Cooling air delivery components

Abstract

A gas turbine engine includes a combustion section, a turbine section, and a compressor section. The combustion section includes a combustor housing, a burner, a cooling duct, and an outer duct. The combustor housing at least partially defines a diffuser cavity and a fluid inlet. The burner is disposed in the diffuser cavity. The cooling duct is in fluid communication with the fluid inlet in the combustor housing and is configured to carry a cooling air flow. The outer duct surrounds at least a portion of the cooling duct and extends along a portion of the entire length of the cooling duct. The outer duct defines a gap with the cooling duct and is configured to carry a buffer air flow. The turbine section is disposed downstream of the combustion section. The cooling duct is in fluid communication with the turbine section.

Description

Technical Field

[0001] The present disclosure generally relates to cooling air delivery systems for gas turbine engines. Background Art

[0002] A gas turbine engine generally comprises 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.

[0003] In gas turbine engines, thermal management systems are incorporated to cool certain components and prevent damage due to overheating. In existing thermal management systems, air ducts may be provided through the combustion chamber of the gas turbine engine to provide cooling air flow to the turbine section of the gas turbine engine. BRIEF DESCRIPTION OF THE DRAWINGS

[0004] A full and effective 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 cross-sectional view of a compressor section and a combustion section of an exemplary gas turbine showing a first cooling tube assembly according to an exemplary aspect of the present disclosure.

[0007] Figure 3 is a cross-sectional view of a compressor section and a combustion section of an exemplary gas turbine showing a cooling tube assembly according to another exemplary aspect of the present disclosure.

[0008] Figure 4 is a cross-sectional view of a compressor section and a combustion section of an exemplary gas turbine showing a cooling tube assembly according to yet another exemplary aspect of the present disclosure.

[0009] Figure 5 is a cross-sectional view of a compressor section and a combustion section of an exemplary gas turbine showing a cooling tube assembly according to yet another exemplary aspect of the present disclosure.

[0010] Figure 6 is a cross-sectional view of a compressor section and a combustion section of an exemplary gas turbine showing a cooling tube assembly according to yet another exemplary aspect of the present disclosure.

[0011] Figure 7 is a cross-sectional view of a compressor section and a combustion section of an exemplary gas turbine showing a cooling tube assembly according to yet another exemplary aspect of the present disclosure.

[0012] Figure 8 is a cross-sectional view of a compressor section and a combustion section of an exemplary gas turbine showing a cooling tube assembly according to yet another exemplary aspect of the present disclosure. DETAILED DESCRIPTION

[0013] 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 reference numerals in the drawings and the description have been used to refer to like or similar parts of the disclosure.

[0014] The word "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 preferred or advantageous over other implementations. Furthermore, all embodiments described herein are to be considered exemplary unless expressly stated otherwise.

[0015] 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.

[0016] 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, while the aft position refers to the position closer to the engine nozzle or exhaust.

[0017] The terms "upstream" and "downstream" refer to relative directions of fluid flow in a fluid path. For example, "upstream" refers to the direction from which the fluid is flowing, and "downstream" refers to the direction to which the fluid is flowing.

[0018] The terms “coupled,” “fixed,” “attached,” and the like refer to both direct coupling, fixing, or attachment, as well as indirect coupling, fixing, or attachment through one or more intermediate components or features, unless otherwise indicated herein.

[0019] The singular forms "a," "an," and "the" include plural references unless the context clearly dictates otherwise.

[0020] Approximate language used throughout the specification and claims is applied to modify any quantitative expression that can be allowed to vary without causing changes in its related basic function. Therefore, the values modified by one or more terms such as "about", "approximately" and "substantially" are not limited to the specified precise values. In at least some cases, approximate language can correspond to the accuracy of the instrument used to measure the value, or the accuracy of the method or machine used to construct or manufacture parts and / or systems. For example, approximate language can refer to being within 1%, 2%, 4%, 10%, 15% or 20%. These approximate margins can be applied to a single value, any endpoint or two endpoints of a defined numerical range, and / or the margin of the range between the endpoints.

[0021] Here and throughout the specification and claims, range limitations are combined and interchanged, and unless context or language indicates otherwise, such ranges are identified and include all sub-ranges contained therein. For example, all ranges disclosed herein are inclusive of the endpoints, and the endpoints are independently combinable with each other.

[0022] The terms "low" and "high," or their respective comparative degrees (e.g., lower, higher, as applicable), when used with reference to a compressor, turbine, shaft or spool component, etc., each refer to relative speeds within the engine, unless otherwise specified. For example, a "low turbine" or "low-speed turbine" defines a component that is configured to operate at a lower speed (e.g., maximum allowable speed) than a "high turbine" or "high-speed turbine" at the engine.

[0023] The term “turbomachine” or “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.

[0024] 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, and the like.

[0025] 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 assembly. In certain exemplary 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.

[0026] The present disclosure generally relates to a thermal management system for a gas turbine engine having a compressor section, a combustion section, and a turbine section. Air ducts passing through the combustion chamber of a gas turbine engine may provide cooling air to the turbine section of the gas turbine engine. Such air ducts may be susceptible to thermal contamination within the combustion chamber. This thermal contamination of the cooling air can increase the thermal energy of the cooling air, thereby reducing the cooling effectiveness of the cooling air as it reaches its ultimate thermal energy transfer destination. Therefore, the inventors of the present disclosure have discovered that improvements to these air ducts would be beneficial.

[0027] The present disclosure proposes a nested tube configuration that forms a double wall for running a buffer air layer around a cooled cooling air delivery duct. The buffer air, delivered through the boundary layer confined between the double walls, reduces the amount of thermal contamination of the cooled cooling air passing through the cooled cooling air delivery duct. The buffer air can then be delivered to other cavities or components of the engine (e.g., the high-pressure compressor rear cavity or components or cavities of the engine's high-pressure turbine) for thermal management.

[0028] As described herein, the disclosed system provides an isolated delivery loop for a cooling air flow that reduces the amount of thermal contamination caused by the high levels of thermal energy present in the combustion section. Reducing the thermal contamination of the cooling air flow can reduce the flow rate of the cooling air flow required for cooling, thereby enabling a reduction in the size of the source of the cooling air flow (e.g., a cooling air heat exchanger, another heat exchanger, or another thermal management device). Additionally or alternatively, reducing the thermal contamination of the cooling air flow can increase the cooling of the end component receiving the cooling air flow.

