This is a radial view of an embodiment of a static structure having a fluid circuit according to various aspects of this disclosure;

[0024] Figure 13 This is a radial view of an embodiment of a static structure having a fluid circuit according to various aspects of this disclosure; and

[0025] Figure 14 This is a cross-sectional view illustrating an embodiment of a fluid circuit relative to one or more surfaces of a static structure according to various aspects of this disclosure.

[0026] Reference characters are used repeatedly in this specification and drawings to indicate the same or similar features or elements of this disclosure. Detailed Implementation

[0027] Reference will now be made in detail to embodiments of the present disclosure, one or more examples of which are illustrated in the accompanying drawings. Each example is provided by way of explanation and not limitation of the disclosure. In fact, it will be apparent to those skilled in the art that various modifications and variations can be made to the disclosure without departing from the scope or spirit of the disclosure. For example, features 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 these modifications and variations within the scope of the appended claims and their equivalents.

[0028] The term "exemplary" is used herein to mean "serving as an example, instance, or illustration." Any implementation described herein as "exemplary" is not necessarily to be construed as being more preferred or advantageous than other implementations. Furthermore, unless otherwise specifically indicated, all embodiments described herein should be considered exemplary.

[0029] As used in this application, 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 a single component.

[0030] Furthermore, 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, while "downstream" refers to the direction from which the fluid flows.

[0031] 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 closer to the engine inlet, and "rear" refers to the position closer to the engine nozzle or exhaust.

[0032] Unless the context clearly specifies otherwise, the singular forms “a,” “an,” and “the” include plural references.

[0033] This document provides embodiments of cooling structures for gas turbine engines that reduce losses associated with utilizing relatively high-pressure air. The structures and methods depicted and described herein include a gas turbine engine with a fluid circuit in which an injector is formed. The injector is formed with a static structure, such as a housing, frame, or blade assembly, or particularly within an inner band, outer band, or airfoil structure, forming a blade assembly. The fluid circuit has a first inlet opening in fluid communication with a first chamber of relatively low pressure and a second inlet opening at the injector in fluid communication with a second chamber of relatively high pressure. The injector entrains or draws a low-pressure fluid flow from the first chamber into the fluid circuit via a relatively high-pressure fluid flow from the second chamber. The first chamber may include a lower shroud area, a fan housing, a bypass flow path, or other chambers with a large flow rate of low-pressure air (e.g., atmospheric pressure). The second chamber may include a cooling circuit, such as a secondary cooling circuit outside the main compressed air or combustion gas flow path. In this way, it is understood that, as used herein, the term "chamber" refers to any air source and does not necessarily require a complete or substantially complete housing.

[0034] In contrast to all or substantially all of the cooling air from the compressor section, the ejector allows a relatively large amount of low-temperature air to be drawn through the fluid loop by a relatively small amount of relatively high-temperature air (such as from the compressor section).

[0035] The embodiments of gas turbine engines with fluid circuits and injectors described herein allow for improved cooling or thermal decay while reducing the amount or quantity of air removed from the compressor section and from the combustion process. Therefore, the embodiments herein allow for improved engine and combustion efficiency while maintaining or improving the cooling of conventional cooling structures.

[0036] Now refer to the attached diagram, Figure 1 This is a schematic cross-sectional view of one embodiment of the gas turbine engine 10. In the illustrated embodiment, the engine 10 is configured as a turbofan engine. However, in alternative embodiments, the engine 10 may be configured as a propeller fan or open rotor engine, a turbojet engine, a turboprop engine, a turboshaft gas turbine engine, or any other suitable type of gas turbine engine.

[0037] like Figure 1 As shown, engine 10 is defined by longitudinal direction L, radial direction R, and circumferential direction C. Generally, longitudinal direction L extends parallel to longitudinal centerline 12 of engine 10, radial direction R extends orthogonally outward from longitudinal centerline 12, and circumferential direction C extends substantially concentrically around longitudinal centerline 12.

[0038] Generally, the engine 10 includes a fan section 14, a low-pressure (LP) spool 16, and a high-pressure (HP) spool 18, all at least partially surrounded by an annular nacelle 20. More specifically, the fan section 14 may include a fan rotor 22 and a plurality of fan blades 24 (one shown) coupled to the fan rotor 22. In this respect, the fan blades 24 are spaced apart from each other along the circumferential direction C and extend radially outward from the fan rotor 22. Furthermore, the LP spool 16 and the HP spool 18 are positioned downstream of the fan section 14 along a longitudinal centerline 12 (i.e., along the longitudinal direction L). As shown, the LP spool 16 is rotatably coupled to the fan rotor 22, thereby allowing the LP spool 16 to rotate the fan section 14. Additionally, a plurality of outlet guide vanes or struts 26, spaced apart from each other in the circumferential direction C, extend radially R between the outer casing 28 surrounding the LP spool 16 and the HP spool 18 and the nacelle 20. Therefore, the strut 26 supports the nacelle 20 relative to the outer shell 28, such that the outer shell 28 and the nacelle 20 define a bypass airflow passage 30 located therebetween.

[0039] The housing 28 generally surrounds or encloses the compressor section 32, combustion section 34, turbine section 36, and exhaust section 38 in a sequential flow order. For example, in some embodiments, the compressor section 32 may include a low-pressure (LP) compressor 40 of LP spool 16 and a high-pressure (HP) compressor 42 of HP spool 18, the HP compressor 42 being positioned downstream of the LP compressor 40 along a longitudinal centerline 12. Each compressor 40, 42 may instead include one or more rows of stator blades 44 interlaced with one or more rows of compressor rotor blades 46. Furthermore, in some embodiments, the turbine section 36 includes a high-pressure (HP) turbine 48 of HP spool 18 and a low-pressure (LP) turbine 50 of LP spool 16, the low-pressure (LP) turbine 50 being positioned downstream of the HP turbine 48 along a longitudinal centerline 12. Each turbine 48, 50 may instead include one or more rows of stator blades interlaced with one or more rows of turbine rotor blades 54. In a particular embodiment, the turbine section includes a first stator blade assembly or turbine nozzle 52 located downstream of the combustion chamber 106 and upstream of the turbine rotor blades.

