Gas turbine engine

Abstract

A gas turbine engine includes: a fan assembly including a fan; a turbine drivingly coupled to the fan and including a compressor section, a combustion section, and a turbine section arranged in a serial flow sequence, and partially defining a working gas flow path, the gas turbine engine defining a bypass passage above the turbine, the turbine defining an annular cooling passage extending between a CP inlet and a CP outlet, the CP inlet is in airflow communication with the working gas flow path, the bypass channel, or both; an accessory system; a heat exchanger positioned in thermal communication with the annular cooling channel at a location between the CP inlet and the CP outlet, the heat exchanger in thermal communication with the accessory system; and an exhaust cooling system defining a BC inlet in airflow communication with the annular cooling channel at a location between the CP inlet and the CP outlet.

Description

Technical Field

[0001] The present disclosure relates to a gas turbine engine having cooling passages and an exhaust cooling system. Background Art

[0002] A gas turbine engine typically includes a turbine and a rotor assembly. Gas turbine engines, such as turbofan engines, can be 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 surrounded by an outer nacelle. The outer nacelle, together with the turbine, can define a bypass duct.

[0003] Generally speaking, improvements to turbofan engines in terms of thermal management and aerodynamics would be welcome in the art. BRIEF DESCRIPTION OF THE DRAWINGS

[0004] A full and enabling disclosure of the present disclosure, including the best mode thereof, to one of ordinary skill in the art is set forth in the specification with reference to the accompanying drawings, in which:

[0005] Figure 1 is a cross-sectional view of a gas turbine engine according to an exemplary aspect of the present disclosure.

[0006] Figure 2 is a schematic diagram of a portion of a turbine according to an exemplary embodiment of the present disclosure.

[0007] Figure 3 According to an exemplary embodiment of the present disclosure Figure 2 Schematic cross-sectional view of a cooling channel of a turbine.

[0008] Figure 4 is a schematic diagram of a portion of a turbine according to another exemplary embodiment of the present disclosure.

[0009] Figure 5 yes Figure 4 Schematic diagram of an exemplary turbine shroud.

[0010] Figure 6 is a schematic diagram of a shroud of a turbine according to an exemplary embodiment of the present disclosure.

[0011] Figure 7 is a schematic diagram of a portion of a turbine according to another exemplary embodiment of the present disclosure.

[0012] Figure 8 is a schematic diagram of an ejector according to another exemplary embodiment of the present disclosure.

[0013] Figure 9 is a schematic cross-sectional view of a cooling passage and a variable exhaust assembly according to an exemplary embodiment of the present disclosure.

[0014] Figure 10 is a schematic diagram of a portion of a turbine according to another exemplary embodiment of the present disclosure.

[0015] Figure 11 It is along Figure 10 Line 11-11 Figure 10 Schematic diagram of an exemplary turbine.

[0016] Figure 12 is a schematic diagram of a portion of a turbine according to yet another exemplary embodiment of the present disclosure.

[0017] Figure 13 is a schematic diagram of a portion of a turbine according to yet another exemplary embodiment of the present disclosure. DETAILED DESCRIPTION

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

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

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

[0021] The term "at least one" in the context of, for example, "at least one of A, B, and C" means only A, only B, only C, or any combination of A, B, and C.

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

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

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

[0025] The terms "low" and "high" or their respective comparatives (e.g., lower, higher, as applicable) when used with reference to a compressor, turbine, shaft or spool component, etc., refer to relative speeds within the engine, unless otherwise specified. For example, a "low-pressure turbine" or "low-speed turbine" defines a component configured to operate at a lower speed (e.g., maximum allowable speed) than a "high turbine" or "high-speed turbine" of the engine.

[0026] The terms "fore" and "aft" refer to relative positions within a gas turbine engine or vehicle and refer to the normal operating attitude of the gas turbine engine or vehicle. For example, for a gas turbine engine, the front position refers to the position closer to the engine inlet, and the aft position refers to the position closer to the engine nozzle or exhaust.

[0027] The terms "upstream" and "downstream" refer to relative directions relative to the flow of fluid in a fluid path. For example, "upstream" refers to the direction from which the fluid is flowing, and "downstream" refers to the direction toward which the fluid is flowing.

[0028] As used herein, the terms “first,” “second,” and “third” may be used interchangeably to distinguish one component from another and are not intended to indicate the position or importance of each component.

[0029] The present disclosure generally relates to gas turbine engines, such as turbofan engines.

[0030] As turbofan engine design standards require operating conditions reaching higher pressures and temperatures, it has been recognized in the art that increased cooling capabilities would be welcome. Some gas turbine engines bleed air from, for example, the high-pressure compressor and utilize the bleed air to cool various accessory systems of the turbofan engine. However, this configuration can result in inefficiencies because the air bleed from the high-pressure compressor undergoes a relatively large amount of work and is typically at a higher temperature.

[0031] Therefore, in one exemplary aspect, a gas turbine engine is provided having a turbine defining a cooling passage extending between an inlet and an outlet. The inlet is in gaseous communication with a working gas flow path of the turbine at a location upstream of a compressor section, in gaseous communication with a bypass passage of the gas turbine engine at a location outboard of the turbine, or both. The outlet is in gaseous communication with the bypass passage. A heat exchanger is disposed in thermal communication with the gas flow through the cooling passage. The heat exchanger may be used to cool one or more accessory systems of the gas turbine engine.

[0032] In such an exemplary aspect, the airflow provided to the heat exchanger may be relatively cool and may be provided from a location upstream of a substantial amount of compression (which would result in work losses).

[0033] In another exemplary aspect of the present disclosure, there is provided a device for promoting airflow through a cooling passage (such as the cooling passage described above). Specifically, using an exemplary aspect of the present disclosure, there is provided a turbine of a gas turbine engine having a variable exhaust assembly having a variable exhaust duct extending between a variable exhaust (VB) inlet and a VB outlet. The VB inlet is connected to the working gas flow path airflow at a position downstream of the cooling passage inlet of the cooling passage, and the VB outlet is connected to the cooling passage airflow for promoting airflow through the annular cooling passage. In this way, a small amount of high-pressure airflow can be utilized to promote low-pressure airflow through the cooling passage. This can allow the low-pressure airflow to be used to cool various components of the turbine, which can result in an overall more efficient cooling system and gas turbine engine.

[0034] In another exemplary aspect, an exhaust cooling system for a gas turbine engine draws airflow from a cooling passage. The exhaust cooling system may be, for example, an under-cowl cooling system, a clearance control system, a combination thereof, or other systems of a gas turbine engine that utilize cooling air without requiring high pressure treatment of the cooling air. In such an exemplary aspect, the airflow provided to the exhaust cooling system of the gas turbine engine may be relatively cool and may be provided from a location upstream of a significant amount of compression (which would result in work losses).

[0035] Referring now to the drawings, wherein like numerals refer to like elements throughout, Figure 1 is a schematic cross-sectional view of a gas turbine engine 10 according to an exemplary embodiment of the present disclosure. More specifically, Figure 1 In the embodiment of the present invention, the gas turbine engine 10 is a high bypass turbofan jet engine, sometimes also referred to as a "turbofan engine". Figure 1As shown, the gas turbine engine 10 defines an axial direction A (extending parallel to a longitudinal centerline 12 for reference), a radial direction R, and a circumferential direction C extending about the longitudinal centerline 12. Generally speaking, the gas turbine engine 10 includes a fan section 14 and a turbine 16 disposed downstream of the fan section 14, the turbine 16 drivingly coupled to a fan 38 of the fan section 14.

[0036] The exemplary turbine 16 shown generally includes a substantially tubular casing 18 that defines an annular turbine inlet 20. Casing 18 encloses, in serial flow relationship, a compressor section, including a supercharger or low-pressure (LP) compressor 22 and a high-pressure (HP) compressor 24; a combustion section 26; a turbine section, including 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) drivingly connects the HP turbine 28 to the HP compressor 24. A low-pressure (LP) shaft 36 (which may additionally or alternatively be a spool) drivingly connects the LP turbine 30 to the LP compressor 22. The compressor section, combustion section 26, turbine section, and exhaust nozzle section 32 together define a working gas flow path 37. In this manner, it should be understood that the turbine inlet 20 is the inlet to the working gas flow path 37.

[0037] In the illustrated embodiment, the turbine inlet 20 is positioned directly downstream of the fan 38 (i.e., there are no intervening structures such as blades, buckets, or struts therebetween). Furthermore, it should be understood that for the illustrated embodiment, the LP compressor 22 is located downstream of the fan 38, and there are no intervening compression stages between the fan 38 and the LP compressor 22.