[0029] Furthermore, in at least certain exemplary aspects, a buffer air flow can also be used by the high-pressure compressor via the rear cavity. The buffer air flow can be used to reduce the temperature of the air within the rear cavity, resulting in cooler air being delivered to the high-pressure compressor for cooling purposes. Since this cooler cooling air is delivered to the high-pressure compressor via the rear cavity, a dedicated cooling air flow from a heat exchanger (which would otherwise be required) can be eliminated due to the cooling advantage gained by utilizing the buffer air flow as cooling air for the high-pressure compressor. In this manner, the disclosed system can result in increased rotor temperature, improved cooling efficiency, improved combustion efficiency, or a combination thereof, for the high-pressure compressor and high-pressure turbine of the engine.

[0030] Referring now to the drawings, wherein like numerals refer to like elements throughout, Figure 1 is a schematic cross-sectional view of a propulsion system 10 according to an exemplary embodiment of the present disclosure. More specifically, Figure 1In an embodiment of the present invention, the propulsion system 10 includes a gas turbine engine 12. However, it should be understood that the term "gas turbine engine 12" may also refer to a high-bypass turbofan jet engine. In one example, the turbofan engine 12 may be a high-bypass turbofan jet engine. Figure 1 As shown, the turbofan engine 12 defines an axial direction A (extending parallel to a longitudinal centerline 14 for reference) and a radial direction R. Generally speaking, the turbofan engine 12 includes a fan section 16 and a turbine 18 disposed downstream of the fan section 16 .

[0031] The depicted exemplary turbomachine 18 generally includes a substantially tubular outer casing 20 defining an annular inlet 22. The outer casing 20 encloses, in series flow sequence / relationship: a compressor section, including a supercharger or low-pressure compressor 24 (“LP compressor 24”) and a high-pressure compressor 26 (“HP compressor 26”); a combustion section 28; a turbine section, including a high-pressure turbine 30 (HP turbine 30”) and a low-pressure turbine 32 (“LP turbine 32”); a high-pressure shaft or spool 34 (“HP spool 34”) drivingly connecting the HP turbine 30 to the HP compressor 26. A low-pressure shaft or spool 36 (“LP spool 36”) drivingly connecting the LP turbine 32 to the low-pressure compressor 24.

[0032] For the depicted embodiment, fan section 16 includes a variable pitch fan 38 having a plurality of fan blades 40 coupled to a disk 42 in a spaced-apart manner. As depicted, fan blades 40 generally extend outwardly from disk 42 in a radial direction R. Each fan blade 40 is rotatable relative to disk 42 about a pitch axis P by being operably coupled to a suitable actuation member 44 that is configured to collectively, e.g., uniformly, change the pitch of fan blades 40. Fan blades 40, disk 42, and actuation members 44 are rotatable together about longitudinal centerline 14 via LP spool 36 through power gearbox 46. Power gearbox 46 includes a plurality of gears for reducing the rotational speed of LP spool 36 to a more efficient fan rotational speed.

[0033] Still refer to Figure 1 In an exemplary embodiment of the present invention, the disk 42 is covered by a rotatable forward hub 48 having an aerodynamic profile to facilitate airflow through the plurality of fan blades 40. Additionally, the fan section 16 includes an annular fan casing or outer nacelle 50 that circumferentially surrounds the variable pitch fan 38 and / or at least a portion of the turbine 18. It should be understood that in some embodiments, the nacelle 50 is configured to be supported relative to the turbine 18 by a plurality of circumferentially spaced outlet guide vanes 52. Furthermore, a downstream section 54 of the nacelle 50 extends over an outer portion of the turbine 18 to define a bypass airflow passage 56 therebetween.

[0034] During operation of turbofan engine 12, a volume of air 58 enters turbofan engine 12 through nacelle 50 and / or associated inlet 60 of fan section 16. As volume of air 58 passes through fan blades 40, a first portion of air 58, as indicated by arrow 62, is directed or channeled into bypass airflow passage 56, and a second portion of air 58, as indicated by arrow 64, is directed or channeled into LP compressor 24. The ratio between first portion 62 of air and second portion 64 of air is generally referred to as the bypass ratio. The pressure of second portion 64 of air is then increased as it passes through high pressure (HP) compressor 26 and enters combustion section 28, where it is mixed with fuel and combusted to provide combustion gases 66. Combustion gases 66 are then directed through HP turbine 30 and LP turbine 32, where a portion of the thermal and / or kinetic energy from combustion gases 66 is extracted.

[0035] The combustion gases 66 are then directed through the combustion section 28 of the turbine 18 to provide propulsive thrust. Simultaneously, the pressure of the first portion of air 62 is significantly increased as the first portion of air 62 is directed through the bypass airflow passage 56 before being discharged from the fan nozzle exhaust section 68 of the turbofan engine 12, which also provides propulsive thrust.

[0036] However, it should be understood that Figure 1 The turbofan engine 12 depicted in the drawings is merely an example, and in other exemplary embodiments, aspects of the present disclosure may additionally or alternatively be applied to any other suitable gas turbine engine. For example, in other exemplary embodiments, the turbofan engine 12 may alternatively be any other suitable aviation gas turbine engine, such as a turbojet engine, a turboshaft engine, a turboprop engine, etc. Additionally, in other exemplary embodiments, the turbofan engine 12 may include or be operatively connected to any other suitable accessory system.

[0037] Now refer to Figure 2 , Figure 2 is a cross-sectional view of a portion of the turbomachine 18 and shows the HP compressor 26 , the combustion section 28 , and the HP turbine 30 .

[0038] like Figure 2 As shown, the HP compressor 26 includes a diffuser nozzle 70 and defines a rear cavity 72. The diffuser nozzle 70 is disposed at the rear end of the HP compressor 26 and straightens or redirects the air flow from the HP compressor 26 to the combustion section 28. The rear cavity 72 is also disposed at the rear end of the HP compressor and stores and supplies, for example, cooling air for use by the HP compressor 26. As will be discussed herein, the air within the rear cavity 72 can be used to provide thermal energy transfer benefits to the HP compressor 26.

[0039] The combustion section 28 includes a combustor casing 74 and a combustor 80. The combustor casing 74 of the combustion section 28 defines a fluid inlet 76 and at least partially defines a chamber configured to accommodate the combustor 80, referred to herein as a diffuser cavity 78. The fluid inlet 76 extends through a portion of the combustor casing 74. The gas turbine engine 10 also includes a cooling duct assembly that includes a cooling duct 82. The fluid inlet 76 is configured to deliver a cooling air flow 84 through the combustor casing 74 to the cooling duct 82.