[0040] Additionally, the LP spool 16 includes a low-pressure (LP) spool 56, and the HP spool 18 includes a high-pressure (HP) spool 58 concentrically positioned around the LP spool 56. In such an embodiment, the HP spool 58 is rotatably coupled to the rotor blades 54 of the HP turbine 48 and the rotor blades 46 of the HP compressor 42, such that rotation of the HP turbine rotor blades 54 rotatably drives the HP compressor rotor blades 46. As shown, the LP spool 56 is directly coupled to the rotor blades 54 of the LP turbine 50 and the rotor blades 46 of the LP compressor 40. Furthermore, the LP spool 56 is coupled to the fan section 14 via a gearbox 60. In this respect, rotation of the LP turbine rotor blades 54 rotatably drives the LP compressor rotor blades 46 and the fan blades 24.

[0041] In several embodiments, engine 10 can generate thrust to propel the aircraft. More specifically, during operation, air 62 enters the inlet section 64 of engine 10. Fan section 14 supplies a first portion of air 62 (shown by arrow 66) to bypass airflow passage 30 and a second portion of air 62 (shown by arrow 68) to compressor section 32. The second portion 68 of air 62 first flows through LP compressor 40, where rotor blades 46 progressively compress the second portion 68 of air 62. Next, the second portion 68 of air 62 flows through HP compressor 42, where rotor blades 46 continue to progressively compress the second portion 68 of air 62. The compressed second portion 68 of air 62 is then delivered to combustion section 34. In combustion section 34, the second portion 68 of air 62 mixes with fuel and burns to produce high-temperature, high-pressure combustion gas 70. Thereafter, combustion gas 70 flows through HP turbine 48, from which HP turbine rotor blades 54 extract a first portion of kinetic and / or thermal energy. This energy extraction causes HP shaft 58 to rotate, thereby driving HP compressor 42. The combustion gases 70 then flow through the LP turbine 50, where the LP turbine rotor blades 54 extract a second portion of kinetic and / or thermal energy. This energy extraction causes the LP shaft 56 to rotate, thereby driving the LP compressor 40 and fan section 14 via the gearbox 60. The combustion gases 70 then exit the engine 10 through the exhaust section 38.

[0042] The above and Figure 1 The configuration of the gas turbine engine 10 shown is only intended to place the subject matter within an exemplary field of application. Therefore, the subject matter can be readily adapted to any type of gas turbine engine configuration, including other types of aerospace gas turbine engines, marine gas turbine engines, and / or land-based / industrial gas turbine engines.

[0043] Figure 2 This is a cross-sectional view of one embodiment of the combustion section 34 of the gas turbine engine 10. As shown, the combustion section includes multiple fuel nozzles 112 (although in...). Figure 2 An annular burner assembly 100 with only one fuel nozzle 112 is shown. In several embodiments, the combustion section 34 includes a compressor discharge housing 118. In these embodiments, the compressor discharge housing 118 at least partially surrounds or otherwise encloses the burner assembly 100 in the circumferential direction C. In this respect, a compressor discharge chamber 120 is defined between the compressor discharge housing 118 and bushings 102, 104. The compressor discharge chamber 120 is instead configured to supply compressed air to the burner assembly 100. Specifically, as shown, air 68 exiting the HP compressor 42 is guided into the compressor discharge chamber 120 via inlet guide vanes 122. The air 68 within the compressor discharge chamber 120 is then supplied by the fuel nozzle 112 to the combustion chamber 106 of the burner assembly 100 for combustion of fuel.

[0044] The burner assembly 100 includes an inner bushing 102 extending circumferentially C. The burner assembly 100 also includes an outer bushing 104 positioned radially R outward from the inner bushing 102. The outer bushing 104 extends circumferentially C. In this respect, the inner bushing 102 and the outer bushing 104 define a combustion chamber 106 therebetween. Each bushing 102, 104 includes a first bushing or front bushing section 108 and a second bushing or rear bushing section 110, the second bushing or rear bushing section 110 being positioned downstream of the front bushing section 108 relative to the flow direction of fluid through the burner assembly 100, such as the flow of combustion gas 70. The burner assembly 100 includes a fuel nozzle 112 extending through a baffle assembly 107, the baffle assembly 107 providing a wall at an upstream end 121 of the combustion chamber 106. Each fuel nozzle 112 supplies a mixture of gaseous and / or liquid fuel with an oxidizer (such as air 68) to the combustion chamber 106. The fuel and air mixture is burned within the combustion chamber 106 to produce combustion gases 70. Although Figure 2 An annular burner assembly 100 is shown; in other embodiments, the combustion section 34 may include multiple burner assemblies 100. Other burner assembly configurations include cylindrical burners and cylindrical annular burners. There are also several other burner assembly configurations including vortex burners, detonation burners, or combinations of one or more types described herein.

[0045] refer to Figure 1-2 The engine 10 includes one or more static structures 290 defining a housing or frame for the engine 10. The static structure 290 is generally located upstream or downstream of the rotor assembly and may provide structural support for bearing assemblies, lubrication systems, gearbox assemblies, or drive shafts at the LP and / or HP spools. In various embodiments, such as those further described herein, the static structure 290 is located at the compressor section 32, the combustion section 34, or the turbine section 36.