[0038] For the illustrated embodiment, the fan section 14 includes a fan 38 having a plurality of fan blades 40 coupled to a disk 42 in a spaced-apart manner. As shown, the fan 38 is a single-stage fan, and the fan blades 40 extend outwardly from the disk 42 generally in a radial direction R. Each fan blade 40 is rotatable relative to the disk 42 about a pitch axis P by virtue of the fan blades 40 being operably coupled to a suitable pitch change mechanism 44, the pitch change mechanism 44 being configured to collectively change the pitch of the fan blades 40, e.g., in unison. The gas turbine engine 10 further includes a reduction gearbox 46, and the fan blades 40, disk 42, and pitch change mechanism 44 are rotatable together about the longitudinal centerline 12 via the LP shaft 36 across the reduction gearbox 46. The reduction gearbox 46 includes a plurality of gears for adjusting the rotational speed of the fan 38 relative to the rotational speed of the LP shaft 36 so that the fan 38 can rotate at a more efficient fan speed.

[0039] Still refer to Figure 1In the exemplary embodiment of the fan section 14 , the disk 42 is covered by a rotatable front hub 48 (sometimes also referred to as a “spinner”) of the fan section 14 . The front hub 48 has an aerodynamic profile to facilitate airflow through the plurality of fan blades 40 .

[0040] Additionally, the exemplary fan section 14 includes an annular fan casing or outer nacelle 50 that circumferentially surrounds the fan 38 and / or at least a portion of the turbine 16. It should be understood that, in the illustrated embodiment, the outer nacelle 50 is supported relative to the turbine 16 by a plurality of circumferentially spaced outlet guide vanes 52. Furthermore, a downstream section 54 of the outer nacelle 50 extends above an outer portion of the turbine 16 to define a bypass passage 56 therebetween. The bypass passage 56 is at least partially defined above the turbine 16.

[0041] During operation of gas turbine engine 10, a volume of air 58 enters gas turbine engine 10 through outer nacelle 50 and associated inlet 60 of fan section 14. As air 58 passes through fan blades 40, a first portion of air 62 is directed or channeled into bypass passage 56, while a second portion of air 64 (as indicated by arrow 64) is directed or channeled into working gas flow path 37, or more specifically, into LP compressor 22. The ratio between first portion of air 62 and second portion of air 64 is generally referred to as the bypass ratio. The pressure of second portion of air 64 is then increased as it is directed through HP compressor 24 and into combustion section 26, where it is mixed with fuel and combusted to provide combustion gases 66.

[0042] Combustion gases 66 are directed through HP turbine 28, wherein a portion of thermal and / or kinetic energy from combustion gases 66 is extracted via successive stages of HP turbine stator blades 68 coupled to casing 18 and HP turbine rotor blades 70 coupled to HP shaft 34, thereby rotating HP shaft 34, which supports operation of HP compressor 24. Combustion gases 66 are then directed through LP turbine 30, wherein a second portion of thermal and kinetic energy is extracted from combustion gases 66 via successive stages of LP turbine stator blades 72 coupled to casing 18 and LP turbine rotor blades 74 coupled to LP shaft 36, thereby rotating LP shaft 36, which supports operation of LP compressor 22 and / or rotation of fan 38.

[0043] The combustion gases 66 are then directed through the exhaust nozzle section 32 of the turbine 16 to provide propulsive thrust. Simultaneously, the pressure of the first portion of air 62 is significantly increased as it is directed through the bypass passage 56 before exiting the fan nozzle exhaust section 76 of the gas turbine engine 10, also providing propulsive thrust. The HP turbine 28, the LP turbine 30, and the exhaust nozzle section 32 at least partially define a hot gas path 78 for directing the combustion gases 66 through the turbine 16.

[0044] Briefly, it should also be appreciated that the gas turbine engine 10 further includes one or more cooling systems that may require a relatively cool air flow under one or more operating conditions.

[0045] For example, the gas turbine engine 10, or more specifically, the turbine 16, defines one or more under-cowl regions 80 (regions inboard of the outer casing 18 in the radial direction R and outboard of the working gas flow path 37 in the radial direction R). The illustrated turbine 16 includes an under-cowl ventilation cooling system 82 for providing a relatively cool air flow during operation to maintain components and structures within these under-cowl regions 80 at a desired temperature. Specifically, the under-cowl ventilation cooling system 82 can provide relatively cool air to the under-cowl regions 80 in the radial direction R outwardly of (and aligned with in the axial direction A in alignment with) the combustion section 26, the HP turbine 28, the LP turbine 30, the exhaust port 32, or a combination thereof.

[0046] The one or more cooling systems further include a clearance control system 84. The clearance control system 84 is essentially a thermal management system for the outer shroud to control thermal growth under certain operating conditions, thereby controlling the clearance (in the radial direction R) between the outer shroud and the rotating rotor blades inside thereof. For example, the illustrated clearance control system 84 includes a shroud cooling assembly that is thermally coupled to one or more shrouds or other structures outside of the corresponding turbine to provide this function. The clearance control system 84 (and the shroud cooling assembly) can more specifically receive a relatively cool airflow 86 during one or more operating periods to perform this function.

[0047] The gas turbine engine 10 may additionally or alternatively include other cooling systems. It is noted that for the embodiment shown, the cooling systems (under-cowl ventilation cooling system 82 and clearance control system 84) may not require the relatively cool airflow to be at a high pressure to operate as intended. Thus, the exhaust gas flow from the annular duct described below (see, for example, Figure 2 etc.) can provide the relatively cool airflow required for these cooling systems.

[0048] However, it should be understood that Figure 1The exemplary gas turbine engine 10 shown in FIG. 1 is provided as an example only, and in other exemplary embodiments, the gas turbine engine 10 may have any other suitable configuration. For example, although the illustrated gas turbine engine 10 is configured as a ducted gas turbine engine (i.e., including an outer nacelle 50 ), in other embodiments, the gas turbine engine 10 may be an unducted gas turbine engine (such that the fan 38 is an unducted fan and the outlet guide vanes 2 are cantilevered from the casing 18 ). Additionally or alternatively, although the illustrated 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 are non-rotatable about the pitch axis P), or both. It should also be understood that in other exemplary embodiments, aspects of the present disclosure may be incorporated into any other suitable gas turbine engine. For example, in other exemplary embodiments, aspects of the present disclosure may be incorporated into, for example, a turboprop gas turbine engine, as appropriate.

[0049] Now refer to Figure 2 , showing Figure 1 A close-up schematic diagram of a portion of an exemplary gas turbine engine 10 is provided. Specifically, Figure 2 The view is Figure 1 A close-up of the turbine 16 is shown, illustrating the turbine inlet 20, the compressor section including the LP compressor 22 and the HP compressor 24, the LP shaft 36, and the reduction gearbox 46. Additionally, the casing 18 of the turbine 16 is shown extending around at least a portion of the compressor section, with a bypass passage 56 partially defined thereby.

[0050] As from Figure 2 As will be appreciated from the view of FIG, the turbine 16 further includes a compressor front frame 100 and a compressor midframe 102. The compressor front frame 100 includes struts 104 that extend through the working gas flow path 37 at a location upstream of the LP compressor 22 and downstream of the turbine inlet 20. Similarly, the compressor midframe 102 includes struts 106 that extend through the working gas flow path 37 at a location downstream of the LP compressor 22 and upstream of the HP compressor 24. The compressor front frame 100 and the compressor midframe 102 may provide structural support for various components of the gas turbine engine 10.

[0051] It should be appreciated that the exemplary gas turbine engine 10 shown includes one or more accessory systems 108 for facilitating operation of the gas turbine engine 10. The one or more accessory systems 108 may include one or more of an oil lubrication system, a fuel delivery system, a cooled cooling air (CCA) system, an engine controller cooling system, and the like. Figure 2 A single accessory system 108 is schematically shown by way of example in FIG.

[0052] Additionally, to aid in cooling one or more accessory systems 108, such as the illustrated accessory system 108, the turbine 16 further defines a cooling passage 110 extending between a cooling passage (CP) inlet 112 and a CP outlet 114. As should be understood from the description herein, the cooling passage 110 is an annular cooling passage.

[0053] The CP inlet 112 is in gaseous communication with the working gas flow path 37 at a location upstream of the compressor section, the bypass passage 56, or both of the turbine 16. More specifically, for the illustrated embodiment, the CP inlet 112 is in gaseous communication with the working gas flow path 37 at a location upstream of the LP compressor 22 and downstream of the turbine inlet 20. More specifically, for the illustrated embodiment, the CP inlet 112 is aligned with the compressor front frame 100 along the axial direction A of the gas turbine engine 10. In this manner, the CP inlet 112 is configured to receive gas flow from the working gas flow path 37 supplied by the fan section 14 (see FIG. Figure 1 For example, the pressure of the airflow received and provided to the cooling passage 110 through the CP inlet 112 may be substantially equal to (eg, within 10% of) the pressure of the airflow provided through the turbine inlet 20 .

