Sealing system for turbomachine

Abstract

A turbomachine includes a nozzle section including an inner shroud defining a bottom surface and a nozzle flange defining a forward side surface and an aft side surface. The floating rotor seal is coupled to the nozzle flange via the carrier flange. The carrier flange includes a front wall and a rear wall. A nozzle flange is positioned between the forward wall and the aft wall defining a flow path therebetween. A sealing pocket is defined in one of the front wall or the rear wall and is in fluid communication with the flow path. At least one linear sealing section is partially disposed within the sealing pocket. The linear sealing section is configured to form a seal against the nozzle flange or the bottom surface in response to pressurization of the sealing pocket via the working fluid in the flow path.

Description

Technical Field

[0001] This disclosure relates to a gas turbine engine having a turbine. More specifically, this disclosure relates to a sealing system for the turbine of a gas turbine engine. Background Technology

[0002] Gas turbine engines typically consist of a compressor section, a combustion section, and a turbine section in a sequential flow order. At least one rotor shaft extends axially through the compressor section, combustion section, and turbine section. During operation, compressed air for cooling and combustion, as well as combustion gases, are directed through various flow paths defined within the gas turbine engine. Leakage or backflow of compressed air or combustion gases into certain areas of the gas turbine engine can negatively impact engine component life and engine efficiency. Attached Figure Description

[0003] The complete and feasible disclosure of this disclosure, including its best mode, is set forth in the specification with reference to the accompanying drawings, for those skilled in the art, wherein:

[0004] Figure 1 This is a perspective view of an exemplary aircraft according to exemplary embodiments of the present disclosure.

[0005] Figure 2 This is a cross-sectional schematic diagram of a gas turbine engine including a turbine according to an exemplary aspect of this disclosure.

[0006] Figure 3 This is an enlarged cross-sectional schematic diagram of a portion of a high-pressure turbine according to an exemplary embodiment of the present disclosure, the high-pressure turbine including high-pressure turbine stator blades, high-pressure turbine rotor blades, and an exemplary nozzle section.

[0007] Figure 4 This is based on exemplary embodiments of the present disclosure. Figure 3 The diagram shows an enlarged cross-sectional view of a portion of the nozzle section, including the guide vane and a portion of the inner shroud.

[0008] Figure 5 This is a rear-to-front view of a plurality of nozzle segments arranged in a ring around a longitudinal centerline according to an exemplary embodiment of the present disclosure.

[0009] Figure 6 This is a schematic cross-sectional view of a portion of a first linear sealing section and a portion of a second linear sealing section according to an embodiment of the present disclosure.

[0010] Figure 7 This is an enlarged cross-sectional schematic diagram of a portion of an exemplary nozzle segment according to an exemplary embodiment of the present disclosure, including a guide vane and a portion of an inner shroud.

[0011] Figure 8 This is an enlarged cross-sectional schematic diagram of a portion of an exemplary nozzle segment according to an exemplary embodiment of the present disclosure.

[0012] Figure 9 This is based on exemplary embodiments of the present disclosure. Figure 3 The circle (A) in the diagram shows an enlarged cross-sectional view of a portion of the high-pressure turbine.

[0013] Figure 10 This is an enlarged cross-sectional schematic diagram of a portion of a high-pressure turbine according to an exemplary embodiment of the present disclosure, including a portion of an exemplary nozzle segment. Detailed Implementation

[0014] Reference will now be made in detail to embodiments of the present disclosure, one or more examples of which are illustrated in the accompanying drawings. Detailed descriptions use numbers and letters to denote features in the drawings. Similar or related names in the drawings and specification are used to refer to similar or related portions of the present disclosure.

[0015] As used herein, the term “exemplary” means “as an example, instance, or illustration.” Any implementation described herein as “exemplary” is not necessarily to be construed as being better or more advantageous than other implementations.

[0016] As used herein, the terms “first,” “second,” and “third” are used interchangeably to distinguish one component from another and do not imply the position or importance of the components.

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

[0018] Unless otherwise specified herein, the terms “connection,” “fixed,” “attached to,” etc., refer to both direct connection, fixation, or attachment, and indirect connection, fixation, or attachment through one or more intermediate couplings or features.

[0019] Unless the context clearly indicates otherwise, the singular forms “a,” “one,” and “the” include plural references.