[0046] Now for reference Figure 3 , Figure 4 as well as Figure 5 A perspective view is provided of an exemplary embodiment configured as part of a static structure 290 of a blade assembly 300 according to aspects of this disclosure. One embodiment of the blade assembly 300 includes an airfoil 310 extending through a gas flow path 302 of the engine 10. The blade assembly 300 may include a plurality of circumferentially arranged airfoils 310. In a particular embodiment, the airfoil 310 includes a leading edge 312, a trailing edge 314, a pressure side 316, and an intake side 318 (in... Figure 12-13 (As shown in the diagram). However, in other embodiments, the airfoil 310 may be symmetrical, such that the pressure side 316 is a first side, and the suction side 318 is a second side constructed substantially similarly to the first side. The impeller assembly 300 includes an outer band 320 extending from the airfoil 310 and forming an outer radius surface or outer flow path surface 322 of the gas flow path 302. The impeller assembly 300 may include an inner band 330 forming an inner radius surface or inner flow path surface 332 of the gas flow path 302. One or more of the leading edge 312, trailing edge 314, pressure side 316, suction side 318, outer flow path surface 322, or inner flow path surface 332 may define a flow path wall 304 at which the fluid circuit 340 extends through the impeller assembly 300. More specifically, as will be understood from the description below, the fluid circuit 340 extends through the impeller assembly 300 at locations in thermal communication with the flow path wall.

[0047] The fluid circuit 340 includes a first inlet opening 342 in fluid communication with the first cavity to receive a first fluid flow through the fluid circuit 340 (schematically shown via arrow 344). Brief Reference Figure 6-7 According to two exemplary embodiments of the present disclosure, a cross-sectional view of a portion of a fluid circuit 340 is depicted. The fluid circuit 340 includes an injector 350 formed and positioned at the fluid circuit 340. Return to Reference Figure 3 350 injector Figure 6-7 ) including fluid loop flow path 306 ( Figure 6-7 The second inlet opening 352 is in fluid communication with the fluid circuit 340, through which the first fluid flow 344 flows via the fluid circuit flow path 306. The second inlet opening 352 is in fluid communication with the second cavity to receive the second fluid flow (schematically shown via arrow 354) entering the fluid circuit 340 via the fluid circuit flow path 306.

[0048] As mentioned above, regarding Figure 3-5The provided embodiments are configured substantially similarly to each other. In various embodiments, the fluid circuit 340 extends along the radial direction R, circumferential direction C, and / or longitudinal direction L in a tortuous, torsion, or serpentine circuit. Figure 3-5 In the middle, the fluid loop flow path 306 extends through the impeller assembly 300 at the outer zone 320 in a torsional or serpentine flow path, such as along the circumferential direction C and / or the longitudinal direction L. Figure 4 The embodiments further depict a fluid loop flow path 306 that extends through the interior of the airfoil 310 in a torsional or serpentine flow path, for example, extending radially R along the span of the airfoil 310. However, it should be understood that the embodiments depicting the fluid loop 340 at the outer zone 320 may be additionally or alternatively applied to the inner zone 330.

[0049] The torsional or serpentine fluid circuit 340 includes a longitudinal direction L (e.g., Figure 5 As shown in the diagram), or along radial R (e.g., in...). Figure 8-9 As shown), or along the circumferential direction C (e.g., Figure 12-13 (As shown) Extending straight portion 341. The fluid circuit 340 further includes a curved portion 343 configured to circulate the fluid flow. The straight portion 341 and the curved portion 343 together allow the flow within the fluid circuit 340 to enter thermal communication in the region passing through the flow path wall 304.

[0050] Embodiments of the impeller assembly 300 (such as those depicted and described herein) allow a first fluid flow 344 from the first chamber to have a lower pressure and lower temperature relative to a second fluid flow 354 from the second chamber, providing cooling and thermal attenuation to the impeller assembly 300. The injector 350 entrains or pulls the lower-pressure first fluid flow 344 into the fluid circuit 340 and through the fluid circuit flow path 306 via the relatively high-pressure second fluid flow 354 and the second inlet opening 352. The injector 350 allows a relatively large amount of the low-temperature first fluid flow 344 to be pulled through the fluid circuit 340 by a relatively small amount of the relatively high-temperature second fluid flow 354. Similarly, the embodiments provided herein allow for improved component and engine cooling, engine performance, combustion efficiency, and fuel consumption, as well as improved thermal efficiency, by reducing the amount of fluid removed from the compressor section for cooling in other parts of the engine, or especially high-pressure, high-temperature compressed air. Furthermore, the embodiments provided herein allow for the utilization of relatively low-pressure fluids from fan bypass flow, third-flow bypass, lower shroud cavity or lower housing cavity, or atmospheric conditions.

[0051] Now for reference Figure 6-7The injector 350 may include a nozzle 356, which is positioned downstream of the second inlet opening 352 relative to the second fluid flow 354. The nozzle 356 is configured to have an outlet opening 348 relative to the second fluid flow 354 from the second inlet opening 352 toward the fluid circuit 340. Figure 3 The convergent cross-sectional area of ​​). In some embodiments, such as in Figure 7 As depicted, fluid circuit 340 forms a converging-diverging (CD) nozzle 358 positioned downstream of nozzle 356 in a fluid circuit flow path 306. The CD nozzle 358 is a portion of the fluid circuit flow path 306 where the flow path is compressed or narrowed to provide a reduced cross-sectional area, and then expands from the throat of the CD nozzle 358. The CD nozzle 358 is configured to accelerate the formation of a third fluid flow from the mixture of the first fluid flow 344 and the second fluid flow 354, schematically shown by arrow 346.