[0054] Still refer to Figure 2 , the CP outlet 114 is in gaseous communication with the bypass passage 56. Specifically, for the illustrated embodiment, the CP outlet 114 is in gaseous communication with the bypass passage 56 at a location rearward of the CP inlet 112 and forward of the compressor midframe 102. Specifically, for the illustrated embodiment, the CP outlet 114 is aligned with the compressor section in the axial direction A.

[0055] Furthermore, the turbine 16 includes a heat exchanger 116 in thermal communication with the airflow through the cooling passage 110. Specifically, the heat exchanger 116 is positioned within or defines a portion of the cooling passage 110. In this manner, the heat exchanger 116 may be configured to transfer heat from the fluid to the airflow through the cooling passage 110.

[0056] Specifically, for the illustrated embodiment, heat exchanger 116 is further in thermal communication with accessory systems 108 of gas turbine engine 10 to transfer heat from accessory systems 108 to the airflow through cooling passage 110. In this manner, cooling passage 110 may provide cooling for accessory systems 108.

[0057] It should be understood that for the illustrated embodiment, while a single heat exchanger 116 and a single accessory system 108 are depicted, in other exemplary embodiments, other suitable configurations may be provided.

[0058] For example, now briefly refer to Figure 3 , provides a schematic cross-sectional view of a turbine 16 including a cooling passage 110 according to an exemplary aspect of the present disclosure. Figure 3 The exemplary cooling passages 110 and heat exchangers 116 may be described with reference to the above Figure 2 The exemplary cooling passages 110 and heat exchangers 116 are depicted as being configured in a similar manner.

[0059] However, for the illustrated embodiment, the heat exchanger 116 is a first heat exchanger 116A of a plurality of heat exchangers 116 arranged along a circumferential direction C of the gas turbine engine 10 . The plurality of heat exchangers 116 are each positioned within the cooling passage 110 .

[0060] Notably, for the illustrated embodiment, the accessory system 108 is also the first accessory system 108A in the plurality of accessory systems 108. Each of the plurality of accessory systems 108 utilizes one or more of the plurality of heat exchangers 116. For example, the first accessory system 108A is in thermal communication with the first heat exchanger 116A. The second accessory system 108B in the plurality of accessory systems 108 is in thermal communication with the second heat exchanger 116B, the third heat exchanger 116C, and the fourth heat exchanger 116D. Notably, the second heat exchanger 116B, the third heat exchanger 116C, and the fourth heat exchanger 116D are arranged in a serial flow sequence. Furthermore, the third accessory system 108C in the plurality of accessory systems 108 is in thermal communication with the fifth heat exchanger 116E and the sixth heat exchanger 116F.

[0061] In this manner, it should be appreciated that the airflow through the cooling passage 110 may serve as a heat sink for the various accessory systems 108 of the gas turbine engine 10 .

[0062] Furthermore, it should be understood that in other exemplary embodiments, other suitable structures may be provided to assist in generating airflow through the cooling passage 110 of the present disclosure.

[0063] For example, now refer to Figure 4, provides a close-up schematic illustration of a gas turbine engine 10 according to another exemplary aspect of the present disclosure. Figure 4 The exemplary gas turbine engine 10 shown in FIG. 1 may be used in the same manner as described above with reference to FIG. Figure 1 and Figure 2 The exemplary gas turbine engine 10 depicted is configured in a similar manner. Like or similar numbers may refer to like or similar parts.

[0064] For example, Figure 4 The exemplary gas turbine engine 10 shown in FIG generally includes a turbine 16 defining a cooling passage 110 extending between a CP inlet 112 and a CP outlet 114. The CP inlet 112 is in gaseous communication with the working gas flow path 37 of the turbine 16 at a location upstream of the compressor section of the turbine 16. The CP outlet 114 is in gaseous communication with a bypass passage 56 of the gas turbine engine 10, which is defined between the outer nacelle 50 and the turbine 16.

[0065] It is worth noting that, as in the above-described embodiment, the CP inlet 112 is configured to receive the CP inlet 112 from the gas turbine engine 10 (see Figure 1 ) of the fan 38 compressed air flow, but is located upstream of any additional compression stage. Therefore, to help generate air flow through the cooling passage 110, the turbine 16 includes a device for promoting air flow through the cooling passage 110.

[0066] More specifically, for the embodiment shown, the apparatus includes an inlet scoop 118 extending into the working gas flow path 37 at the CP inlet 112 to divert a portion of the gas flow through the working gas flow path 37 through the CP inlet 112 of the cooling passage 110. In the embodiment shown, the inlet scoop 118 is an annular scoop extending 360° about the longitudinal centerline 12 of the gas turbine engine 10. Furthermore, for the embodiment shown, the inlet scoop 118 is a fixed structure.

[0067] However, it should be understood that in other exemplary embodiments, the inlet scoop 118 may be configured in any other suitable manner (e.g., it may be configured as a plurality of separate inlet scoops 118 arranged along the circumferential direction C of the gas turbine engine 10, it may be a variable scoop that can be deployed and retracted, etc.).

[0068] Furthermore, for the illustrated embodiment, the means for promoting airflow through the cooling passage 110 additionally includes a shroud 120 extending above the CP outlet 114 of the cooling passage 110. The shroud 120 extends into the bypass passage 56 such that the cross-sectional area of ​​the bypass passage 56 at the shroud 120 is smaller than the cross-sectional area of ​​the bypass passage 56 immediately upstream of the shroud 120. In this manner, the shroud 120 can form a nozzle to increase the velocity of the airflow through the bypass passage 56, thereby reducing the static pressure at the CP outlet 114 of the cooling passage 110. Thus, the shroud 120 can generate an incremental pressure to promote airflow through the cooling passage 110.

[0069] Briefly, refer to Figure 5 , shows a partial schematic cross-sectional view of the cover 120 viewed along the axial direction A. As shown, for the illustrated embodiment, the cover 120 is a relatively smooth cover 120 .

[0070] Alternatively, a brief reference Figure 6 , showing a partial schematic cross-sectional view of a cover 120 according to another exemplary embodiment of the present disclosure, which may be incorporated into Figure 4 The gas turbine engine 10. Figure 6 In an embodiment, the shroud 120 is configured as a mixer having a plurality of lobes 121 spaced apart along the circumferential direction C, the plurality of lobes 121 having a continuous series of peaks and valleys such that the radial height of the shroud 120 (and the lobes 121 ) defines a sinusoidal pattern along the circumferential direction C. The plurality of lobes 121 can locally raise the radial height of the shroud 120 along the circumferential direction C, thereby minimizing the outlet static pressure at the CP outlet 114 at the local location above the lobes 121 of the shroud 120 by increasing the Mach number of the fan airflow at the local location.

[0071] Furthermore, in other exemplary embodiments, the device for causing airflow through the cooling passage 110 may be any other suitable device, such as a pump or compressor, or an ejector in airflow communication with a high-pressure air source. For example, in certain exemplary embodiments, the device may include a device in airflow communication with a variable exhaust assembly and further in airflow communication with the cooling passage 110 (see, e.g., below). Figure 8 Alternatively, in other embodiments, the high pressure air source may be any other suitable high pressure air source, such as LP compressor exhaust, HP compressor exhaust, turbine exhaust, or a combination thereof. In certain embodiments, the ejector may be positioned downstream of the heat exchanger 116 (e.g., Figure 8 Alternatively, in other embodiments, the ejector may be positioned upstream of the heat exchanger 116 ( Figure 2 and Figure 4 ).

[0072] Reference again Figure 2 and Figure 4 In one or more of the above exemplary embodiments, it will be appreciated that the cooling passages 110 may be configured to receive sufficient air flow therethrough to provide a desired amount of cooling for one or more accessory systems 108 of the gas turbine engine 10. For example, in certain exemplary embodiments, it will be appreciated that during operation of the gas turbine engine 10 at a first operating condition, the cooling passages 110 may be configured to receive 2% to 20% of the total airflow through the working gas flow path 37 at a location upstream of the CP inlet 112 and downstream of the turbine inlet 20. For example, in certain exemplary embodiments, the cooling passages 110 may be configured to receive 4% to 12% of the total airflow through the working gas flow path 37 at a location upstream of the CP inlet 112 and downstream of the turbine inlet 20. In certain exemplary aspects, the first operating condition may be a high power operating condition (e.g., a takeoff operating condition) in which a relatively high amount of cooling may be desired for the gas turbine engine 10.