[0020] The approximate language used throughout this specification and claims is intended to modify any quantitative expression that may allow for variation without altering its underlying function. Therefore, values ​​modified by one or more terms (e.g., “approximately,” “approximately,” and “basically”) are not limited to specified exact values. In at least some cases, approximate language may correspond to the precision of the instrument used to measure the value, or the precision of the method or machine used to construct or manufacture the part and / or system. For example, approximate language might refer to a margin of 10%.

[0021] The terms "front" and "rear" refer to relative positions within a gas turbine engine or carrier, and specifically to the normal operating posture of the gas turbine engine or carrier. For example, in the case of a gas turbine engine, "front" refers to the position closer to the engine inlet, while "rear" refers to the position closer to the engine nozzle or exhaust port.

[0022] The various embodiments shown and described herein provide a sealing system for a gas turbine engine. The sealing system includes multiple linear sealing segments arranged end-to-end around the longitudinal centerline of the gas turbine engine and at least partially disposed within corresponding sealing pockets. The linear sealing segments are disposed between axially adjacent stationary components, such as, but not limited to, the nozzle flange and carrier flange of a nozzle segment, or the turbine blade shroud and the outer shroud of the turbine nozzle segment. The system may include one or more linear sealing segments per stationary component, which may have different lengths, heights, and stiffnesses, and may be angled to make room for higher sealing pockets, or to allow for inward movement of the piston rod at the end of each adjacent stationary component.

[0023] Linear sealing sections use linear contact to seal between stationary components. The linear sealing section moves dynamically with each corresponding individual nozzle section or stationary component to load the linear sealing section onto the corresponding contact surface using pressure increments during gas turbine engine operation, thereby maintaining a continuous seal. Spring-biasing members, such as wave springs, can be used to preload or bias the linear sealing section toward the corresponding sealing surface. Linear sealing sections can be used to seal the front or rear of a nozzle flange and can be configured to provide axial or radial sealing.

[0024] Now refer to the attached diagram, Figure 1 This is a perspective view of an aircraft 10 that can be incorporated into at least one exemplary embodiment of this disclosure. (See figure) Figure 1 As shown, the aircraft 10 has a fuselage 12, wings 14 attached to the fuselage 12, and a tail 16. The aircraft 10 also includes a propulsion system 18, which generates propulsive thrust to propel the aircraft 10 during flight, taxiing operations, etc. Although the propulsion system 18 is shown attached to the wing 14, in other embodiments it may additionally or alternatively include one or more aspects coupled to other parts of the aircraft 10, such as the tail 16, the fuselage 12, or both.

[0025] The propulsion system 18 includes at least one engine. In the exemplary embodiment shown, the aircraft 10 includes a pair of gas turbine engines 20. Each gas turbine engine 20 is mounted to the aircraft 10 in an underwing configuration. Each gas turbine engine 20 is capable of selectively generating propulsive thrust for the aircraft 10. The gas turbine engines 20 can be configured to burn various forms of fuel, including but not limited to jet fuel / aviation turbine fuel and hydrogen fuel, unless otherwise specified.

[0026] Figure 2 This is a cross-sectional side view of a gas turbine engine 20 according to an exemplary embodiment of the present disclosure. More specifically, for Figure 2 In one embodiment, the gas turbine engine 20 is a multi-axis, high-bypass turbofan jet engine, sometimes also referred to as a "turbofan engine." For example... Figure 2 As shown, the gas turbine engine 20 defines an axial direction A (extending parallel to a longitudinal centerline 22 provided for reference), a radial direction R, and a circumferential direction C extending about the longitudinal centerline 22. Typically, the gas turbine engine 20 includes a fan section 24 and a turbine 26 disposed downstream of the fan section 24.

[0027] The turbine 26 depicted typically includes an outer casing 28 defining an annular core inlet 30. The outer casing 28 at least partially surrounds an axial compressor section, a combustion section 36, a turbine section, and an exhaust nozzle 42 in a series flow relationship. The axial compressor section includes a turbocharger or low-pressure compressor 32 and a high-pressure compressor 34, and the turbine section includes a high-pressure turbine 38 and a low-pressure turbine 40.