[0052] In a particular embodiment, such as Figure 7 The CD nozzle 358 shown can be positioned in the straight portion 341 of the fluid circuit 340. The correspondingly positioned CD nozzle 358 allows the mixed first and second fluid flows (i.e., the third fluid flow 346) to approach sonic flow conditions at the throat, and then expand to supersonic flow conditions as the cross-sectional area of ​​the fluid circuit flow path 306 increases downstream of the CD nozzle 358. The CD nozzle 358, positioned in the straight portion 341 of the fluid circuit 340, allows for increased flow velocity before the flow approaches the bend 343, and any flow losses associated with bends, turns, or curves in the fluid circuit flow path 306. This arrangement can mitigate stagnation of the fluid flow through the fluid circuit 340, or especially stagnation of the relatively low-pressure first fluid flow 344.

[0053] Now for reference Figure 8-9 An embodiment of a blade assembly 300 is provided, showing an exemplary cross-sectional view of the interior of an embodiment of an airfoil 310 passing through the blade assembly 300 along the longitudinal direction L. Regarding... Figure 8-9 The provided embodiment depicts a fluid loop 340 that extends torsionally through the airfoil 310 along radial R and longitudinal L. Figure 8 The embodiments depicted in the figure illustrate the following: Figure 4 The longitudinal cross-sectional view of the blade assembly 300 in the diagram shows the fluid circuit 340 extending torsionally from or near the leading edge 312 of the airfoil 310 to or near the trailing edge 314. It should be understood that "near the leading edge 312" refers to within 20% of the chord of the airfoil 310 at the leading edge 312. Similarly, it should be understood that "near the trailing edge 314" refers to within 20% of the chord of the airfoil 310 at the trailing edge 314.

[0054] refer to Figure 9 In the embodiment depicted, the fluid loop 340 extends longitudinally L through the interior of the airfoil 310 from a location between the leading edge 312 and the trailing edge 314. It should be understood that the fluid loop 340 may be configured to extend torsionally through the airfoil 310 from any portion of the airfoil 310 based on thermal connectivity at the airfoil 310. In various embodiments, the fluid loop 340 extends to a specific portion of the airfoil 310 at least based on a desired thermal attenuation or reduction of the thermal gradient at the airfoil 310.

[0055] Now for reference Figure 10-11 Basically based on the information Figure 3-9 The provided description provides exemplary embodiments. Figure 10-11 In this configuration, the fluid circuit 340 is arranged as a mesh or lattice structure. (Reference) Figure 10 The fluid circuit 340 forming a grid or mesh structure may include a plurality of branches 345 extending from a first base 347 to a second base 349. In some embodiments, the first base 347 may extend from an upstream end of the fluid circuit 340 or a first inlet opening 342. The second base 349 may extend from a downstream end of the fluid circuit 340 or an outlet opening 348. A second inlet opening 352 may be located at one or more of the branches 345 between the first base 347 and the second base 349. In some embodiments, in Figure 3-9 The straight portion 341 depicted and described in the text can form and include references. Figure 10 The base of the description and depiction is 347, 349.

[0056] refer to Figure 11 The fluid loop 340 forming a grid or lattice structure includes a reference flow path centerline 351. (The remaining text is omitted for clarity.) Figure 10 The surrounding wall of the fluid circuit 340 shown. Figure 11 In this context, the fluid circuit 340 may include multiple branches 345 and / or bases 347, 349 having small-diameter channels. The small-diameter channels can be formed via additive manufacturing processes, allowing for the formation of a mesh or grid structure at a portion of the airfoil 310 or along the entire span or chord of the airfoil 310. As provided herein, the fluid circuit 340 may be formed at specific portions of the static structure 290, such as the airfoil 310, based at least in part on desired thermal connectivity or thermal attenuation at or around the flow path of the component.

[0057] include Figure 3-4 and Figure 8-9 The twisted flow path shown or Figure 10-11 The grid structure shown Figure 10-11The fluid circuit 340 shown may include dimensions, diameters, or number of turns, intersections, branches, or other geometries to allow for the desired cooling effectiveness or thermal connectivity of the area extended by the fluid circuit 340 with respect to the engine 10 and the thermal loads, flow rates, or pressures experienced during operation. The fluid circuit 340 may form a tortuous or serpentine flow path, grid, grille, intersection, mesh, or other suitable pattern, or a combination thereof, connecting adjacent channels to intersect or connect channels that are in contact with each other. (As per this document regarding...) Figure 12-14 As further described, based on the desired thermal connectivity and / or flow characteristics of the surrounding fluid, the fluid loop 340 may extend within the surface, through the surface, or protrude into one or more flow paths as needed.

[0058] Return to reference Figure 8-9 An exemplary embodiment of a first cavity 307 and a second cavity 309 is provided, wherein a first fluid flow 344 is drawn from the first cavity 307 through a first inlet opening 342, and a second fluid flow 354 is drawn from the second cavity 309 through a second inlet opening 352. In some embodiments, the first cavity 307 is spaced from the second cavity 309 by a core housing 360 surrounding the impeller assembly 300. The core housing 360 may be positioned outward along the radial direction R of an outer belt 320 and extend along the longitudinal direction L and the circumferential direction C. The second cavity 309 is spaced from the first cavity 307 to allow for different pressures and / or temperatures of the fluid in the respective cavities. The outer belt 320 further spaces the second cavity 309 from a gas flow path 302. In other embodiments, the second cavity 309 is formed inward along the radial direction R and is spaced from the gas flow path 302 and via an inner belt 330.