[0073] It should be appreciated that a turbine including defined cooling passages according to one or more exemplary aspects of the present disclosure can allow relatively cool airflow to be used to cool one or more accessory systems of the gas turbine engine before the airflow is subjected to additional compression stages. In this manner, the cooling passages can provide a relatively efficient means of cooling various accessory systems, as the airflow is utilized before being applied to them in compressed form for operation. Furthermore, by transferring heat to the airflow before providing it to the bypass passage, additional energy can be transferred to the bypass passage, which can improve the overall propulsive efficiency of the gas turbine engine.

[0074] Now refer to Figure 7 , provides a close-up schematic illustration of a gas turbine engine 10 according to yet another exemplary aspect of the present disclosure. Figure 7 The exemplary gas turbine engine 10 shown in FIG. 1 may be used in the same manner as described above with reference to FIG. Figure 1 and Figure 2 The exemplary gas turbine engine 10 depicted is configured in a similar manner. Like or similar numbers may refer to like or similar parts.

[0075] For example, Figure 7 The exemplary gas turbine engine 10 shown in FIG generally includes a turbine 16 defining a cooling passage 110 extending between a CP inlet 112 and a CP outlet 114. The CP inlet 112 is in gaseous communication with the working gas flow path 37 of the turbine 16 at a location upstream of the compressor section of the turbine 16. The CP outlet 114 is in gaseous communication with a bypass passage 56 of the gas turbine engine 10, which is defined in the outer nacelle 50 (not shown; see FIG. Figure 1 ) and the turbine 16.

[0076] It is worth noting that, as in the above-described embodiment, the CP inlet 112 is configured to receive the CP inlet 112 from the gas turbine engine 10 (see Figure 1 ) of the fan 38 compressed air flow, but is located upstream of any additional compression stage. Therefore, to help generate air flow through the cooling passage 110, the turbine 16 includes a device for promoting air flow through the cooling passage 110.

[0077] More specifically, for Figure 7 In the embodiment of the present invention, the gas turbine engine 10 further includes a variable exhaust assembly 122. The variable exhaust assembly 122 includes a variable exhaust duct 124 extending between a VB inlet 126 and a VB outlet 128.

[0078] The VB inlet 126 is in gaseous communication with the working gas flow path 37 at a location downstream of the CP inlet 112. Specifically, for the illustrated embodiment, the CP inlet 112 is in gaseous communication with the working gas flow path 37 at a location upstream of the compressor of the compressor section, and more specifically, upstream of the LP compressor 22 of the compressor section, and the VB inlet 126 is in gaseous communication with the working gas flow path 37 at a location downstream of the LP compressor 22. More specifically, for the illustrated embodiment, the VB inlet 126 is in gaseous communication with the working gas flow path 37 at a location upstream of the HP compressor 24 and is aligned with the compressor midframe 102 of the turbine 16 along the axial direction A of the gas turbine engine 10.

[0079] Still refer to Figure 7 , VB outlet 128 is in airflow communication with cooling passage 110. More specifically, for the illustrated embodiment, VB outlet 128 is in airflow communication with cooling passage 110 at a location downstream of heat exchanger 116 (heat exchanger 116 is in thermal communication with the airflow through cooling passage 110).

[0080] As from Figure 7 It should be appreciated that in the exemplary embodiment shown, substantially all of the airflow through the variable exhaust duct 124 (ie, at least 90% of the airflow through the variable exhaust duct 124 ) is provided to the cooling passage 110 through the VB outlet 128 in the illustrated embodiment.

[0081] However, it should be understood that in other embodiments, it may not be necessary to provide substantially all of the airflow through the variable exhaust duct 124 to the cooling passage 110 through the VB outlet 128. In this manner, the VB outlet 128 may be a first VB outlet, and the variable exhaust duct 124 may further include a second VB outlet 128'. The second VB outlet 128', as shown in dashed lines, may be in direct airflow communication with the bypass passage 56 (i.e., the airflow from the variable exhaust duct 124 may be provided to the bypass passage 56 without merging or mixing with any other airflow upstream of the bypass passage 56). In the illustrated embodiment, the variable exhaust duct 124 splits to extend to the first VB outlet 128 and the second VB outlet 128'.

[0082] It should be appreciated that providing airflow through variable bleed duct 124 to cooling passage 110 via VB outlet 128 may facilitate airflow through cooling passage 110 . In this manner, variable bleed assembly 122 may be a device for facilitating airflow through cooling passage 110 .

[0083] In order to adjust the amount of airflow through the cooling passage 110, the variable exhaust assembly 122 can change the amount of airflow provided therethrough to the cooling passage 110. Figure 7 In the embodiment shown, the variable exhaust assembly 122 includes a variable exhaust valve 130 for varying the amount of air flowing through the variable exhaust conduit 124. In the embodiment shown, the variable exhaust valve 130 is located at the upstream end of the variable exhaust conduit 124 and includes a VB inlet 126.

[0084] More specifically, for the embodiment shown, the variable exhaust assembly 122 further includes an actuator 132 coupled to the variable exhaust valve 130, the actuator 132 being configured to actuate the variable exhaust valve 130 about a pin 134 to pivot the variable exhaust valve 130 between a deployed position (shown) and a stowed position (not shown), and optionally various positions therebetween, as indicated by arrow 135. The variable exhaust valve 130 can be moved from the deployed position to the stowed position by rotating about the pin 134 in a clockwise direction in the illustrated view, such that the VB inlet 126 is no longer exposed to the working gas flow path 37.

[0085] The variable exhaust valve 130 can be moved between a fully extended position (shown) and a fully stowed position, and any suitable position therebetween (one or more partially extended positions), wherein in the fully extended position, the variable exhaust duct 124 extracts a maximum amount of airflow from the working gas flow path 37, and in the fully stowed position, the variable exhaust valve 130 extracts substantially no airflow from the working gas flow path 37 (i.e., less than 5% of the maximum amount of airflow extracted).

[0086] The airflow from the variable exhaust duct 124 may be provided to the cooling passage 110 from the VB outlet 128 in any suitable manner. For example, referring briefly to Figure 8 It should be understood that in certain exemplary embodiments, the VB outlet 128 and the cooling passage 110 at least partially form an ejector 136. Specifically, for Figure 8 In the embodiment of the present invention, the VB outlet 128 includes a fluid nozzle 138 configured to provide a relatively high-pressure fluid flow. Furthermore, the cooling passage 110 includes a nozzle section 140 comprising a converging inlet nozzle 142, a diffuser throat 144, and a diverging outlet diffuser 146 arranged in a serial order, wherein the fluid nozzle 138 of the VB outlet 128 is oriented to provide a relatively high-pressure fluid into the converging inlet nozzle 142. It should be appreciated that providing a high-pressure fluid flow through the fluid nozzle 138 using the downstream-located nozzle section 140 can encourage a relatively low-pressure fluid flow therethrough, thereby increasing the amount of airflow through the cooling passage 110.

[0087] Furthermore, it should be appreciated from the description herein and the accompanying drawings that the variable exhaust duct 124 having the VB inlet 126 and the VB outlet 128 may be the first variable exhaust duct 124 of the plurality of variable exhaust ducts 124 of the variable exhaust assembly 122. For example, with brief reference to Figure 9 , provides a schematic cross-sectional view showing the above reference Figure 7 The variable exhaust assembly 122 and the cooling passage 110 are described. As described above, the cooling passage 110 is an annular cooling passage. Further, the variable exhaust assembly 122 includes a plurality of variable exhaust ducts 124 spaced apart along the circumferential direction C of the gas turbine engine 10. Each of the variable exhaust ducts 124 is generally spaced apart from the working gas flow path 37 ( Figure 7 ) extends between a corresponding VB inlet 126 in airflow communication with the cooling channel 110 and a corresponding VB outlet 128 in airflow communication with the cooling channel 110. Figure 9 Each of the plurality of variable exhaust ducts 124 shown in FIG. 1 may be configured as described above with reference to FIG. Figure 7 and Figure 8 The depicted exemplary variable exhaust conduit 124 is configured in a similar manner.

[0088] Now refer back to Figure 7For the illustrated embodiment, it should be understood that the gas turbine engine 10 further includes a controller 150 operably coupled to the variable exhaust assembly 122, and one or more sensors 152. The one or more sensors 152 may be configured to sense data indicative of operating conditions of the gas turbine engine 10. For example, the one or more sensors 152 generally include a bypass duct sensor 152A (e.g., configured to sense one or more of a pressure, a temperature, or an airflow rate of the airflow through the bypass duct 56), an accessory system sensor 152B (e.g., configured to sense data indicative of a condition of the accessory system 108), and a turbine sensor 152C (e.g., configured to sense data indicative of an operating condition of the turbine, such as a speed sensor, a temperature sensor, a pressure sensor, etc.).