[0040] In addition to the cooling duct 82, the cooling duct assembly also includes an outer duct 86. The cooling duct 82 and the outer duct 86 are tubes or conduits. In this example, the cooling duct 82 and the outer duct 86 have a generally circular cross-sectional shape. For the illustrated embodiment, the diameter of the cooling duct 82 and the diameter of the outer duct 86 are generally constant along the length of the cooling duct 82 and the outer duct 86.

[0041] A cooling duct 82 is connected to a portion of the combustor casing 74 and is in fluid communication with the fluid inlet 76 in the combustor casing 74. The cooling duct 82 extends from the fluid inlet 76, through a portion of the diffuser cavity 78, and between the HP compressor 26 and the combustor 80. For example, a portion of the cooling duct 82 is disposed inwardly from the combustor 80 in a radial direction R and extends in an axial direction A. The cooling duct 82 is configured to convey a cooling airflow 84 from the fluid inlet 76, through the diffuser cavity 78, and to a destination. In certain exemplary embodiments, the cooling duct 82 may convey the cooling airflow 84 to a portion or component associated with the HP turbine 30.

[0042] In this example, the cooling air flow 84 is a cooled cooling air flow from a cooled cooling air heat exchanger (see, e.g., Figure 5 However, it should be understood that in other exemplary embodiments, cooling air flow 84 may come from other cooling air sources, such as an ambient source, a bleed air source, a thermal management system of propulsion system 10, or other air sources.

[0043] The outer duct 86 surrounds at least a portion of the cooling duct 82 and extends along a portion of the entire length of the cooling duct 82. The term "entire length" with respect to the cooling duct 82 refers to the length of the cooling duct 82 from the location where the cooling duct 82 connects to the combustor casing 74 at the inlet 76 to a terminal end point of the cooling duct 82 opposite the end of the cooling duct at the inlet 76.

[0044] In this example, the first end 88 of the outer duct 86 is disposed adjacent to a portion of the combustor casing 74 surrounding the fluid inlet 76. As used herein, the term "adjacent" means next to or adjacent to, without any intervening objects or parts positioned therebetween. The second end 90 of the outer duct 86 is disposed at an end of the outer duct 86 opposite the first end 88. More specifically, in at least some exemplary aspects, the second end 90 of the outer duct 86 is disposed inwardly along the radial direction R of the turbofan engine and rearwardly of the first end 88 of the outer duct 86 (e.g., Figure 2 As shown, the forward direction points to the left and the rearward direction points to the right).

[0045] A gap 92 is formed between and defined by the cooling duct 82 and the outer duct 86. The outer duct 86 is configured to receive a buffer air flow 94 from the diffuser cavity 78. Figure 2 As shown, outer duct 86 receives a buffer air flow 94 into gap 92 at both first end 88 and second end 90 of outer duct 86. After buffer air flow 94 flows into gap 92, buffer air flow 94 is drawn through gap 92 toward supply line 98.

[0046] Within gap 92 are a plurality of struts 96. Struts 96 are small pieces of rigid or semi-rigid material that support or brace outer conduit 86 relative to cooling conduit 82 and extend between outer conduit 86 and cooling conduit 82. For example, struts 96 may comprise a minimum cross-sectional area in the direction of flow through gap 92.

[0047] like Figure 2 As shown, the outer tube 86 and the cooling tube 82 form a nested tube configuration, wherein the cooling tube 82 is nested within the outer tube 86. Specifically, the cooling tube 82 is nested within the outer tube 86. More specifically, in at least some exemplary aspects, the cooling tube 82 is disposed within the outer tube 86. In this particular example, the cooling tube 82 includes a portion that extends through or beyond the first end 88 and the second end 90 of the outer tube 86. In other examples, the outer tube 86 extends along a majority of the cooling tube 82 (e.g., greater than or equal to 50% of the entire length of the cooling tube 82, such as greater than or equal to 75% of the entire length of the cooling tube 82, such as greater than or equal to 85% of the entire length of the cooling tube 82, such as greater than or equal to 95% of the entire length of the cooling tube 82).

[0048] The cooling duct 82 and the outer duct 86 are generally coaxially disposed relative to one another. However, it should be understood that in certain exemplary embodiments, at certain points along the length of the cooling duct 82 and the outer duct 86, the axial centerlines of the cooling duct 82 and the outer duct 86 may be slightly offset due to curvature of the ducts or due to bending caused by gravity.

[0049] The nested configuration of the cooling duct 82 and the outer duct 86 forms a double-walled structure through which a buffer air flow 94 can flow. Here, the gap 92 forms a boundary layer that flows through the double-walled structure of the cooling duct 82 and the outer duct 86. The buffer air flow 94 (e.g., a boundary layer) sent through the gap 92 forms a separation layer between the air disposed in the diffuser cavity 78 and the cooling air flow 84 flowing through the cooling duct 82. In this way, the buffer air flow 94 absorbs heat energy from the air disposed in the diffuser cavity 78, thereby reducing the amount of heat energy transferred to the cooling air flow 84.

[0050] A supply line 98 extends between the outer duct 86 and the aft cavity 72 of the HP compressor 26 and fluidly connects the outer duct 86 and the aft cavity 72. In certain exemplary embodiments, a buffer air flow 94 is drawn into the supply line 98 from the gap 92 and delivered to the aft cavity 72. In this manner, a portion of the buffer air flow 94 is diverted to the aft cavity 72 of the HP compressor 26 of the propulsion system 10. After being delivered to the aft cavity 72, the buffer air flow 94 can be utilized by the HP compressor 26 for thermal energy transfer benefits. More specifically, in at least certain exemplary aspects, the portion of the buffer air flow 94 diverted to the aft cavity 72 can be used to reduce the thermal energy of the aft cavity 72 (and the air within the aft cavity 72) of the HP compressor 26 of the propulsion system 10.

[0051] In this example, the HP turbine 30 includes inlet guide vanes 100 ("IGV 100") and first stage blades 102. Figure 2 , a single IGV 100 and a single first-stage blade 102 are shown. However, it should be understood that the propulsion system 10 includes multiple IGVs 100 and multiple first-stage blades 102 extending in a circumferential direction about the longitudinal centerline 14. The IGV 100 is a stationary airfoil for directing or redirecting the flow of fluid passing through the IGV 100. Here, the IGV 100 straightens or changes the direction of the flow of combustion gases flowing from the combustion section 28 to the HP turbine 30. The first-stage blade 102 is an airfoil configured to rotate about the longitudinal centerline 14 (e.g., in conjunction with the rotor disk 108). For example, when combustion gases are exhausted from the combustor 80, the combustion gases push against the first-stage blade 102, causing the first-stage blade 102 to rotate about the longitudinal centerline 14.