[0059] Now for reference Figure 12-13 An exemplary view of an embodiment of the impeller assembly 300 along the radial direction R is provided. Regarding... Figure 3-5 The provided embodiments can be configured as described above. Figure 12-13 As depicted in the provided embodiments. In various embodiments, the fluid circuit 340 extends from the outer band 320 and / or the inner band 330 and enters the airfoil 310. The airfoil 310 may include an airfoil flow path surface 311 in fluid communication with the gas flow path 302. The airfoil flow path surface 311 is formed at the pressure side 316 and the suction side 318 of the airfoil 310. The airfoil 310 may further include an inner airfoil surface 313 inside the airfoil flow path surface 311. The airfoil flow path surface 311 and the inner airfoil surface 313 may together form a double-wall structure at the airfoil 310. Still in some embodiments, the airfoil 310 may include a hollow airfoil cavity 315 inside the inner airfoil surface 313.

[0060] about Figure 3-11 The provided embodiments can be configured, such as regarding Figure 12-13 One or two of the embodiments shown and described are used to depict and describe. Figure 3-11 The fluid loop 340 shown can extend through the airfoil 310 within the double-walled structure between the airfoil flow path surface 311 and the inner airfoil surface 313. Figure 11 In a particular embodiment depicted, a second inlet opening 352 into the fluid circuit 340 is in fluid communication with an airfoil cavity 315 defining a relatively high-pressure second cavity. In such an embodiment, a second fluid flow 354 is drawn from the airfoil cavity 315 to entrain a first fluid flow 344 through the fluid circuit flow path 306.

[0061] It should be understood that other embodiments of the airfoil 310 may include those without airfoil 310. Figure 12-13 The hollow airfoil cavity 315 shown has a solid or substantially solid volume. In some embodiments, the airfoil flow path surface 311 is in fluid communication with the gas flow path 302, while the inner airfoil surface 313 may represent a reference thermal gradient entering the airfoil 310, at which cooling, thermal attenuation, or thermal gradient reduction may be applied via the fluid loop 340 described herein. It should be further appreciated that the embodiments depicted and described herein can allow for improvements in the aerodynamic performance of the airfoil, such as by allowing for reductions in airfoil thickness, reductions in cross-sectional area relative to the gas flow path, or other dimensional variations that allow for increases or decreases in airfoil size relative to known blade assemblies.

[0062] Now for reference Figure 14 The diagram provides a radial view of an exemplary embodiment of the airfoil 310 along the radial direction R. Figure 14 The illustrated embodiments are configured as follows: Figure 3-13 The configurations described in the various embodiments shown. Figure 14 Exemplary locations through the airfoil 310 are depicted where the fluid circuit 340 may extend. In one embodiment, the fluid circuit 340 may extend within the airfoil cavity 315, as shown at fluid circuit 340a. Fluid circuit 340a may be attached to the inner airfoil surface 313, thereby allowing the fluid circuit flow path 306 to be formed at least partially by the walls of the fluid circuit 340 and the inner airfoil surface 313. Fluid circuit 340a may allow thermal communication at the airfoil 310, such as fluid circuit 340b described below. Additionally or alternatively, fluid circuit 340a may be formed at the airfoil cavity 315 and attached to the inner airfoil surface 313, or attached to the interior of the airfoil 310 at the airfoil flow path surface 311 (not shown) without the inner airfoil surface 313. The fluid circuit 340a allows thermal communication at the airfoil 310 and additionally allows thermal communication with fluids (e.g., air, lubricant, hydraulic fluid, fuel, etc.) within the airfoil cavity 315.

[0063] In another embodiment, the fluid loop 340 may extend within the double-walled structure of the airfoil 310 between the inner airfoil surface 313 and the airfoil flow path surface 311, as shown at fluid loop 340b. In yet another embodiment, the fluid loop 340 may protrude at least partially into the gas flow path 302, as shown at fluid loop 340c. In contrast to fluid loop 340b, which is formed inside the airfoil flow path surface 311 and enters the airfoil 310, fluid loop 340c may form corrugations, ridges, waves, or other surface features protruding into the gas flow path 302. Therefore, fluid loop 340c may allow greater thermal communication with the gas flow path 302. Additionally or alternatively, fluid loop 340c may produce specific flow characteristics for the flow of combustion gas 70 through the airfoil 310. Such flow characteristics may include turbulence, eddies, vortices, flow separation from the airfoil flow path surface 311, or other characteristics that may increase diffusivity, rotation, dissipation, or irregularity. Conversely, fluid circuit 340b may allow thermal connectivity at airfoil flow path surface 311 and / or inner airfoil surface 313, while allowing fluid (e.g., combustion gas 70) to flow laminarly across airfoil 310.

[0064] Return to reference Figure 1-2 In a particular embodiment, the static structure 290, including the impeller assembly 300 described herein, is positioned between the fan section 14 and the LP compressor 40. Figure 1 ), or between LP compressor 40 and HP compressor 42 ( Figure 1 ), or at the outlet of the HP compressor 42 at the inlet guide vane 122 of the combustion section 34 ( Figure 2 ), or at the outlet of combustion section 34 at the turbine nozzle 52 at the inlet of turbine section 36 ( Figure 2 ), or between LP Turbo 50 and HP Turbo 48 ( Figure 1 ), or downstream of the HP turbine 48 in exhaust section 38 ( Figure 1 ).

[0065] The embodiments of the static structure 290 and impeller assembly 300 provided herein can be configured as a turbine center frame, turbine impeller frame, or turbine rear frame located at or within the turbine section 36, between the combustion section 34 and the turbine section 36, or between the turbine section 36 and the exhaust section 38. Other embodiments can be configured as a compressor intermediate frame, fan intermediate frame, or diffuser or pre-diffuser impeller located at or within the compressor section 32, between the compressor section 32 and the combustion section 34, or between the fan section 14 and the compressor section 32.