[0089] As mentioned above, Figure 7 The exemplary controller 150 shown in FIG. 1 is configured to receive sensed data from one or more sensors (sensors 152A, 152B, 152C for the illustrated embodiment) and may make control decisions for the variable exhaust assembly 122 based on the received data, for example.

[0090] In one or more exemplary embodiments, Figure 7 The controller 150 shown in the figure may be a stand-alone controller 150 for the variable exhaust assembly 122, or alternatively, may be integrated into one or more of a controller of the gas turbine engine 10 integrated with the variable exhaust assembly 122, a controller of an aircraft including the gas turbine engine 10 integrated with the variable exhaust assembly 122, and the like.

[0091] With particular reference to the operation of controller 150, in at least some embodiments, controller 150 may include one or more computing devices 154. Computing devices 154 may include one or more processors 154A and one or more memory devices 154B. One or more processors 154A may include any suitable processing device, such as a microprocessor, a microcontroller, an integrated circuit, a logic device, and / or other suitable processing device. One or more memory devices 154B may include one or more computer-readable media, including, but not limited to, non-transitory computer-readable media, RAM, ROM, a hard drive, a flash drive, and / or other memory devices.

[0092] One or more memory devices 154B may store information accessible by one or more processors 154A, including computer-readable instructions 154C executable by one or more processors 154A. Instructions 154C may be any set of instructions that, when executed by one or more processors 154A, cause one or more processors 154A to perform operations. In some embodiments, instructions 154C may be executed by one or more processors 154A to cause one or more processors 154A to perform operations, such as any of the operations and functions for which controller 150 and / or computing device 154 are configured, operations for operating variable exhaust assembly 122 (e.g., method 200) (as described herein), and / or any other operations or functions of one or more computing devices 154. Instructions 154C may be software written in any suitable programming language or implemented in hardware. Additionally and / or alternatively, instructions 154C may be executed in logically and / or virtually independent threads on one or more processors 154A. One or more memory devices 154B may also store data 154D accessible by one or more processors 154A. For example, data 154D may include data indicative of power flow, data indicative of engine / aircraft operating conditions, and / or any other data and / or information described herein.

[0093] The computing device 154 may also include a network interface 154E for communicating, for example, with other components of the variable exhaust assembly 122, the gas turbine engine 10 including the variable exhaust assembly 122, an aircraft including the gas turbine engine 10, and the like. For example, in the illustrated embodiment, as described above, the gas turbine engine 10 and / or the variable exhaust assembly 122 include one or more sensors for sensing data indicative of one or more parameters of the gas turbine engine, the variable exhaust assembly 122, the cooling passage 110, the accessory system 108, or a combination thereof. The controller 150 of the variable exhaust assembly 122 is operably coupled to the one or more sensors via, for example, the network interface 154E, such that the controller 150 can receive data indicative of various operating parameters sensed by the one or more sensors during operation. Furthermore, for the illustrated embodiment, the controller 150 is operably coupled to, for example, the actuator 132. In this manner, the controller 150 can be configured to vary the amount of airflow passing through the variable exhaust assembly 122 and into the cooling passage 110 in response to, for example, data sensed by the one or more sensors.

[0094] The network interface 154E may include any suitable components for interfacing with one or more networks, including, for example, a transmitter, a receiver, a port, a controller, an antenna, and / or other suitable components.

[0095] The technology discussed herein relates to computer-based systems and the actions taken by computer-based systems and the information sent to and from computer-based systems. Those skilled in the art will recognize that the inherent flexibility of computer-based systems allows for various possible configurations, combinations, and partitioning of tasks and functions between and within components. For example, the processes discussed herein can be implemented using a single computing device or multiple computing devices working in combination. Databases, memories, instructions, and application programs can be implemented on a single system or distributed across multiple systems. Distributed components can operate sequentially or in parallel.

[0096] For the illustrated embodiment, the controller 150 is configured to actuate the variable exhaust valve 130 to increase or decrease the amount of gas flowing through the variable exhaust duct 124. In certain exemplary aspects, the controller 150 may also be configured to receive data (e.g., sensed data from one or more sensors 152) indicative of operating conditions of the gas turbine engine 10. In certain exemplary aspects, the controller 150 may actuate the variable exhaust valve 130 in response to the data indicative of the operating conditions to increase or decrease the amount of gas flowing through the variable exhaust duct 124.

[0097] For example, in response to receiving data indicating an operating condition that may require a relatively high amount of cooling, the controller 150 may be configured to actuate the variable exhaust valve 130 to increase the amount of airflow through the variable exhaust duct 124 (e.g., move the variable exhaust valve 130 to a fully extended position), thereby increasing the amount of airflow through the cooling passage 110. Conversely, in response to receiving data indicating an operating condition that may require a relatively low amount of cooling, the controller 150 may be configured to actuate the variable exhaust valve 130 to decrease the amount of airflow through the variable exhaust duct 124 (e.g., move the variable exhaust valve 130 to a fully retracted position), thereby decreasing the amount of airflow through the cooling passage 110.

[0098] Now refer to Figure 10 , a gas turbine engine 10 according to another exemplary aspect of the present disclosure is provided. Figure 10 The exemplary gas turbine engine 10 may be configured as described above with reference to Figure 1 and Figure 2 The exemplary gas turbine engine 10 depicted is configured in a similar manner. Like or similar numbers may refer to like or similar parts.

[0099] For example, Figure 10The exemplary gas turbine engine 10 shown in FIG generally includes a turbine 16 having a casing 18 and defining an inlet 20 and a (portion of) a bypass passage 56 above the casing 18. The turbine 16 further defines an annular cooling passage, referred to herein simply as a cooling passage 110, extending between a CP inlet 112 and a CP outlet 114. The cooling passage 110 may extend substantially continuously in a circumferential direction C of the gas turbine engine 10 (e.g., at least 300 degrees, such as at least 330 degrees, such as 360 degrees, excluding various structural components extending therethrough).

[0100] The CP inlet 112 is in flow communication with the working gas flow path 37, the bypass passage 56, or both. Figure 2 Same as the example, Figure 10 The CP inlet 112 of the cooling passage 110 is shown in FIG. 1 and is in gas flow communication with the working gas flow path 37. In this manner, as in the above-described embodiment, the CP inlet 112 is configured to receive the working gas from the gas turbine engine 10 (see FIG. 1 ) prior to any additional compression stages. Figure 1 ) of the fan 38 to compress the airflow. This can avoid the work loss when utilizing the airflow.

[0101] As described above, gas turbine engine 10 includes accessory system 108. As with the exemplary embodiments described above, accessory system 108 may be an oil cooling system, a cooled cooling air system, an electric machine cooling system, or a combination thereof.

[0102] Additionally, the turbine 16 includes a heat exchanger 116 positioned in thermal communication with the cooling passage 110 (e.g., in thermal communication with the airflow through the cooling passage 110 during operation of the gas turbine engine 10) at a location between the CP inlet 112 and the CP outlet 114. Specifically, the heat exchanger 116 is positioned within and extends through the cooling passage 110. In this manner, the heat exchanger 116 may define a portion of the cooling passage 110. In this manner, the heat exchanger 116 may be configured to transfer heat from the fluid to the airflow through the cooling passage 110.

[0103] Further, for the illustrated embodiment, heat exchanger 116 is in thermal communication with accessory system 108 of gas turbine engine 10 for transferring heat from accessory system 108 to the airflow through cooling passage 110. In this manner, cooling passage 110 may provide cooling for accessory system 108.

[0104] It should be understood that for the illustrated embodiment, although a single heat exchanger 116 and a single accessory system 108 are depicted, in other exemplary embodiments, other suitable configurations may be provided (see, e.g., Figure 3 ).

[0105] Still refer to Figure 10 In an exemplary embodiment, the gas turbine engine 10 further includes an exhaust cooling system 160. The exhaust cooling system 160 defines an exhaust cooling (BC) inlet 162 and a duct assembly 164, which is in fluid communication with the BC inlet 162 for receiving airflow from the BC inlet 162. The BC inlet 162 is in airflow communication with the cooling passage 110 at a position between the CP inlet 112 and the CP outlet 114. In certain exemplary embodiments, the exhaust cooling system 160 may be a clearance control system (see, for example, the above reference Figure 1 The gap control system 84) can be a ventilation cooling system under the fairing (see, for example, the above reference Figure 1 The under-fairing ventilation cooling system 82), or a combination thereof.