[0028] The high-pressure shaft 44 drives the high-pressure turbine 38 to the high-pressure compressor 34. The low-pressure shaft 46 drives the low-pressure turbine 40 to the low-pressure compressor 32. The low-pressure compressor 32, the high-pressure compressor 34, the combustion section 36, the high-pressure turbine 38, the low-pressure turbine 40, and the exhaust nozzle 42 together define a working gas flow path 48 through the gas turbine engine 20.

[0029] In the depicted embodiment, fan section 24 includes a fan 50 having a plurality of fan blades 52 spaced apart and coupled to disk 54. As shown, the fan blades 52 generally extend outward from disk 54 along a radial direction R. Each fan blade 52 is operably coupled to a pitch changing mechanism 56 by means of the fan blades 52, and the pitch changing mechanism is configured to collectively, for example, uniformly change the pitch of the fan blades 52.

[0030] The gas turbine engine 20 also includes a power gearbox 58. The fan blades 52, disk 54, and pitch-changing mechanism 56 can rotate together about a longitudinal centerline 22 via a low-pressure shaft 46 passing through the power gearbox 58. The power gearbox 58 includes multiple gears for adjusting the rotational speed of the fan 50 relative to the low-pressure shaft 46, allowing the fan 50 and the low-pressure shaft 46 to rotate at a more efficient relative speed.

[0031] Still referencing Figure 2 In an exemplary embodiment, the disk 54 is covered by a rotatable front hub 60 (sometimes referred to as a "rotor") of the fan section 24. The front hub 60 has an aerodynamic profile to facilitate airflow through a plurality of fan blades 52. Furthermore, the fan section 24 includes an annular fan housing or outer nacelle 62 circumferentially surrounding at least a portion of the fan 50 and / or turbine 26. In the depicted embodiment, the outer nacelle 62 is supported relative to the turbine 26 by a plurality of circumferentially spaced struts or outlet guide vanes 64. Additionally, a downstream section 66 of the outer nacelle 62 extends over the outer portion of the turbine 26 to define a bypass airflow passage 68 therebetween.

[0032] However, it should be understood that Figure 2 The gas turbine engine 20 depicted herein is provided as an example only, and in other exemplary embodiments, the gas turbine engine 20 may have other configurations. For example, although the depicted gas turbine engine 20 is configured as a ducted gas turbine engine (e.g., including an outer nacelle 62), in other embodiments, the gas turbine engine 21 may be a non-ducted or ductless gas turbine engine (such that the fan 50 is a non-ducted fan, and the outlet guide vane 64 cantilevered from the housing 28).

[0033] Alternatively, while the depicted gas turbine engine 20 is configured as a geared gas turbine engine (e.g., including a power gearbox 58) and a variable-pitch gas turbine engine (e.g., including a fan 50 configured as a variable-pitch fan), in other embodiments, the gas turbine engine 20 may be configured as a direct-drive gas turbine engine (such that the low-pressure shaft 46 rotates at the same speed as the fan 50), a fixed-pitch gas turbine engine (such that the fan 50 includes fan blades 52 that cannot rotate about the pitch axis P), or both. It should also be understood that, in other exemplary embodiments, aspects of this disclosure may be incorporated into any other suitable gas turbine engine. For example, in other exemplary embodiments, aspects of this disclosure may (as appropriate) be incorporated into, for example, a turboprop gas turbine engine, a turboshaft gas turbine engine, or a turbojet gas turbine engine.

[0034] During operation of the gas turbine engine 20, a volume of air 70 enters the gas turbine engine through the relevant inlet 72 of the outer nacelle 62 and the fan section 24. As the volume of air 70 passes through the fan blades 52, a first portion of the air 74 is directed or directed into the bypass airflow passage 68, and a second portion of the air 76 is directed or directed into the working gas flow path 48, or more specifically, into the low-pressure compressor 32. The ratio between the first portion of air 74 and the second portion of air 76 is commonly referred to as the bypass ratio.

[0035] As the second portion of air 76 enters the low-pressure compressor 32, one or more sequential stages of the low-pressure compressor stator blades 78 and rotor blades 80, coupled to the low-pressure shaft 46, progressively compress the second portion of air 76 flowing through the low-pressure compressor 32 on its way to the high-pressure compressor 34. Next, one or more sequential stages of the high-pressure compressor stator blades 82 and rotor blades 84, coupled to the high-pressure shaft 44, further compress the second portion of air 76 flowing through the high-pressure compressor 34. This supplies compressed air to the combustion section 36, where it mixes with fuel and burns to provide combustion gases 86.