[0052] A forward wheelspace cavity 104 is formed between the IGV 100 and the first-stage turbine blades 102. More specifically, in at least some exemplary aspects, the forward wheelspace cavity 104 is defined and formed by components corresponding to the IGV 100 and the first-stage turbine blades 102, such as a frame 106 that supports the IGV 100 and a rotor disk 108 to which the first-stage rotor blades are attached.

[0053] Furthermore, engine 10 includes components for providing cooling air flow 84 of cooling air from cooling duct 82 to rotor disk 108 and first-stage turbine blades 102. More specifically, frame 106 includes nozzle 103 configured to divert cooling air flow 84 from cooling duct 82 to at least partially match the rotation of rotor disk 108, and rotor disk 108 includes a rotary seal 105 that can rotate therewith. Air from nozzle 103 is then provided through opening 107 defined in the rotary seal to an inlet 109 of an internal cooling duct 111 (depicted in dashed lines) defined within rotor disk 108. Cooling air flow 84 can then be provided through rotor disk 108 to first-stage turbine blades 102.

[0054] In this example, at least a portion of the cooling air flow 84 flows through the seal formed between the frame 106 and the rotating seal and is subsequently delivered to the forward wheelspace cavity 104 .

[0055] The combustion section 28 also defines a forward outer axial cavity 110 . The forward outer axial cavity 110 is disposed generally inwardly from the cooling duct 82 in the radial direction R. The forward outer axial cavity 110 is fluidly connected to the HP turbine 30 via the forward wheelspace cavity 104 .

[0056] Here, the propulsion system 10 having the outer duct 86 surrounding the cooling duct 82 provides an insulated delivery loop for the cooling air flow 84 that reduces the amount of thermal contamination caused by the high levels of thermal energy present in the combustion section 28. Reducing thermal contamination of the cooling air flow 84 can reduce the flow rate of the cooling air flow 84 required for cooling, thereby enabling a reduction in the size of the source of the cooling air flow 84 (e.g., a cooling air heat exchanger, another heat exchanger, or another thermal management device).

[0057] Furthermore, in at least some exemplary aspects, a buffer air flow 94 can also be used by the HP compressor 26 via the aft cavity 72. The buffer air flow 94 can be used to reduce the temperature of the air within the aft cavity 72, thereby causing cooler air to be delivered to the HP compressor 26 for cooling. Because this cooler cooling air is delivered to the HP compressor 26 via the aft cavity 72, in some exemplary embodiments, a dedicated cooling air flow from a heat exchanger (which would otherwise be required) can be eliminated due to the cooling benefits gained by utilizing the buffer air flow 94 as cooling air for the HP compressor 26. In this manner, the propulsion system 10 having the cooling duct 82 and the outer duct 86 can improve the rotor temperature of the HP compressor 26 as well as the HP turbine 30.

[0058] Now refer to Figure 3 , Figure 3is a cross-sectional view of a portion of the turbomachine 18 and illustrates the HP compressor 26 , the combustion section 28 , the HP turbine 30 , and the cooling duct assembly according to another exemplary embodiment of the present disclosure.

[0059] Figure 3 The examples provided in Figure 2 The structure is similar to that described. Figure 3 Including the above Figure 2 The same or similar components as described herein include the supply line 98, IGV 100, first stage blades 102, forward wheel space cavity 104, and additional components as described below. Figure 2 The same or similar descriptions provided in the parts also apply to Figure 3 The corresponding parts shown in .

[0060] For the illustrated embodiment, the outer duct 86 is fluidly connected to a forward outer seal 112 and a rotary seal 114. In this example, the forward outer seal 112 is disposed downstream of the cooling duct 82 (e.g., as shown in FIG. Figure 3 A rotary seal 114 is disposed downstream of the forward outer seal 112 (shown from left to right).

[0061] In certain exemplary embodiments, the forward outer seal 112 may be positioned adjacent the rotor disk 108 and / or the first stage blades 102 (see, e.g., Figure 2 For example, the front outer seal 112 and the rotary seal 114 may be arranged along the longitudinal centerline 14 relative to the flow direction (at Figure 3 to the right) is positioned in front of the rotor disk 108 (in Figure 3 In this example, the cooling air flow 84 may flow through the nozzle 103 toward the opening 107 defined in the rotating seal 114 to the rotor disk 108, as described above with respect to FIG. Figure 2 In contrast, the buffer air flow 94 can flow through the front outer seal 112 and leak through the rotating seal 114 to the front seal cavity 104 (see Figure 2 ).

[0062] Furthermore, in this example, a portion of the cooling conduit 82 forms an inner venturi 116, while a portion of the outer conduit 86 forms an outer venturi 124. Figure 3 As shown, the outer venturi 124 surrounds the inner venturi 116 .

[0063] More specifically, in at least some exemplary aspects, the inner venturi 116 of the cooling duct 82 includes an inner narrowing portion 118 and an inner widening portion 120. The inner narrowing portion 118 comprises a portion of the cooling duct 82 in which an inner diameter 122 of the cooling duct 82 decreases along a downstream direction D of the cooling air flow 84 flowing through the cooling duct 82. The inner widening portion 120 comprises a portion of the cooling duct 82 in which an inner diameter 122 of the cooling duct 82 increases along a downstream direction D of the cooling air flow 84 flowing through the cooling duct 82.

[0064] Furthermore, outer duct 86 includes an outer narrowing portion 126 and an outer expanding portion 128. Outer narrowing portion 126 comprises a portion of outer duct 86 in which an inner diameter 130 of outer duct 86 decreases in a downstream direction D of cooling air flow 84 flowing through cooling duct 82. Outer expanding portion 128 comprises a portion of outer duct 86 in which an inner diameter 130 of outer duct 86 increases in a downstream direction D of cooling air flow 84 flowing through cooling duct 82. With respect to inner venturi 116, inner diameter 122 of cooling duct 82 is minimized at the interface between inner narrowing portion 118 and inner expanding portion 120. Similarly, inner diameter 130 of outer duct 86 is minimized at the interface between outer narrowing portion 126 and outer expanding portion 128.

[0065] During operation, the cooling air flow 84 is restricted at the location where the diameter of the cooling duct 82 is smallest (e.g., at the intersection between the inner narrowed portion 118 and the inner expanded portion 120). In other words, the cooling air flow 84 is restricted at the point along the cooling duct 82 where the cooling duct 82 forms the inner venturi 116.

[0066] Here, the configuration of the inner venturi 116 and outer venturi 124 provides a smaller outer surface area of the outer conduit 86 than a tube assembly without the venturi configuration. This smaller outer surface area of the outer conduit 86 can reduce the impact on the aerodynamic behavior of the air flowing within the combustion section 28.