[0106] More specifically, for the illustrated embodiment, the BC inlet 162 of the exhaust cooling system 160 is a first BC inlet 162A, and the exhaust cooling system 160 also defines a second BC inlet 162B and a third BC inlet 162C.

[0107] For the illustrated embodiment, the first BC inlet 162A of the exhaust cooling system 160 is located downstream of the CP inlet 112 and upstream of the heat exchanger 116. In this manner, the first BC inlet 162A can be configured to receive airflow from within the cooling passage 110 at a lower temperature (relative to the downstream location within the cooling passage 110) because the airflow has not yet passed through the heat exchanger 116 to exchange heat with the fluid.

[0108] Furthermore, for the illustrated embodiment, the second BC inlet 162B of the exhaust cooling system 160 is co-located with the heat exchanger 116. More specifically, the second BC inlet 162B is in airflow communication with the cooling passage 110 at a location aligned (along the axial direction A) with the location of the cooling passage 110, wherein the heat exchanger 116 is in thermal communication with the airflow through the cooling passage 110. Positioning the second BC inlet 162B at such a location allows the exhaust cooling system 160 to receive an airflow having desired aerodynamic properties, thermal properties, or a combination thereof.

[0109] Furthermore, for the illustrated embodiment, the third BC inlet 162C of the exhaust cooling system 160 is located downstream of the heat exchanger 116 and upstream of the CP outlet 114. In this manner, the third BC inlet 162C can be configured to receive airflow from within the cooling passage 110 at a higher temperature (relative to an upstream location within the cooling passage 110). However, providing airflow to the exhaust cooling system 160 through the third BC inlet 162C can allow for reduced back pressure within the cooling passage 110, for example, increasing the effectiveness of the heat exchanger 116.

[0110] However, it should be understood that although Figure 10 The exemplary exhaust cooling system 160 shown in FIG includes each of the first BC inlet 162A, the second BC inlet 162B, and the third BC inlet 162C, but in other exemplary embodiments, the exhaust cooling system 160 may include only the first BC inlet 162A, may include only the second BC inlet 162B, may include only the third BC inlet 162C, or may include a combination of only two of the first BC inlet 162A, the second BC inlet 162B, and the third BC inlet 162C.

[0111] Furthermore, it should be understood that in one or more of these exemplary embodiments, the BC inlets 162 described at particular locations along the cooling passage 110 (e.g., the first BC inlet 162A, the second BC inlet 162B, and the third BC inlet 162C) may include a plurality of BC inlets 162 spaced apart along the circumferential direction C to allow extraction of airflow through the cooling passage 110 along an annulus of the cooling passage 110 (as will be described below with reference to FIG. Figure 11 described in more detail).

[0112] Specifically, now refer to Figure 11 , provides along Figure 10 Line 11-11 Figure 10 110, it should be understood that the BC inlet 162 of the exhaust cooling system 160, and more specifically, the first BC inlet 162A of the exhaust cooling system 160, is one of a plurality of first BC inlets 162A spaced apart along the circumferential direction C of the gas turbine engine 10. Each of the plurality of first BC inlets 162A is in gas flow communication with the cooling channel 110 at a location between the CP inlet 112 and the CP outlet 114, and more specifically, in gas flow communication with the cooling channel 110 at a location downstream of the CP inlet 112 and upstream of the heat exchanger 116.

[0113] Notably, in the illustrated embodiment, the duct assembly 164 of the exhaust cooling system 160 includes a plurality of first ducts 166, one or more second ducts 168, and a third duct 170. More specifically, the plurality of first ducts 166 are a plurality of radial ducts extending from the respective plurality of first BC inlets 162A to receive airflow from the cooling passage 110 provided through the respective plurality of first BC inlets 162A. In the illustrated embodiment, the one or more second ducts 168 are a plurality of circumferential ducts in airflow communication with the plurality of radial ducts, receiving airflow from each of the plurality of radial ducts. Furthermore, the third duct 170 is an axial duct that receives airflow from the plurality of circumferential ducts and delivers the received airflow to a rearward location of the exhaust cooling system 160 (e.g., to one or more under-cowl areas in embodiments where the exhaust cooling system 160 is an under-cowl ventilation cooling system; to a shroud cooling assembly in embodiments where the exhaust cooling system 160 is a clearance control system; etc.).

[0114] like Figure 11 As shown, the exhaust cooling system 160 may include one or more features for promoting airflow into and through the plurality of BC inlets 162. Specifically, for the illustrated embodiment, the exhaust cooling system 160 further includes a plurality of scoops 171, each scoop 171 being positioned adjacent a corresponding BC inlet 162 of the plurality of BC inlets 162 to direct airflow into and through the plurality of BC inlets 162. The scoops 171 define a height 174 in the radial direction R that is greater than or equal to 2% of a height 176 of the cooling channel 110 at the same axial position and less than or equal to 25% of the height 176 of the cooling channel 110 at the same axial position.

[0115] although Figure 11 A plurality of discrete fixed scoops 171 are depicted in FIG, but in other embodiments, other configurations may be provided to divert airflow into the BC inlet 162. For example, in other embodiments, the exhaust cooling system 160 may instead include one or more scoops 171 extending along the circumferential direction C (similar to, for example, the scoops described above with reference to FIG). Figure 4 The annular scoop 171 described above), one or more variable scoops 171 that can be expanded and retracted, etc.

[0116] Briefly, refer back to Figure 10 The exhaust cooling system 160 is shown to further include a valve in airflow communication with the duct assembly 164 to regulate the amount of airflow through the duct assembly 164 of the exhaust cooling system 160. Specifically, for the illustrated embodiment, the valve is a variable flow valve 178, and the gas turbine engine 10 includes a controller 150 in operable communication with the variable flow valve 178. Figure 10 The controller 150 in the above reference Figure 7The exemplary controller 150 described is configured in a similar manner. In this manner, the controller 150 can be configured to vary the amount of airflow through the exhaust cooling system 160 in response to, for example, operating conditions of the gas turbine engine 10, data sensed from one or more sensors of the gas turbine engine 10, and the like.

[0117] However, it should be understood that in other exemplary embodiments, the exhaust cooling system 160 may additionally or alternatively include one or more similar valves at other locations in the conduit assembly 164, such as with, for example, the plurality of first conduits 166, the one or more second conduits 168 (see FIG. Figure 11 ), a third conduit 170 at one or more alternative locations, or a combination thereof, is connected to the airflow.

[0118] Furthermore, in other exemplary embodiments, the gas turbine engine 10 may have various other configurations. For example, referring now to Figure 12 , a gas turbine engine 10 having a cooling passage 110 according to another exemplary embodiment of the present disclosure is provided. Figure 12 The exemplary gas turbine engine 10 and cooling passage 110 may be described with reference to Figure 10 The exemplary gas turbine engine 10 and cooling passages 110 depicted are configured in a similar manner. Like or similar numbers may refer to like or similar parts.

[0119] For example, Figure 12 The exemplary gas turbine engine 10 generally includes an exhaust cooling system 160. However, for the illustrated embodiment, the exhaust cooling system 160 is a first exhaust cooling system 160A, and the BC inlet 162 of the first exhaust cooling system 160A is a first BC inlet 162A.

[0120] The exemplary gas turbine engine 10 further includes a second exhaust cooling system 160B, which defines a second BC inlet 162B in gaseous communication with the cooling passage 110 at a second location between the CP inlet 112 and the CP outlet 114. The second BC inlet 162B is located downstream of the first BC inlet 162A. In addition, the illustrated exemplary gas turbine engine 10 further includes a third exhaust cooling system 160C, which defines a third BC inlet 162C in gaseous communication with the cooling passage 110 at a third location between the CP inlet 112 and the CP outlet 114. The third BC inlet 162C is located downstream of the second BC inlet 162B. In short, similar to the above reference Figure 10Like the BC inlet 162 described above, the first BC inlet 162A, the second BC inlet 162B, and the third BC inlet 162C may each represent a plurality of first BC inlets 162A, second BC inlets 162B, or third BC inlets 162C spaced apart along the circumferential direction C of the gas turbine engine 10 (see FIG. Figure 11 ).

[0121] For the embodiment shown, first BC inlet 162A is positioned between CP inlet 112 and heat exchanger 116 in thermal communication with cooling channel 110. Second BC inlet 162B is co-located with heat exchanger 116. Third BC inlet 162C is downstream of heat exchanger 116 and upstream of CP outlet 114.

[0122] However, it should be understood that in other exemplary embodiments, two or more of the BC inlets 162 may be positioned at similar locations (e.g., between the CP inlet 112 and the heat exchanger 116, co-located with the heat exchanger 116, or between the heat exchanger 116 and the CP outlet 114).