[0036] Combustion gas 86 is guided through high-pressure turbine 38, where a portion of the thermal and / or kinetic energy from the combustion gas 86 is extracted via a sequential stage of high-pressure turbine stator blades 88 connected to the turbine housing and high-pressure turbine rotor blades 90 connected to the high-pressure shaft 44, thereby rotating the high-pressure shaft 44 and supporting the operation of high-pressure compressor 34. Combustion gas 86 is then guided through low-pressure turbine 40, where a second portion of the thermal and kinetic energy is extracted from the combustion gas 86 via a sequential stage of low-pressure turbine stator blades 92 connected to the turbine housing and low-pressure turbine rotor blades 94 connected to the low-pressure shaft 46, thereby rotating the low-pressure shaft 46 and supporting the operation of low-pressure compressor 32 and / or fan 50.

[0037] Combustion gas 86 is then directed through the exhaust nozzle 42 of turbine 26 to provide propulsive thrust. Simultaneously, the pressure of the first portion of air 74 increases significantly as it is directed through bypass airflow passage 68 before exiting the exhaust section 96 of the fan nozzle of gas turbine engine 20, also providing propulsive thrust. High-pressure turbine 38, low-pressure turbine 40, and exhaust nozzle 42 at least partially define a hot gas path 98 for directing combustion gas 86 through turbine 26. Each stage of the high-pressure turbine stator blades 88 includes multiple nozzle sections, which will be described in more detail below, arranged circumferentially around the longitudinal centerline 22 of gas turbine engine 20.

[0038] Figure 3This is an enlarged schematic diagram of a portion of a high-pressure turbine 38 of a turbine 26 according to an exemplary embodiment of the present disclosure, the high-pressure turbine including high-pressure turbine stator blades 88, high-pressure turbine rotor blades 90, and a nozzle section 100. It should be understood that, although... Figure 3 The image shows one nozzle segment, but the high-pressure turbine 38 typically includes multiple nozzle segments 100 arranged annularly around the longitudinal centerline 22. Figure 3 As shown, the nozzle section 100 includes a guide vane 102. Although Figure 3 Only one guide vane is shown, but it should be understood that the nozzle section 100 may include multiple guide vanes spaced circumferentially around the longitudinal centerline 22. For example... Figure 3 As shown, the guide vane 102 extends between the inner shroud 104 and the outer shroud 106 in a radial direction R generally perpendicular to the longitudinal centerline 22 and an axial direction A parallel to the longitudinal centerline 22. The guide vane 102, the inner shroud 104, and the outer shroud 106 at least partially define the hot gas path 98 through the high-pressure turbine 38.

[0039] Figure 4 This is based on exemplary embodiments of the present disclosure. Figure 3 An enlarged view of a portion of the nozzle section 100 shown, including the guide vane 102 and a portion of the inner shroud 104. Figure 4 As shown, nozzle section 100 includes nozzle flange 108. Nozzle flange 108 extends radially inward from bottom surface 110 of inner shroud 104 relative to the radial direction R. Floating rotor seal 112 is coupled to nozzle flange 108 via mechanical fasteners 114 (e.g., but not limited to pins or bolts). Floating rotor seal 112 includes carrier flange 116 configured or formed to be mounted to nozzle flange 108. Floating rotor seal 112 may also include or define sealing block 118 configured or formed to abut against rotor shaft 122 (e.g., Figure 2 The outer surface 120 of the high-pressure shaft 44 or low-pressure shaft 46 shown is sealed.

[0040] like Figure 4As shown, nozzle flange 108 defines a front surface 124 and a rear surface 126. Carrier flange 116 is formed or shaped to receive at least a portion of nozzle flange 108. In the illustrated exemplary embodiment, carrier flange 116 includes a front wall 128 defining a rearward-facing surface 130 and a rear wall 132 defining a forward-facing surface 134. Mechanical fastener 114 may extend through the front wall 128, nozzle flange 108, and rear wall 132 to engage floating rotor seal 112 to nozzle flange 108. The rearward-facing surface 130 of front wall 128, the front surface 124 of nozzle flange 108, the forward-facing surface 134 of rear wall 132, and the rear surface 126 of nozzle flange 108 at least partially define a flow path 136 therebetween.