[0066] In various embodiments, the first cavity 307 may be formed at or within the nacelle 20. The nacelle 20 may form a lower cover cavity or air chamber. As described above, the first cavity 307 is a low-pressure region relative to the second cavity 309 having a large flow rate of fluid (such as air). The first cavity 307 may receive airflow from atmospheric conditions or from downstream of the fan section 14. In some embodiments, the first cavity 307 is formed by a bypass airflow passage 30. The strut 26 may be configured with one or more flow path ducts to direct the first fluid flow to, for example, the impeller assembly 300 described herein. In yet another embodiment, the first cavity 307 is formed within a housing 28, as described with respect to the nacelle 20. In various embodiments, the nacelle 20 or the housing 28 or other suitable portions of the engine 10 may each include a first housing defining the first cavity 307.

[0067] refer to Figure 2 In one embodiment, the second cavity 309 may be formed within the compressor discharge housing 118 at the compressor discharge chamber 120. In another embodiment, the second cavity 309 may be formed at the turbine section 36, inside or outside the gas flow path through which the combustion gas 70 flows. As described above, the second cavity 309 is a high-pressure region relative to the first cavity 307. The high-pressure region is formed at least partially by unburned compressed air from the compressor section 32. The compressed air may be drawn in or discharged from the compressor section 32, or drawn out from the compressor discharge chamber 120. The compressed air received from the compressor section 32 and used to provide the second fluid flow 354 at the second cavity 309 may generally have the pressure and temperature corresponding to the compressed air at one or more stages of the compressor section 32. Conversely, the first fluid flow 344 in the first cavity 307 may generally have pressure and temperature corresponding to the external environment or atmospheric conditions surrounding the engine 10, or corresponding to airflow from the fan section 14, or corresponding to airflow from one or more stages at the compressor section 32, and the second fluid flow 354 is received from upstream of these one or more stages from the second cavity 309. In some embodiments, the first fluid flow 344 may be received from the LP compressor 40, while the second fluid flow 354 may be received from the HP compressor 42. In various embodiments, the core housing 360, the compressor discharge housing 118, or other suitable portions of the engine 10 may include a second housing forming the second cavity 309.

[0068] In an exemplary embodiment of engine 10, during operation at rated power output (i.e., maximum steady-state operating conditions, or maximum steady-state operating conditions that enable safe or stable operation of the engine, such as takeoff or full-load conditions), the first fluid flow 344 may have a first pressure between 9 pounds per square inch (psi) and 14.8 psi. The second fluid flow 354 may have a second pressure of at least 20 psi. In some embodiments, the second fluid flow 354 may have a second pressure of up to 250 psi. In various embodiments, the first fluid flow 344 and the second fluid flow 354 may include a temperature difference between 100 degrees Fahrenheit and 400 degrees Fahrenheit. However, it should be understood that the second pressure may be limited by the maximum pressure output at compressor section 32. Therefore, embodiments of engine 10 and impeller assembly 300 may allow a second pressure greater than 250 psi. During engine 10 operation, the second fluid flow 354 may entrain or pull the first fluid flow 344 through fluid circuit 340 via the pressure difference between the injector 350 and the fluid flow. The third fluid flow 346 (i.e., a mixture of the first fluid flow and the second fluid flows 344, 354) flows out through the outlet opening 348. In some embodiments, such as those described herein, the outlet opening 348 purges the third fluid flow 346 into one or more embodiments of the first cavity 307.

[0069] The embodiments provided herein allow the cooled shroud, impeller assembly, or nozzle to be supplied with a flow received from and purged from the first chamber via a fluid loop. The embodiments provided herein allow the impeller assembly, frame, or housing to be formed at locations such as those described herein using materials of a relatively lower grade, at a lower cost, or with greater ease of manufacture, due to the improved cooling of the first fluid flow primarily from the first chamber, which is significantly cooler than the relatively hot air from the compressor section. Additionally or alternatively, the embodiments provided herein may utilize known, higher-grade materials and allow for increased gas flow path temperatures and increased combustion gas outlet temperatures. Furthermore, one or more of these benefits can be obtained without increasing the amount of compressed air from the compressor section. Additionally, one or more of these benefits can be obtained while further reducing the amount of compressed air from the compressor section.

[0070] All or part of the static structure 290 and / or blade assembly 300, fluid loop 240, and ejector 350 may be formed via one or more additive manufacturing or 3D printing processes. The blade assembly 300 may be formed with the fluid loop 240 and ejector 350 described herein as a single, integral, monolithic, or integrated structure. In other embodiments, the static structure 290, blade assembly 300, or portions thereof may be formed as separate or detachable components attached together via one or more joining processes (such as welding, brazing, or using mechanical fasteners such as nuts, bolts, screws, tie rods, etc.). Still in other embodiments, the structures provided herein may be formed from forgings, machined materials, castings, or other suitable manufacturing processes. It should be understood that additive manufacturing can particularly allow the formation of the fluid loop 340, ejector 350, and other openings, conduits, flow paths, torsional loops, mesh structures, grid structures, double-wall structures, or specific locations, outer bands, inner bands, or airfoils within a double-wall structure.