[0123] Notably, for the illustrated embodiment, each of the first exhaust cooling system 160A, the second exhaust cooling system 160B, and the third exhaust cooling system 160C includes a separate duct assembly 164 (i.e., a first duct assembly 164A, a second duct assembly 164B, and a third duct assembly 164C), and valves 178A, 178B, 178C in airflow communication with the respective duct assemblies 164A, 164B, 164C to, for example, regulate the amount of airflow therethrough.

[0124] Furthermore, in other exemplary embodiments of the present disclosure, the gas turbine engine 10 may have other configurations. For example, referring now to Figure 13 , a gas turbine engine 10 having a cooling passage 110 according to another exemplary embodiment of the present disclosure is provided. Figure 13 The exemplary gas turbine engine 10 and cooling passage 110 may be described with reference to Figure 10 The exemplary gas turbine engine 10 and cooling passages 110 depicted are configured in a similar manner. Like or similar numbers may refer to like or similar parts.

[0125] For example, Figure 13 The exemplary gas turbine engine 10 generally includes a cooling passage 110 extending between a CP inlet 112 and a CP outlet 114. The gas turbine engine 10 further includes an exhaust cooling system 160, which may be similar to that described herein above, for example with reference to Figures 10 to 13 One or more of the described exemplary exhaust cooling systems 160 are configured in a similar manner.

[0126] The CP inlet 112 is in flow communication with the working gas flow path 37 of the turbine 16 of the gas turbine engine 10, the bypass passage 56 of the gas turbine engine 10, or both. Figure 13 In the exemplary embodiment of the present invention, the CP inlet 112 of the cooling passage 110 is in airflow communication with the bypass passage 56. It is worth noting that, as in the above-described embodiment, this configuration allows the CP inlet 112 to receive the gas turbine engine 10 (see Figure 1 ) of the fan section 14 without performing additional work / pressurization to avoid exhausting an airflow that has applied excessive work.

[0127] More specifically, the turbine 16 of the gas turbine engine 10 includes a casing 18 having an outer surface 180. The turbine 16 further includes a cowling 182 positioned outside the casing 18 in a radial direction R of the gas turbine engine 10 and, together with the casing 18, at least partially defining the cooling passage 110. The cowling 182 extends from a CP inlet 112 to a CP outlet 114. For the illustrated embodiment, the heat exchanger 116 is coupled to the cowling 182.

[0128] It should be understood that the term “casing” as used herein broadly encompasses the outermost structure of the turbine 16 (without the cowling 182 ) and may be formed from multiple components coupled together.

[0129] The housing 18 defines a front stationary point 184 located directly forward of the CP inlet 112 and a rear stationary point 186 located directly rearward of the CP outlet 114. The term "stationary point" refers to the position of the housing 18 in the axial and radial planes (i.e., in the plane defined by the axial direction A and the radial direction R, such as Figure 13 18 (the plane of view shown in FIG), along the outer surface 180 of the casing 18, wherein tangents to the outer surface 180 of the casing 18 in the axial and radial planes are parallel to the centerline 12 of the gas turbine engine 10. Additionally, the term "directly forward" relative to the position of the forward stationary point 184 relative to the CP inlet 112 refers to the first stationary point forward of the CP inlet 112. Similarly, the term "directly aft" relative to the position of the aft stationary point 186 relative to the CP outlet 114 refers to the first stationary point aft of the CP outlet 114.

[0130] Furthermore, the housing 18 defines a reference line 188 extending in the axial and radial planes shown from the forward stationary point 184 to the aft stationary point 186. A cowling 182 of the turbine 16, which at least partially defines the cooling passage 110, is positioned at least partially outboard of the reference line 188 in the radial direction R. Specifically, for the exemplary embodiment shown, a forward end 190 of the cowling 182 is positioned completely outboard of the reference line 188 in the radial direction R, and an aft end 192 of the cowling 182 is also positioned completely outboard of the reference line 188 in the radial direction R.

[0131] Furthermore, cooling passage 110 may be configured to receive less than 10% of the airflow through bypass passage 56, such as less than 8% of the airflow through bypass passage 56, and at least 0.05% of the airflow through bypass passage 56. These percentages may be calculated based on the mass flow rate of the airflow through bypass passage 56 during cruise operating conditions of gas turbine engine 10.

[0132] Such a configuration ensures that the required airflow is provided through the cooling passage 110 for cooling accessory systems of the gas turbine engine 10 (using the heat exchanger 116 ) and for providing cooling airflow to the exhaust cooling system 160 of the gas turbine engine 10 .

[0133] Further aspects are provided by the subject matter of the following clauses:

[0134] A gas turbine engine comprises: a fan assembly, the fan assembly including a fan; a turbine, the turbine being drivingly coupled to the fan and including a compressor section, a combustion section, and a turbine section arranged in a serial flow order and partially defining a working gas flow path, the gas turbine engine defining a bypass passage above the turbine, the turbine defining an annular cooling passage extending between a CP inlet and a CP outlet, the CP inlet being in communication with the working gas flow path airflow, and the CP outlet being in communication with the bypass passage airflow; and a variable exhaust assembly, the variable exhaust assembly comprising a variable exhaust duct extending between a VB inlet and a VB outlet, the VB inlet being in communication with the working gas flow path airflow at a position downstream of the CP inlet, and the VB outlet being in communication with the annular cooling passage airflow for promoting airflow through the cooling passage.

[0135] A gas turbine engine according to any of the preceding clauses, wherein the compressor section includes a compressor, wherein the CP inlet is connected to the working gas flow path airflow at a position upstream of the compressor, and wherein the VB inlet is connected to the working gas flow path airflow at a position downstream of the compressor.

[0136] A gas turbine engine as claimed in any preceding clause, wherein the compressor is a low pressure compressor.

[0137] A gas turbine engine as claimed in any preceding clause, wherein the compressor section further comprises a high pressure compressor, wherein the VB inlet is in gaseous communication with the working gas flow path at a location upstream of the high pressure compressor.

[0138] The gas turbine engine according to any of the preceding clauses, wherein the variable exhaust assembly comprises a variable exhaust valve for varying the amount of gas flowing through the variable exhaust duct.

[0139] The gas turbine engine according to any of the preceding items further includes: a controller operably connected to the variable exhaust valve, wherein the controller is configured to actuate the variable exhaust assembly to increase the air flow through the variable exhaust duct in response to the operating conditions of the gas turbine engine, thereby increasing the air flow through the annular cooling channel.

[0140] A gas turbine engine according to any preceding clause, wherein the VB outlet at least partially forms an ejector.

[0141] A gas turbine engine as claimed in any preceding clause, wherein substantially all of the airflow through the variable exhaust duct is provided to the cooling passage through the VB outlet.

[0142] The gas turbine engine according to any preceding clause, wherein the VB outlet is a first VB outlet, and wherein the variable exhaust duct further comprises a second VB outlet, wherein the second VB outlet is in direct airflow communication with the bypass passage.

[0143] A gas turbine engine as claimed in any preceding clause, wherein the turbine comprises a heat exchanger in thermal communication with the gas flow through the cooling passage.

[0144] A gas turbine engine according to any preceding clause, wherein the VB outlet is in airflow communication with the cooling passage at a location downstream of the heat exchanger.

[0145] A gas turbine engine according to any preceding clause, wherein the gas turbine engine defines a circumferential direction, and wherein the heat exchanger is a first heat exchanger of a plurality of heat exchangers arranged along the circumferential direction within the annular cooling passage.

[0146] A method of operating a gas turbine engine including a fan assembly and a turbine drivingly coupled to a fan of the fan assembly, the method comprising: receiving data indicative of an operating condition of the gas turbine engine; and, in response to the received data, varying a variable exhaust gas flow rate provided to an annular cooling passage through a variable exhaust duct, the annular cooling passage extending between a CP inlet in airflow communication with a working gas flow path of the turbine and a CP outlet in airflow communication with a bypass passage of the gas turbine engine.

[0147] The method of any preceding clause, wherein the operating condition is a low fan power operating condition, and wherein changing the variable exhaust gas flow rate provided to the annular cooling passage through the variable exhaust duct comprises increasing the variable exhaust gas flow rate provided to the annular cooling passage through the variable exhaust duct.

[0148] A method according to any preceding clause, wherein the low fan power operating condition is a ground idling operating condition or a flight idling descent operating condition.

[0149] A method according to any preceding clause, wherein the operating condition is indicative of ambient temperature.

[0150] The method of any preceding clause, wherein the operating condition indicates that the ambient temperature is greater than a threshold value, and wherein changing the variable exhaust gas flow rate provided to the annular cooling passage through the variable exhaust duct comprises increasing the variable exhaust gas flow rate provided to the annular cooling passage through the variable exhaust duct.