[0041] exist Figure 4 In the exemplary embodiment shown, the rear wall 132 defines a sealing recess 138 defined along the forward-facing surface 134. A linear sealing segment 140 is partially disposed within the sealing recess 138. In the exemplary embodiment, as Figure 4 As shown, a biasing member 142, such as, but not limited to, a wave spring, may be at least partially disposed in the sealing recess 138. The biasing member 142 may contact the linear sealing section 140 in a manner that provides an axial force or "axial action force" relative to the axial direction A, a radial force or "radial action force" relative to the radial direction R, or both axial and radial forces to the linear sealing section 140. The biasing member 142 may bias the linear sealing section 140 toward the forward-facing surface 134 of the rear wall 132.

[0042] In an exemplary embodiment, such as Figure 4 As shown by the dashed lines, the sealing recess 138 can be radially offset from the top surface 144 of the front wall 128 of the carrier flange 116 relative to the radial direction R. In other words, the rear wall 132 can be higher than the front wall 128 in the radial direction R. This configuration allows for improved fabrication of the carrier flange 116 and formation of the sealing recess 138.

[0043] Figure 5 Provided according to Figure 4 The exemplary embodiment shown is a rear-to-front view of a plurality of nozzle segments 100 arranged in a ring around a longitudinal centerline 22. For clarity, the guide vanes, outer shroud, and floating rotor seal have been removed. Figure 5 As shown, each nozzle segment 100 may include at least one linear sealing segment 140, which is at least partially disposed in a corresponding sealing recess 138. For example, in an exemplary embodiment, each nozzle segment 100 includes two linear sealing segments 140a, 140b. (Common Reference) Figure 4 and Figure 5The linear sealing section 140 can be formed to extend linearly between the forward-facing surface 134 of the rear wall 132 and the rear side surface 126 of the nozzle flange 108.

[0044] Figure 6 Schematic diagrams of a portion of linear sealing section 140a and a portion of linear sealing section 140b according to embodiments of the present disclosure are provided. Figure 6 As shown, the end 146 of the linear sealing section 140a may be formed or shaped to at least partially overlap with the adjacent end 148 of the linear sealing section 140b.

[0045] For reference again Figure 3 and Figure 4 The first pressure chamber 150 is at least partially defined between the outer surface 120 of the rotor shaft 122, the bottom surface 110 of the inner shroud 104, the front surface 124 of the nozzle flange 108 of the nozzle section 100, and the floating rotor seal 112. The second pressure chamber 152 is at least partially defined between the outer surface 120 of the rotor shaft 122, the bottom surface 110 of the inner shroud 104, the rear surface 126 of the nozzle flange 108, and the floating rotor seal 112.

[0046] In operation, such as Figure 3 and Figure 4 As shown, the working fluid (WF) enters the first pressure chamber 150 at a first pressure (P1). The working fluid WF may at least partially comprise combustion gas 86 or compressed air. The second pressure chamber 152 is at a second pressure (P2) lower than the first pressure P1. A portion of the working fluid WF flows into the flow path 136, around the end portion 154 of the nozzle flange 108, and into the sealing recess 138.

[0047] like Figure 4 As shown, the working fluid WF pressurizes the sealing recess 138 and applies a radially outward force (Fro) relative to the radial direction R to the linear sealing section 140, thereby causing the linear sealing section 140 to abut against the upper surface 156 of the sealing recess 138. Alternatively, the working fluid WF applies a forward axial force (Faf) relative to the axial direction A to the linear sealing section 140, thereby pressing the linear sealing section 140 against the rear surface 126 of the nozzle flange 108 and forming a seal therebetween. The linear sealing section 140 prevents or blocks the working fluid WF from leaking from the first pressure chamber 150 to the second pressure chamber 152.