[0067] Now refer to Figure 4 , Figure 4 is a cross-sectional view of a portion of the turbomachine 18 and shows the HP compressor 26 , the combustion section 28 , the HP turbine 30 , and the cooling duct assembly according to another exemplary embodiment of the present disclosure.

[0068] An exemplary cooling duct assembly includes a Figure 3 The cooling duct 82 is constructed in substantially the same manner as the cooling duct 82. However, as Figure 4 As depicted, the cooling duct 82 also includes swirl vanes 134 and defines slots 136. The swirl vanes 134 are airfoils configured to impart a swirl to the cooling air flow 84 passing through the swirl vanes 134. Figure 4As shown, the vortex vanes 134 are arranged along the radial direction R of the cooling duct 82. CD Provided on the inner surface of the cooling pipe 82 .

[0069] In one example, the swirl vanes 134 can be configured as a cyclone separator or a centrifugal separator. For example, with this configuration, the swirl vanes 134 are configured to impart rotational and centrifugal forces to the cooling air flow 84. As the cooling air flow 84 passes through the swirl vanes 134, the cooling air flow 84 is primarily divided into two streams, an inner cooling air flow 84A and an outer cooling air flow 84B. For example, as the cooling air flow 84 is swirled by the swirl vanes 134, the heavier portion of the cooling air flow 84 (i.e., the outer cooling air flow 84B) swirls outward in a radial direction of the cooling duct 82. The cooling air flow 84B can include air of a higher density, a higher amount of particulates, or both, than the cooling air flow 84A.

[0070] In this example, the swirl vanes 134 are disposed in the inner venturi 116. However, it should be understood that in other exemplary embodiments, the swirl vanes 134 may be disposed or positioned upstream or downstream along the downstream direction D of the cooling air flow 84 from the inner venturi 116, or may be positioned in a cooling duct assembly without the inner venturi 116 or the outer venturi 124.

[0071] In one example, the slot 136 can be a single, full-annular opening extending throughout the entire 360° of the cooling duct 82. In other examples, the slot 136 can include one or more openings, perforations, slits, slots, or holes disposed around the circumference of the cooling duct 82. For the illustrated embodiment, the slot 136 is disposed downstream in the downstream direction D of the cooling air flow 84 from the vortex vanes 134. In this manner, as the outer cooling air flow 84B flows outward in the radial direction of the cooling duct 82, a portion of the outer cooling air flow 84B exits the cooling duct 82 through the slot 136 and enters the gap 92 between the cooling duct 82 and the outer duct 86. The outer cooling air flow 84B continues through the gap 92 as buffer air 94.

[0072] Here, with swirl vanes 134 positioned within cooling duct 82, particle-laden (e.g., dirty) cooling air flow 84B is ejected from cooling duct 82 into outer duct 86, leaving cooling air flow 84A as a clean cooling air flow flowing from cooling duct 82 to downstream components, such as first-stage blades 102 and rotor disk 108. Thus, this configuration minimizes the risk of particles flowing into the internal cooling passages of rotor disk 108 and first-stage blades 102, potentially blocking or clogging such internal cooling passages.

[0073] In certain exemplary embodiments, the dirty or particulate laden cooling air flow 84B may be delivered to the forward outer seal 112 and through the rotating seal 114 for the forward cavity 104 .

[0074] Now refer to Figure 5 , Figure 5 is a cross-sectional view of a portion of the turbomachine 18 and illustrates the HP compressor 26, combustion section 28, HP turbine 30, and cooling duct assembly according to another exemplary embodiment of the present disclosure. Here, the turbomachine 18 and cooling duct assembly are shown in FIG. Figure 2 The turbine 18 and cooling duct assembly are constructed in essentially the same manner. However, for Figure 5 In the embodiment, the turbine 18 and cooling duct assembly further includes a separator 138 and a heat exchanger 140 .

[0075] Separator 138 is a component for separating a single air stream into two or more air streams. In certain exemplary embodiments, separator 138 can be a cyclone separator, a centrifugal separator, a particle separator, or another type of separator. In this example, separator 138 is disposed outside of combustor casing 74 and is fluidly connected to and between the compressor section and heat exchanger 140. During operation, separator 138 receives air stream 142 from the compressor section (e.g., HP compressor 26) and separates the air stream from the compressor section into a clean air stream 84A and a dirty air stream 84B. The separator then delivers clean air stream 84A to heat exchanger 140 and delivers dirty air stream 84B to outer duct 86 and into gap 92 as a buffer air stream 94. In the illustrated embodiment, the ends of cooling duct 82 and outer duct 86 extend through fluid inlet 76 defined by combustor casing 74.

[0076] In certain exemplary embodiments, heat exchanger 140 can comprise a cooling air heat exchanger for cooling. Here, heat exchanger 140 fluid is connected to separator 138 and cooling duct 82 and between them. During operation, heat exchanger 140 removes heat energy from clean air stream 84A and delivers cooled clean air stream 84A to cooling duct 82. Heat exchanger 140 can be in thermal communication with any suitable cooling fluid source. For example, heat exchanger 140 can be configured to receive hot fluid from a heat bus, bypass air from a bypass channel, fuel from a fuel source, etc.

[0077] like Figure 5As shown, the dirty air flow 84B serves as a buffer air utilized by the outer duct 86 to isolate the cooling duct 82 from thermal energy contamination within the diffuser cavity 78. After being drawn through the gap 92, the dirty air flow 84B may ultimately be delivered to the aft cavity 72 or the forward outer seal 112 of the HP compressor 26. For example, in certain exemplary embodiments, a portion of the dirty air flow 84B may be diverted from the gap 92 to the supply line 98 and delivered to the aft cavity 72 of the HP compressor 26.

[0078] Now refer to Figure 6 , Figure 6 is a cross-sectional view of a portion of a turbomachine 18 and shows an HP compressor 26, a combustion section 28, an HP turbine 30, and a cooling duct assembly according to another exemplary embodiment of the present disclosure. Here, the turbomachine 18 and the cooling duct assembly are shown in a manner similar to that shown in FIG. Figure 2 and 3 The turbine 18 and cooling duct assembly are constructed in essentially the same manner. However, for Figure 6 In the embodiment of FIG. 1 , the turbine 18 and cooling duct assembly further includes an additional third duct 144 having one or more openings 146 and a third air flow 148 .