[0151] The method according to any of the preceding clauses, wherein the variable exhaust assembly comprises a variable exhaust valve for varying the amount of air passing through the variable exhaust duct.

[0152] A method as in any preceding clause, wherein the turbine comprises a heat exchanger in thermal communication with the air flow through the cooling passage.

[0153] A method according to any preceding clause, wherein the VB outlet is in fluid communication with the cooling channel at a location downstream of the heat exchanger.

[0154] 1. A gas turbine engine comprising: a fan assembly, the fan assembly including a fan; a turbine, the turbine being drivingly coupled to the fan and including a compressor section, a combustion section, and a turbine section arranged in a serial flow order and partially defining a working gas flow path, the gas turbine engine defining a bypass passage above the turbine, the turbine defining an annular cooling passage (CP), the annular CP extending between a CP inlet and a CP outlet, the CP inlet being in airflow communication with the working gas flow path, the bypass passage, or both; an accessory system; a heat exchanger positioned to be in thermal communication with the annular cooling passage at a position between the CP inlet and the CP outlet, the heat exchanger being in thermal communication with the accessory system; and an exhaust cooling system, the exhaust cooling system defining a BC inlet, the BC inlet being in airflow communication with the annular cooling passage at a position between the CP inlet and the CP outlet.

[0155] A gas turbine engine according to any preceding clause, wherein the accessory system is in thermal communication with the accessory system for cooling an oil cooling system, a cooled cooling air system, an electric machine cooling system, or a combination thereof.

[0156] The gas turbine engine according to any of the preceding clauses, wherein the exhaust cooling system is a clearance control system, an under-cowl ventilation cooling system, or a combination thereof.

[0157] A gas turbine engine according to any preceding clause, wherein the BC inlet of the exhaust cooling system is located downstream of the CP inlet and upstream of the heat exchanger.

[0158] A gas turbine engine as claimed in any preceding clause, wherein the BC inlet of the exhaust cooling system is co-located with the heat exchanger.

[0159] A gas turbine engine according to any preceding clause, wherein the BC inlet of the exhaust cooling system is located downstream of the heat exchanger and upstream of the CP outlet.

[0160] A gas turbine engine according to any of the preceding clauses, wherein the exhaust cooling system is a first exhaust cooling system, wherein the BC inlet is a first BC inlet, and wherein the gas turbine engine further comprises: a second exhaust cooling system, the second exhaust cooling system defining a second BC inlet, the second BC inlet being in airflow communication with the annular cooling channel at a second position between the CP inlet and the CP outlet, wherein the second BC inlet is located downstream of the first BC inlet.

[0161] A gas turbine engine according to any of the preceding items, wherein the BC inlet of the exhaust cooling system is the first BC inlet of a plurality of BC inlets spaced apart in a circumferential direction of the gas turbine engine, wherein each of the plurality of BC inlets is in airflow communication with the annular cooling channel at the position between the CP inlet and the CP outlet.

[0162] The gas turbine engine of any preceding clause, wherein the BC inlet of the exhaust cooling system is a first BC inlet, wherein the exhaust cooling system further defines a second BC inlet co-located with or downstream of the heat exchanger and upstream of the CP outlet.

[0163] A gas turbine engine according to any preceding clause, wherein the CP inlet is in gas flow communication with the bypass passage.

[0164] A gas turbine engine according to any preceding clause, wherein the turbine comprises a cowling and a casing defining an outer surface, wherein the cowling and the outer surface of the casing at least partially define the annular cooling passage.

[0165] A gas turbine engine according to any of the preceding clauses, wherein the casing defines a forward stationary point immediately forward of the CP inlet, an aft stationary point immediately aft of the CP outlet, and a reference line extending from the forward stationary point to the aft stationary point, and wherein the fairing is at least partially located outside of the reference line.

[0166] A gas turbine engine according to any of the preceding clauses, wherein the fairing defines a front end and a rear end, wherein the front end is positioned completely outside the reference line along a radial direction of the gas turbine engine, and wherein the rear end is also positioned completely outside the reference line along the radial direction.

[0167] A gas turbine engine as claimed in any preceding clause, wherein the CP inlet is in gas flow communication with the working gas flow path.

[0168] The gas turbine engine according to any of the preceding clauses, wherein the exhaust cooling system further comprises a duct assembly in airflow communication with the BC inlet for receiving airflow from the BC inlet.

[0169] The gas turbine engine according to any preceding clause, wherein the exhaust cooling system further comprises a valve in airflow communication with the duct assembly to regulate the amount of airflow through the duct assembly of the exhaust cooling system.

[0170] The gas turbine engine of any preceding clause, further comprising a controller, wherein the valve is a variable flow valve, and wherein the controller is operatively coupled to the variable flow valve.

[0171] A gas turbine engine according to any of the preceding clauses, wherein the BC inlet is one of a plurality of BC inlets spaced apart in a circumferential direction of the gas turbine engine, wherein the duct assembly comprises a plurality of radial ducts, wherein each radial duct is in airflow communication with a corresponding BC inlet of the plurality of BC inlets.

[0172] The gas turbine engine of any preceding clause, wherein the turbine defines an under-fairing region, wherein the duct assembly further comprises an axial duct in gas flow communication with the plurality of radial ducts, wherein the axial duct extends to the under-fairing region.

[0173] The gas turbine engine of any preceding clause, wherein the exhaust cooling system is a clearance control system having a shroud cooling assembly, wherein the duct assembly further comprises an axial duct in airflow communication with the plurality of radial ducts, wherein the axial duct extends to the shroud cooling assembly.

[0174] This written description uses examples to disclose the disclosure, including the best mode, and also to enable any person skilled in the art to practice the disclosure, including making and using any devices or systems and performing any incorporated methods. The patentable scope of the disclosure is defined by the claims, and may include other examples that occur to those skilled in the art. Such other examples are intended to be within the scope of the claims if they include structural elements that do not differ from the literal language of the claims, or if they include equivalent structural elements with insubstantial differences from the literal language of the claims.

Claims

1. A gas turbine engine, characterized in that: include: A fan assembly, the fan assembly comprising a fan; a turbine drivingly coupled to the fan and comprising a compressor section, a combustion section, and a turbine section arranged in serial flow order and partially defining a working gas flow path, the gas turbine engine defining a bypass passage above the turbine, the turbine defining an annular cooling passage (CP), the annular CP extending between a CP inlet and a CP outlet, the CP inlet being in flow communication with the working gas flow path, the bypass passage, or both; Accessory system; a heat exchanger positioned in thermal communication with the annular cooling passage at a location between the CP inlet and the CP outlet, the heat exchanger being in thermal communication with the accessory system; as well as An exhaust cooling-BC system defines a BC inlet in airflow communication with the annular cooling passage at a location between the CP inlet and the CP outlet.

2. The gas turbine engine according to claim 1, wherein: The accessory system is in thermal communication with the accessory system for cooling an oil cooling system, a cooling air system, a motor cooling system, or a combination thereof.

3. The gas turbine engine according to claim 1, wherein: The exhaust cooling system is a gap control system, an under-fairing ventilation cooling system or a combination thereof.

4. The gas turbine engine according to claim 1, wherein: The BC inlet of the exhaust cooling system is located downstream of the CP inlet and upstream of the heat exchanger.

5. The gas turbine engine according to claim 1, wherein The BC inlet of the exhaust cooling system is located at the same position as the heat exchanger.

6. The gas turbine engine according to claim 1, wherein: The BC inlet of the exhaust cooling system is located downstream of the heat exchanger and upstream of the CP outlet.

7. The gas turbine engine according to claim 1, wherein: wherein the exhaust cooling system is a first exhaust cooling system, wherein the BC inlet is a first BC inlet, and wherein the gas turbine engine further comprises: A second exhaust cooling system defines a second BC inlet in airflow communication with the annular cooling passage at a second location between the CP inlet and the CP outlet, wherein the second BC inlet is located downstream of the first BC inlet.

8. The gas turbine engine according to claim 1, wherein: The BC inlet of the exhaust cooling system is a first BC inlet among a plurality of BC inlets spaced apart in a circumferential direction of the gas turbine engine, wherein each of the plurality of BC inlets is in airflow communication with the annular cooling channel at the position between the CP inlet and the CP outlet.

9. The gas turbine engine according to claim 1, wherein: The BC inlet of the exhaust cooling system is a first BC inlet, wherein the exhaust cooling system further defines a second BC inlet co-located with the heat exchanger or downstream of the heat exchanger and upstream of the CP outlet.

10. The gas turbine engine according to claim 1, wherein: The CP inlet is in airflow communication with the bypass channel.

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