[0048] Figure 7 This is an enlarged view of a portion of a nozzle segment 100 according to another exemplary embodiment of the present disclosure, the nozzle segment 100 including a guide vane 102 and a portion of an inner shroud 104. Figure 7As shown, the front wall 128 defines a sealing recess 238 defined along the rearward-facing surface 130. A linear sealing segment 240 is partially disposed within the sealing recess 238. In an exemplary embodiment, as Figure 7 As shown, a biasing member 242, such as, but not limited to, a wave spring, may be at least partially disposed in the sealing recess 238. The biasing member 242 may contact the linear sealing section 240 in a manner that provides an axial force relative to the axial direction A, a radial force relative to the radial direction R, or both axial and radial forces to the linear sealing section 240. In some embodiments, the front wall 128 may define at least one channel 258 that defines a flow path for fluid communication between the first pressure chamber 150 and the sealing recess 238.

[0049] In an exemplary embodiment, such as Figure 7 As shown by the dashed lines, the sealing recess 238 can be radially offset from the top surface 244 of the rear wall 132 of the carrier flange 116 relative to the radial direction R. In other words, the front wall 128 can be higher than the rear wall 132 in the radial direction R. This configuration allows for improved manufacturing of the carrier flange 116 and formation of the sealing recess 238.

[0050] In operation, such as Figure 7 As shown, the working fluid WF enters the first pressure chamber 150 at a first pressure P1. The working fluid WF may at least partially comprise combustion gas 86 or compressed air. The second pressure chamber 152 is at a second pressure P2, lower than the first pressure P1. A portion of the working fluid WF flows into the flow path 136 and into the sealing recess 238. The working fluid WF pressurizes the sealing recess 238 and applies a radially inward force (Fri) relative to the radial direction R to the linear sealing section 240, thereby causing the linear sealing section 240 to abut against the lower surface 256 of the sealing recess 238. Alternatively, the working fluid WF applies a rearward axial force (Faa) relative to the axial direction A to the linear sealing section 240, thereby pressing the linear sealing section 240 against the front surface 124 of the nozzle flange 108 and forming a seal therebetween. The linear sealing section 240 prevents or stops the working fluid WF from leaking from the first pressure chamber 150 or flowing into the second pressure chamber 152.

[0051] Figure 8 This is an enlarged view of a portion of a nozzle segment 100 according to another exemplary embodiment of the present disclosure, including a guide vane 102 and a portion of an inner shroud 104. Figure 8As shown, the nozzle flange 108 may include a secondary flange 162. The secondary flange 162 defines a front end 164 and a rear end 166. A rear wall 132 defines a sealing recess 338 along its rear end 168. A linear sealing section 340 is partially disposed within the sealing recess 338. In some embodiments, the rear wall 132 may define at least one channel 358 that defines a flow path for fluid communication between the first pressure chamber 150 and the sealing recess 338.

[0052] In operation, such as Figure 8 As shown, the working fluid WF enters the first pressure chamber 150 at a first pressure P1. The working fluid WF may at least partially comprise combustion gas 86 or compressed air. The second pressure chamber 152 is at a second pressure P2, lower than the first pressure P1. A portion of the working fluid WF flows into the flow path 136, around the end portion 154 of the nozzle flange 108, and into the sealing recess 338. The working fluid WF pressurizes the sealing recess 338 and applies a radially inward force (Fri) relative to the radial direction R to the linear sealing section 340, thereby causing the linear sealing section 340 to abut against the lower surface 374 of the sealing recess 338. Alternatively, the working fluid WF applies a rearward axial force (Faa) relative to the axial direction A to the linear sealing section 340, thereby pressing the linear sealing section 340 against the front surface 164 of the secondary flange 162 of the nozzle flange 108 and forming a seal therebetween. The linear sealing section 340 prevents or stops the working fluid WF from leaking or flowing from the first pressure chamber 150 to the second pressure chamber 152.

[0053] In an exemplary embodiment, such as Figure 3 As shown, a sealing recess 438 may be formed in the first fixing member 170, for example in the turbine blade shroud 172. The turbine blade shroud 172 circumferentially surrounds the corresponding row of high-pressure turbine rotor blades 90. Figure 9 This is based on exemplary embodiments of the present disclosure. Figure 3 The circle (A) in the image shows an enlarged view of a portion of the high-pressure turbine 38. (See image for reference.) Figure 9 As shown, the sealing recess 438 is positioned along the surface 474 of the first fixing member 170 and oriented toward the sealing surface 176 of the second fixing member 178, which is axially adjacent to the first fixing member. The second fixing member may include an outer cover 180 of the axially adjacent high-pressure turbine stator blade 184.