[0079] In this example, the third duct 144 is disposed outside and around the outer duct 86, such that the cooling duct 82 and the outer duct 86 are nested within the third duct 144. Here, the third duct 144 is configured to receive an air source (e.g., a third air flow 148) from within the diffuser cavity 78 through one or more openings 146. For the depicted embodiment, the third duct 144 defines a first opening 146A proximate the fluid inlet 76 (e.g., closer to the fluid inlet 76 than the diffuser nozzle 70) and a second opening 146B located inboard of the diffuser nozzle 70 in the radial direction R of the gas turbine engine 10. The first and second openings 146A, 146B are each configured as an opening or slot disposed in or defined by the third duct 144 that enables the third air flow 148 to enter the third duct 144.

[0080] Here, the plurality of openings 146 of the third tube 144 are provided at different positions along the third tube 144 to allow the third air flow 148 to be drawn in at different positions within the diffuser cavity 78. These multiple intake positions allow the third tube 144 to receive third air flows 148 having different thermal energies to enhance the buffering of the cooling air flow 84, and the multiple buffer air layers to protect the cooling air flow 84 from thermal contamination within the diffuser cavity 78.

[0081] Now refer to Figure 7 , Figure 7is a cross-sectional view of a portion of a turbomachine 18 and shows an HP compressor 26, a combustion section 28, an HP turbine 30, and a cooling duct assembly according to another exemplary embodiment of the present disclosure. Here, the turbomachine 18 and the cooling duct assembly are shown in FIG. Figure 6 The turbine 18 and cooling duct assembly are constructed in essentially the same manner. However, for Figure 7 In the embodiment of FIG. 1 , the assembly includes a third tube 144 (having an opening 146 ) and a second gap 150 in a different configuration.

[0082] In this example, cooling duct 82 extends from combustor casing 74 at fluid inlet 76 to a portion of frame 106 located at second end 90 of outer duct 86 .

[0083] A third tube 144 is disposed around a portion of the outer tube 86 and a portion of the cooling tube 82. In this example, a first end of the third tube 144 is located at or near the supply line 98 (e.g., defining a spacing that is less than about 5 inches, such as less than about 3 inches, such as less than about 1 inch), and a second end of the third tube 144 is disposed at and connected to the frame 106 at a point where the cooling tube 82 connects to and passes through the frame 106. In this example, a portion of the cooling tube 82 and a portion of the outer tube 86 are nested within the third tube 144.

[0084] The gap 150 is fluidly connected to the gap 92 at the location of the second end 90 of the outer duct 86. In this example, the fluid channel is formed by the gap 150, the gap 92 and the supply line 98, which is configured to supply the buffer air flow 94 from two separate parts of the diffuser cavity 78 (for example, from the first end 88 of the outer duct 86 and from the opening 146 of the third tube 144) and reach the rear cavity 72 through the gap 92, the gap 150 and the supply line 98. After the buffer air flow 94 enters the third tube 144 through the opening 146, the buffer air flow 94 flows through the gap 150 in a first axial direction. When the buffer air flow 94 flowing through the gap 150 reaches the frame 106 at the second end 90 of the outer duct 86, the buffer air flow 94 turns around the second end 90 of the outer duct 86 and flows through the gap 92 in a second axial direction opposite to the first axial direction. In this example, the first axial direction is from front to back (such as Figure 7 From left to right as shown), the second axial direction is from back to front (as shown Figure 7 shown from right to left).

[0085] In certain exemplary embodiments, the opening 146 of the third tube 144 may be located at a location where the air pressure within the diffuser cavity 78 is higher than the air pressure at or near the first end 88 of the outer duct 86. The buffer air flow 94 entering the outer duct 86 at the first end 88 of the outer duct 86 and the buffer air flow 95 entering the third tube 144 at the opening 146 may cause the buffer air flow 94 to flow through the gap 150 and back through the gap 92 to the supply line 98 more quickly than the portion of the buffer air flow 94 flowing from the first end 88 of the outer duct 86 through the gap 92 to the supply line 98. With the increased velocity of the buffer air flow 94 flowing through the gap 150, thermal energy from the air within the diffuser cavity 78 has less time to be exposed to the buffer air flow 94 flowing through the third tube 144, thereby reducing the amount of thermal energy transferred from the diffuser cavity 78 to the cooling air flow 84 within the cooling duct 82.

[0086] Now refer to Figure 8 , Figure 8 is a cross-sectional view of a portion of the turbomachine 18 and shows the HP compressor 26, combustion section 28, HP turbine 30 and cooling duct assembly according to another exemplary embodiment of the present disclosure. Here, the turbomachine 18 and cooling duct assembly are shown in FIG. Figure 2 The turbine 18 and cooling duct assembly are constructed in essentially the same manner. However, for Figure 5 In the exemplary embodiment, the turbine 18 and cooling duct assembly further includes a bypass port 152 extending from the outer duct 86 to the front axle outer cavity 110 .

[0087] In this example, the combustion section 28 also includes an inner combustor casing 75, and the bypass port 152 is a fluid outlet in fluid communication with the gap 92 between the cooling duct 82 and the outer duct 86. The bypass port 152 extends from the outer duct 86, through a portion of the inner combustor outer casing 75, and into the forward axle outer cavity 110.

[0088] Without the supply line 98 connecting the gap 92 to the aft cavity 72 , all or a majority of the buffer air flow 94 exits the outer duct 86 through the bypass port 152 .

[0089] In this way, a larger amount of buffer air flow 94 or a higher pressure buffer air flow 94 can be supplied to the front axle outer cavity 110 and the rotating seal 114 (see, for example Figure 3 ) or components supplied to the HP turbine 30, such as the front wheel space cavity 104 or the rotor disk 108 (see e.g. Figure 2 In one example, the effect of a greater amount of buffer air flow 94 entering the front shaft outer cavity 110 is that a greater amount of buffer air flow 94 flows to components downstream of the front shaft outer cavity 110 , thereby enhancing the transfer of thermal energy and reducing thermal pollution of the turbine 18 .

[0090] It should be understood that Figures 1 to 8 Any configuration and / or components of the propulsion system 10 shown may be used with Figures 1 to 8 Likewise, the description of any component presented in a given figure is applicable and can be used to refer to or describe similarly numbered components shown in other figures.

[0091] This written description uses examples to disclose the present disclosure, including the best mode, and also to enable any person skilled in the art to practice the present disclosure, including making and using any devices or systems and performing any incorporated methods. The patentable scope of the present disclosure is defined by the claims, and may include other examples that occur to those skilled in the art. If these other examples include structural elements that 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, they are intended to fall within the scope of the claims.