[0071] In various embodiments, the first inlet opening 342, the second inlet opening 352, and the outlet opening 348 are sealed to the respective walls of the first cavity 307 and the second cavity 309 to allow for desired pressure differentials and accommodate relative thermal and mechanical deflection of the engine 10, or to form the walls of the embodiments of the first cavity 307 and the second cavity 309 described herein. Generally, the first cavity 307 and the second cavity 309 are separated from or sealed to allow a pressure and / or temperature difference between the first fluid flow 344 and the second fluid flow 354 for operation of the injector 350. Methods may include forming the first inlet opening 342 and the second inlet opening 352 as integral structures of the respective cavities 307, 309, such as via additive manufacturing methods, casting, forging, or other suitable manufacturing processes. Other methods may include bonding, welding, forming, fastening, or otherwise attaching fittings to the respective walls of the first cavity 307 and / or the second cavity 309 to form the respective first inlet opening 342, second inlet opening 352, or outlet opening 348. According to those skilled in the art, other suitable methods may be used to form the openings described herein to allow for pressure differentials and structural deflection.

[0072] Examples of powder-based additive layer fabrication include, but are not limited to, selective laser sintering (SLS), selective laser melting (SLM), direct metal laser sintering (DMLS), direct metal laser melting (DMLM), and electron beam melting (EBM) processes. Representative examples of suitable powder materials used in embodiments of the devices depicted and described herein may 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 known “superalloys” designed to have good oxidation resistance, 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, ECY768, 282, X45, PWA 1483, and CMSX (e.g., CMSX-4) single-crystal alloys. The objects to be manufactured according to this disclosure may be formed with one or more selected crystalline microstructures, such as directional solidification (DS) or single crystal (SX). However, as provided above, embodiments of engines including, for example, the fluid circuits and injectors described herein can allow the use of lower-strength materials in gas turbine engines at elevated operating temperatures, such as due to low-pressure, cryogenic air from the first chamber and / or improved cooling through the double-walled structure provided herein.

[0073] This written description uses examples to disclose preferred embodiments, including best practices, and also enables those skilled in the art to practice this disclosure, including making and using any device or system and performing any combined methods. The patentable scope of this disclosure is defined by the claims, but may include other examples that would occur to those skilled in the art. Such other examples are intended to be within the scope of the claims if they comprise structural elements that are not different from the literal language of the claims, or if they comprise equivalent structural elements that are not substantially different from the literal language of the claims.

[0074] Other aspects of this disclosure are provided by the subject matter of the following clauses:

[0075] 1. A gas turbine engine, the engine comprising: a blade assembly including a flow path wall, wherein a fluid circuit extends through the flow path wall, and wherein the fluid circuit defines a first inlet opening in fluid communication with a first cavity to receive a first fluid flow through the fluid circuit, and wherein the blade assembly includes an injector positioned at the fluid circuit, wherein the injector defines a second inlet opening in fluid communication with a second cavity to receive a second fluid flow through the injector and into the fluid circuit.

[0076] 2. A gas turbine engine according to one or more of these clauses, wherein the second inlet opening is positioned downstream of the first inlet opening along the fluid loop.

[0077] 3. A gas turbine engine according to one or more of these clauses, wherein the injector includes a nozzle positioned downstream of a second inlet opening relative to a second fluid flow entering the fluid circuit.

[0078] 4. A gas turbine engine according to one or more of these clauses, wherein the nozzle includes a converging cross-sectional area relative to the second fluid flow from the second inlet opening toward the outlet opening of the fluid loop.

[0079] 5. A gas turbine engine according to one or more of these clauses, wherein a fluid circuit forms a converging-diverging nozzle positioned downstream of the nozzle at the fluid circuit.

[0080] 6. A gas turbine engine according to one or more of these clauses, wherein the fluid circuit forms a torsional flow path, a grid structure, or a grille structure through the impeller assembly.

[0081] 7. A gas turbine engine according to one or more of these clauses, wherein the fluid circuit includes a straight portion extending in a longitudinal, radial or circumferential direction, and wherein the fluid circuit includes a curved portion configured to rotate a first fluid flow.

[0082] 8. A gas turbine engine according to one or more of these clauses, wherein the blade assembly includes an airfoil, wherein the flow path wall is the airfoil flow path surface, and wherein the airfoil includes a double-wall structure through which the fluid loop extends.

[0083] 9. A gas turbine engine according to one or more of these clauses, wherein the double-wall structure includes an airfoil flow path surface formed on the pressure side and the intake side of the airfoil, and wherein the double-wall structure includes an inner airfoil surface inside the airfoil flow path surface, and wherein a fluid loop extends between the airfoil flow path surface and the inner airfoil surface.

[0084] 10. A gas turbine engine according to one or more of these clauses, wherein an airfoil forms an airfoil cavity on the inner side of the inner airfoil surface, wherein a second cavity is an airfoil cavity, and wherein a second inlet opening is in fluid communication with the airfoil cavity to receive a second fluid flow from it into a fluid loop.

[0085] 11. A gas turbine engine according to one or more of these clauses, wherein the airfoil includes a leading edge and a trailing edge, and wherein a fluid loop extends from near the leading edge to near the trailing edge.

[0086] 12. A gas turbine engine according to one or more of these clauses, wherein the first inlet opening is close to the leading edge relative to the trailing edge.

[0087] 13. A gas turbine engine according to one or more of these clauses, the engine comprising: a nacelle forming a first cavity; and a core housing forming a second cavity, wherein the blade assembly is configured to receive a first fluid flow from the first cavity, the first fluid flow having a lower pressure relative to a second fluid flow from the second cavity.

[0088] 14. A gas turbine engine according to one or more of these clauses, the engine comprising: a compressor section, a combustion section and a turbine section in a sequential flow order, wherein a blade assembly is located at one or more of the compressor section, the combustion section or the turbine section.

[0089] 15. A gas turbine engine according to one or more of these clauses, wherein the flow path wall, fluid circuit and injector are formed as an integral, one-piece structure.