[0054] The linear sealing segment 440 is at least partially disposed within the sealing recess 438 and extends toward the sealing surface 176 of the second retaining member 178. In a particular embodiment, a biasing member 442, such as, but not limited to, a wave spring, may be at least partially disposed within the sealing recess 438. The biasing member 442 may contact the linear sealing segment 440 in a manner that provides an axial force relative to the axial direction A, a radial force relative to the radial direction R, or both axial and radial forces to the linear sealing segment 440. The biasing member 442 may bias the linear sealing segment 440 toward the sealing surface 176 of the second retaining member 178.

[0055] In operation, such as Figure 9 As shown, the working fluid WF, for example, comes from the high-pressure compressor 34 ( Figure 2 Compressed air flows into and pressurizes the sealing recess 438, applying a radially inward force (Fri) relative to the radial direction R to the linear sealing section 440, thereby sealing the linear sealing section 440 against the surface 182 of the sealing recess 438. Alternatively, the working fluid WF applies an axial force (Faf) relative to the axial direction A to the linear sealing section 440, pressing the linear sealing section 440 against the sealing surface 176 of the second fixing member 178 and forming a seal therebetween. In this configuration, the linear sealing section 440 prevents combustion gas 86 from the hot gas path 98 from leaking between the first fixing member 170 and the second fixing member 178 and from entering areas outside the hot gas path 98.

[0056] Figure 10 A schematic diagram of a portion of a high-pressure turbine according to an exemplary embodiment of the present disclosure is provided, including a portion of a nozzle section 500. Figure 10 As shown, the nozzle segment includes a nozzle flange 508 that extends radially inward from the bottom surface 510 of the inner shroud 504 of the nozzle segment 500 relative to the radial direction R. A floating rotor seal 512 is coupled to the nozzle flange 508 via a carrier flange 516. In a particular embodiment, a mechanical fastener 514, such as, but not limited to, a pin or bolt, can engage the carrier flange 516 to the nozzle flange 508. The carrier flange 516 includes a front wall 528 defining a top surface 544.

[0057] The front wall 528 defines a sealing recess 538 disposed along the top surface 544 and oriented toward the bottom surface 510 of the inner shroud 504. A linear sealing segment 540 is at least partially disposed within the sealing recess 538. In an exemplary embodiment, a biasing member 542, such as, but not limited to, a wave spring, may be at least partially disposed within the sealing recess 538. The biasing member 542 may contact the linear sealing segment 540 in a manner that provides a radial force relative to the radial direction R. The biasing member 542 may bias the linear sealing segment 540 toward the bottom surface 510 of the inner shroud 506 of the nozzle segment 500.

[0058] In operation, such as Figure 10 As shown, the working fluid WF enters the first pressure chamber 150 at a first pressure P1. The working fluid WF may at least partially comprise combustion gas 86 or compressed air. The second pressure chamber 152 is at a second pressure P2, lower than the first pressure P1. A portion of the working fluid WF flows into and pressurizes the sealing recess 538. The working fluid WF exerts a radially outward force (Fro) relative to the radial direction R on the linear sealing section 540, causing the linear sealing section 540 to abut against the bottom surface 510 of the inner shroud 506 of the nozzle section 500. The linear sealing section 540 prevents or blocks the working fluid WF from leaking from the first pressure chamber 150 or flowing into the second pressure chamber 152.

[0059] Further aspects are provided by the following topics:

[0060] A turbine includes: a nozzle section including an inner shroud and a nozzle flange, the inner shroud defining a bottom surface and the nozzle flange defining a front surface and a rear surface; a floating rotor seal coupled to the nozzle flange via a carrier flange, wherein the carrier flange includes a front wall and a rear wall, wherein the nozzle flange is positioned between the front wall and the rear wall, and wherein a flow path is defined between the front wall, the nozzle flange, and the rear wall; a sealing recess defined in one of the front wall or the rear wall, wherein the sealing recess is in fluid communication with the flow path; and at least one linear sealing section partially disposed within the sealing recess, wherein in response to pressurization of the sealing recess via working fluid in the flow path, the at least one linear sealing section forms a seal against the nozzle flange or the bottom surface.