[0092] Further aspects are provided by the subject matter of the following clauses:

[0093] A gas turbine engine includes a combustion section, a turbine section, and a compressor section having a high-pressure compressor. The combustion section includes a combustor housing, a burner, a cooling duct, and an outer duct. The combustor housing at least partially defines a diffuser cavity and a fluid inlet. The burner is disposed in the diffuser cavity. The cooling duct is in fluid communication with the fluid inlet in the combustor housing and is configured to deliver a cooling airflow. The outer duct surrounds at least a portion of the cooling duct and extends along a portion of the entire length of the cooling duct. The outer duct defines a gap with the cooling duct and is configured to deliver a buffer airflow through the gap. The turbine section is disposed downstream of the combustion section. The cooling duct is in fluid communication with the turbine section.

[0094] The gas turbine engine according to one or more of these clauses, wherein the gas turbine engine further comprises a cooled cooling air heat exchanger, wherein the cooling duct is fluidly connected to the cooled cooling air heat exchanger and is configured to receive a cooling air flow from the cooled cooling air heat exchanger.

[0095] The gas turbine engine according to one or more of these clauses, further comprising a particle separator disposed outside the combustor casing, wherein the outer duct is configured to receive the buffer air flow from the diffuser cavity, the cooling duct, the particle separator, or any combination thereof.

[0096] The gas turbine engine according to one or more of these clauses, wherein the cooling duct includes swirl vanes disposed on an inner surface of the cooling duct along a radial direction of the cooling duct, wherein the swirl vanes are configured to impart a swirl to a cooling air flow passing through the swirl vanes.

[0097] The gas turbine engine according to one or more of these clauses, wherein the buffer air flow received from the cooling duct through the outer duct comprises a first bleed air flow drawn from the cooling duct, the first bleed air flow being swirled outwardly in a radial direction of the cooling duct by the swirl vanes.

[0098] The gas turbine engine according to one or more of these clauses, wherein the turbine section includes a high pressure turbine having first stage high pressure turbine rotor blades, and wherein the cooling duct is in fluid communication with the first stage high pressure turbine rotor blades of the high pressure turbine.

[0099] A gas turbine engine according to one or more of these clauses, wherein the outer duct is configured to deliver a portion of the buffer air flow to a rear cavity of the high pressure compressor, a forward outer seal disposed downstream of the cooling duct, a first stage blade of the turbine section, a forward shaft outer cavity of the turbine section, a forward impeller space cavity, or any combination thereof.

[0100] The gas turbine engine according to one or more of these clauses, wherein the outer duct is a first outer duct, wherein the first outer duct defines a first gap between the cooling duct and the outer duct, wherein the combustion section further includes a second outer duct disposed around the first outer duct, wherein the first outer duct is nested inside the second outer duct, wherein the second outer duct defines a second gap between the first outer duct and the second outer duct.

[0101] The gas turbine engine according to one or more of these clauses, further comprising a plurality of struts connected to and extending between the cooling duct and the outer duct.

[0102] A gas turbine engine according to one or more of these clauses, wherein the cooling duct includes an inner narrowing portion and an inner widening portion, wherein the inner narrowing portion includes a portion of the cooling duct: in a portion, the inner diameter of the cooling duct decreases along the downstream direction of the cooling air flow flowing through the cooling duct; wherein the inner widening portion includes a portion of the cooling duct: in a portion, the inner diameter of the cooling duct increases along the downstream direction of the cooling air flow flowing through the cooling duct; wherein the outer duct includes an outer narrowing portion and an outer widening portion, wherein the outer narrowing portion includes a portion of the outer duct: in a portion, the inner diameter of the outer duct decreases along the downstream direction of the cooling air flow flowing through the cooling duct; wherein the outer widening portion includes a portion of the outer duct: in a portion, the inner diameter of the outer duct increases along the downstream direction of the cooling air flow flowing through the cooling duct.

[0103] The gas turbine engine according to one or more of these clauses, wherein a portion of the cooling duct forms an inner venturi, wherein a portion of the outer duct forms an outer venturi, wherein the outer venturi surrounds the inner venturi.

[0104] A method for managing thermal energy in a gas turbine engine includes providing a cooling air flow to a cooling duct located in a diffuser cavity defined by a combustor casing of the gas turbine engine. A buffer air flow may be provided into a gap defined between the cooling duct and an outer duct surrounding the cooling duct to insulate the cooling air flow provided to the cooling duct. The cooling duct is nested within the outer duct. The cooling air flow may be delivered to a turbine section of the gas turbine engine.

[0105] The method of one or more of these clauses, further comprising swirling a portion of the cooling air flow flowing through the cooling duct using swirl vanes, wherein providing the buffer air flow into the gap comprises providing a portion of the swirled cooling air flow into the gap between the cooling duct and the outer duct.

[0106] The method of one or more of these clauses, further comprising diverting a portion of the buffer air flow to an aft cavity of a high pressure compressor of the gas turbine engine.

[0107] The method of one or more of these clauses further comprises reducing thermal energy in a rear cavity of a high pressure compressor of the gas turbine engine using a portion of the buffer air flow.

[0108] The method according to one or more of these clauses further includes delivering the buffer air flow to a rear cavity of the high-pressure compressor, a front outer seal arranged downstream of the cooling duct, a first stage blade of the turbine section, a front axial outer cavity of the turbine section, a front impeller space cavity, a front axial outer cavity of the high-pressure turbine of the gas turbine engine, or any combination thereof.

[0109] The method of one or more of these clauses, further comprising restricting the cooling air flow through the cooling duct at a location where the diameter of the cooling duct is smallest.

[0110] The method of one or more of these clauses, further comprising restricting the flow of cooling air through the cooling duct at a point along the cooling duct where the cooling duct forms a venturi.

[0111] The method of one or more of these clauses, wherein delivering the cooling air flow to a turbine section of the gas turbine engine comprises delivering the cooling air flow to a first stage high pressure turbine rotor blade of a high pressure turbine of the turbine section.

[0112] A cooling duct assembly for a gas turbine engine includes a cooling duct and an outer duct. The gas turbine engine includes a turbine section and a combustor casing, the combustor casing at least partially defining a diffuser cavity and a fluid inlet. The cooling duct is configured to be in fluid communication with the fluid inlet in the combustor casing and with the turbine section when installed in the gas turbine engine. The cooling duct is configured to convey a cooling airflow. The outer duct surrounds at least a portion of the cooling duct and extends along a portion of the entire length of the cooling duct. The outer duct defines a gap with the cooling duct, the cooling duct being configured to convey a buffer airflow through the gap.