[0090] 16. A gas turbine engine according to one or more of these clauses, wherein the impeller assembly includes an outer belt, and wherein the fluid circuit extends along the outer belt along the flow path wall.

[0091] 17. A gas turbine engine according to one or more of these clauses, wherein the outer casing at least partially forms the gas flow path of the engine through which the combustion gases flow.

[0092] 18. A gas turbine engine according to one or more of these clauses, wherein the impeller assembly includes an inner belt, and wherein a fluid circuit extends through the inner belt through a flow path wall.

[0093] 19. A static structure for a gas turbine engine, the static structure comprising: a flow path wall, wherein a fluid loop extends through the flow path wall, and wherein the fluid loop includes a first inlet opening in fluid communication with a first cavity to receive a first fluid flow through the fluid loop, and wherein the static structure includes an injector positioned at the fluid loop, wherein the injector includes a second inlet opening in fluid communication with a second cavity to receive a second fluid flow through the injector and into the fluid loop.

[0094] 20. A static structure according to one or more of these clauses, wherein the static structure includes a double-walled structure through which a fluid loop extends.

Claims

1. A gas turbine engine, characterized in that, The engine includes: A blade assembly includes a flow path wall through which a fluid loop extends, and wherein the fluid loop defines a first inlet opening in fluid communication with the first cavity to receive a first fluid flow through the fluid loop, and wherein the blade assembly includes an ejector positioned at the fluid loop, wherein the ejector defines a second inlet opening in fluid communication with the second cavity to receive a second fluid flow through the ejector and into the fluid loop, the second cavity being fluidly spaced from the first cavity, the second fluid flow having a higher pressure and temperature than the first fluid flow. The injector includes a nozzle positioned downstream of the second inlet opening relative to the second fluid flow toward the fluid circuit; and The nozzle is configured to propel the first fluid flow through the fluid circuit by injecting the second fluid flow from the second inlet opening through the nozzle into the fluid circuit.

2. The gas turbine engine according to claim 1, characterized in that, in, The second inlet opening is positioned downstream of the first inlet opening along the fluid loop.

3. The gas turbine engine according to claim 1, characterized in that, in, The nozzle includes a converging cross-sectional area relative to the second fluid flow from the second inlet opening toward the outlet opening of the fluid circuit.

4. The gas turbine engine according to claim 1, characterized in that, in, The fluid circuit forms a converging-diverging nozzle, which is positioned downstream of the nozzle in the fluid circuit.

5. The gas turbine engine according to claim 1, characterized in that, in, The fluid circuit forms a torsional flow path, a mesh structure, or a grid structure through the impeller assembly.

6. The gas turbine engine according to claim 5, characterized in that, in, The fluid circuit includes a straight portion extending in a longitudinal, radial, or circumferential direction, and wherein the fluid circuit includes a curved portion configured to rotate the first fluid flow.

7. The gas turbine engine according to claim 1, characterized in that, in, The blade assembly includes an airfoil, wherein the flow path wall is an airfoil flow path surface, and wherein the airfoil includes a double-wall structure through which the fluid loop extends.

8. The gas turbine engine according to claim 7, characterized in that, in, The double-wall structure includes the airfoil flow path surface, which is formed on the pressure side and the suction side of the airfoil, and wherein the double-wall structure includes an inner airfoil surface inside the airfoil flow path surface, and wherein the fluid loop extends between the airfoil flow path surface and the inner airfoil surface.

9. The gas turbine engine according to claim 8, characterized in that, in, The airfoil forms an airfoil cavity on the inner side of the inner airfoil surface, wherein the second cavity is the airfoil cavity, and wherein the second inlet opening is in fluid communication with the airfoil cavity to receive the second fluid flow from it into the fluid circuit.

10. The gas turbine engine according to claim 7, characterized in that, in, The airfoil includes a leading edge and a trailing edge, wherein the fluid loop extends from near the leading edge to near the trailing edge.

11. The gas turbine engine according to claim 10, characterized in that, in, The first inlet opening is close to the leading edge relative to the trailing edge.

12. The gas turbine engine according to claim 1, characterized in that, The engine includes: A compressor section, a combustion section, and a turbine section in a sequential flow order, wherein the impeller assembly is located at one or more of the compressor section, the combustion section, or the turbine section.

13. The gas turbine engine according to claim 1, characterized in that, in, The impeller assembly includes an outer belt, and the fluid circuit extends along the outer belt.

14. The gas turbine engine according to claim 13, characterized in that, in, The outer band at least partially forms the gas flow path of the engine through which the combustion gases flow.

15. The gas turbine engine according to claim 1, characterized in that, in, The impeller assembly includes an inner belt, and the fluid circuit extends through the inner belt.

16. A static structure for a gas turbine engine, characterized in that, The static structure includes: A flow path wall, wherein a fluid loop extends through the flow path wall, and wherein the fluid loop includes a first inlet opening in fluid communication with a first cavity to receive a first fluid flow through the fluid loop, and wherein the static structure includes an ejector positioned at the fluid loop, wherein the ejector includes a second inlet opening in fluid communication with a second cavity to receive a second fluid flow through the ejector and into the fluid loop, the second cavity being fluidly separated from the first cavity, the second fluid flow having a higher pressure and temperature than the first fluid flow. The injector includes a nozzle configured to propel the first fluid flow through the fluid circuit by injecting the second fluid flow from the second inlet opening through the nozzle into the fluid circuit. The fluid circuit further includes an outlet opening downstream of the injector. The outlet opening is in fluid communication with the outlet chamber, which is fluidly separated from the gas flow path through the gas turbine engine.

17. The static structure according to claim 16, characterized in that, in, The static structure includes a double-walled structure through which the fluid circuit extends.

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