Claims

1. A gas turbine engine, characterized in that: include: a compressor section, the compressor section including a high-pressure compressor; A combustion section, the combustion section comprising: a combustor housing at least partially defining a diffuser cavity and a fluid inlet; a burner disposed in the diffuser cavity; a cooling duct in fluid communication with the fluid inlet in the combustor casing, wherein the cooling duct is configured to deliver a cooling air flow, wherein the cooling duct includes swirl vanes disposed on an inner surface of the cooling duct along a radial direction of the cooling duct, wherein the swirl vanes are configured to impart a swirl to the cooling air flow passing through the swirl vanes; and an outer duct surrounding at least a portion of the cooling duct and extending along a portion of the entire length of the cooling duct, wherein the outer duct defines a gap with the cooling duct, and wherein the outer duct is configured to convey a buffer air flow through the gap; and A turbine section is disposed downstream of the combustion section, wherein the cooling conduit is in fluid communication with the turbine section.

2. The gas turbine engine according to claim 1, wherein: in, The gas turbine engine further includes a cooled cooling air heat exchanger, wherein the cooling duct is fluidly connected to the cooled cooling air heat exchanger and is configured to receive the cooling air flow from the cooled cooling air heat exchanger.

3. The gas turbine engine according to claim 1, wherein: Further included is a particle separator disposed outside the combustor casing, wherein the outer duct is configured to receive the buffer air flow from the diffuser cavity, the cooling duct, the particle separator, or any combination thereof.

4. The gas turbine engine according to claim 1, wherein: The buffer air flow received from the cooling duct through the outer duct includes a first induced air flow drawn from the cooling duct, and the first induced air flow swirls outward along the radial direction of the cooling duct through the vortex vanes.

5. The gas turbine engine according to claim 1, wherein Wherein the turbine section includes a high-pressure turbine having first-stage high-pressure turbine rotor blades, and wherein the cooling conduit is in fluid communication with the first-stage high-pressure turbine rotor blades of the high-pressure turbine.

6. The gas turbine engine according to claim 1, wherein: in, The outer duct is configured to deliver a portion of the buffer air flow to a rear cavity of the high-pressure compressor, a front outer seal disposed downstream of the cooling duct, a first-stage blade of the turbine section, a front shaft outer cavity of the turbine section, a front impeller space cavity, or any combination thereof.

7. The gas turbine engine according to claim 1, wherein: The outer pipe is a first outer pipe, wherein the first outer pipe defines a first gap between the cooling pipe and the outer pipe, wherein the combustion section further includes a second outer pipe disposed around the first outer pipe, wherein the first outer pipe is nested inside the second outer pipe, and wherein the second outer pipe defines a second gap between the first outer pipe and the second outer pipe.

8. The gas turbine engine according to claim 1, wherein: Further included are a plurality of struts connected to and extending between the cooling duct and the outer duct.

9. The gas turbine engine according to claim 1, wherein: The cooling duct includes an inner narrowing portion and an inner expanding portion, wherein the inner narrowing portion includes a portion of the cooling duct in which the inner diameter of the cooling duct decreases along the downstream direction of the cooling air flow flowing through the cooling duct; wherein the inner expanding portion includes a portion of the cooling duct in which the inner diameter of the cooling duct increases along the downstream direction of the cooling air flow flowing through the cooling duct; wherein the outer duct includes an outer narrowing portion and an outer expanding portion, wherein the outer narrowing portion includes a portion of the outer duct in which the inner diameter of the outer duct decreases along the downstream direction of the cooling air flow flowing through the cooling duct; and wherein the outer expanding portion includes a portion of the outer duct in which the inner diameter of the outer duct increases along the downstream direction of the cooling air flow flowing through the cooling duct.

10. The gas turbine engine according to claim 1, wherein: A portion of the cooling pipe forms an inner venturi tube, a portion of the outer pipe forms an outer venturi tube, and the outer venturi tube surrounds the inner venturi tube.

11. A method of managing thermal energy in a gas turbine engine, characterized in that The method comprises: providing a cooling air flow to a cooling duct located in a diffuser cavity defined by a combustor casing of the gas turbine engine; providing a buffer air flow into a gap defined between the cooling duct and an outer duct surrounding the cooling duct to insulate the cooling air flow provided to the cooling duct, wherein the cooling duct is nested in the outer duct; and delivering the cooling air flow to a turbine section of the gas turbine engine; swirl a portion of the cooling air flow flowing through the cooling duct using swirl vanes, Wherein providing the buffer air flow into the gap includes providing a portion of the swirled cooling air flow into the gap between the cooling duct and the outer duct.

12. The method according to claim 11, characterized in that Further comprising diverting a portion of the buffer air flow to an aft cavity of a high pressure compressor of the gas turbine engine.

13. The method according to claim 12, characterized in that Further comprising utilizing a portion of the buffer air flow to reduce thermal energy in the aft cavity of the high pressure compressor of the gas turbine engine.

14. The method according to claim 11, characterized in that It further includes delivering the buffer air flow to the rear cavity of the high-pressure compressor, the front outer seal arranged downstream of the cooling duct, the first stage blades of the turbine section, the front shaft outer cavity of the turbine section, the front impeller space cavity, the front shaft outer cavity of the high-pressure turbine of the gas turbine engine, or any combination thereof.

15. The method according to claim 11, characterized in that Further included restricting the cooling air flow through the cooling duct at a location where the diameter of the cooling duct is smallest.

16. The method according to claim 11, characterized in that Further included restricting the cooling air flow through the cooling duct at a point along the cooling duct where a venturi is formed.

17. The method according to claim 11, characterized in that Wherein delivering the cooling air flow to a turbine section of the gas turbine engine includes delivering the cooling air flow to first stage high pressure turbine rotor blades of a high pressure turbine of the turbine section.

18. A cooling duct assembly for a gas turbine engine comprising a turbine section and a combustor casing, the combustor casing at least partially defining a diffuser cavity and a fluid inlet, characterized in that The cooling duct assembly comprises: a cooling duct configured to be in fluid communication with the fluid inlet in the combustor casing and with the turbine section when installed in the gas turbine engine, wherein the cooling duct is configured to convey a cooling air flow, wherein the cooling duct includes swirl vanes disposed on an inner surface of the cooling duct in a radial direction of the cooling duct, wherein the swirl vanes are configured to impart a swirl to the cooling air flow passing through the swirl vanes; and An outer duct surrounds at least a portion of the cooling duct and extends along a portion of the entire length of the cooling duct, wherein the outer duct defines a gap with the cooling duct, and wherein the outer duct is configured to convey a buffer air flow through the gap.

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