[0061] According to any of the preceding or following clauses, the turbine, wherein the front wall defines a rearward-facing surface oriented toward the front side surface of the nozzle flange, wherein the sealing recess is defined along the rearward-facing surface.

[0062] According to any of the preceding or subsequent clauses of the turbine, wherein the sealing recess is radially offset from the top surface of the front wall.

[0063] According to any of the preceding or following clauses, the turbine, wherein the rear wall defines a forward-facing surface oriented toward the rear side surface of the nozzle flange, wherein the sealing recess is defined along the forward-facing surface.

[0064] According to any of the preceding or subsequent clauses of the turbine, wherein the sealing recess is radially offset from the top surface of the rear wall.

[0065] According to any of the preceding or following clauses, the turbine, wherein the nozzle flange includes a secondary flange defining a front and a rear, wherein the rear wall of the carrier flange defines a rear surface, wherein the sealing recess is defined along the rear surface of the rear wall, and wherein the sealing recess is oriented toward the front of the secondary flange.

[0066] According to any of the preceding or following clauses, the sealing recess is defined along the top surface of the front wall and toward the bottom surface of the inner shroud.

[0067] According to any of the preceding or subsequent clauses, the at least one linear sealing section includes a first linear sealing section and a second linear sealing section that are at least partially disposed in the sealing recess.

[0068] According to any of the preceding or subsequent clauses, the end of the first linear sealing section overlaps with the adjacent end of the second linear sealing section.

[0069] According to any of the preceding or subsequent clauses, the turbine, wherein the at least one linear sealing section comprises a plurality of linear sealing sections arranged in a ring around the longitudinal centerline of the turbine.

[0070] The turbine according to any of the foregoing or subsequent clauses further includes a biasing member disposed within the sealing recess, wherein the biasing member is configured to apply at least one of a radial force and an axial force to the at least one linear sealing section.

[0071] According to any of the preceding or subsequent clauses, the biasing member is a wave spring.

[0072] The turbine according to any of the foregoing or subsequent clauses further includes a rotor shaft having an outer surface, wherein the nozzle flange and the floating rotor seal are disposed between the inner shroud and the outer surface of the rotor shaft.

[0073] The turbine according to any of the preceding or subsequent clauses, wherein the inner shroud, the nozzle flange, the floating rotor seal, and the rotor shaft at least partially define a first pressure chamber and a second pressure chamber.

[0074] According to any of the preceding or subsequent clauses of the turbine, wherein the at least one linear sealing section is configured to prevent the working fluid from flowing from the first pressure chamber to the second pressure chamber.

[0075] According to any of the preceding or subsequent clauses, the working fluid fills the first pressure chamber, and the first pressure chamber is at a first pressure higher than the second pressure of the second pressure chamber.

[0076] A gas turbine engine includes: a turbine including a high-pressure turbine and a working fluid flowing through the high-pressure turbine, wherein the high-pressure turbine includes: a first fixed member defining a sealing recess; a second fixed member axially adjacent to the first fixed member relative to a longitudinal centerline of the turbine, the second fixed member defining a sealing surface oriented toward the sealing recess; and a linear sealing section partially disposed within the sealing recess, wherein in response to pressurization of the sealing recess via the working fluid, the linear sealing section forms a seal against the sealing surface of the second fixed member.

[0077] According to any of the preceding or subsequent clauses of the gas turbine engine, wherein the first fixed component is an outer shroud of the nozzle section, and the second fixed component is a turbine rotor blade shroud.

[0078] According to any of the preceding or subsequent clauses, the gas turbine engine further includes a biasing member, wherein the biasing member is configured to apply at least one of a radial force and an axial force to the linear sealing section.

[0079] In any of the preceding or subsequent clauses of the gas turbine engine, the biasing member is a wave spring.

[0080] A method for sealing a first pressure chamber and a second pressure chamber of a turbine. The method includes pressurizing a sealing recess with a compressed working fluid at a first pressure, wherein the compressed working fluid applies a radially outward force relative to the radial direction to a linear sealing section, thereby causing the linear sealing section to abut against the bottom surface of an inner shroud of a nozzle section of the turbine, wherein the linear sealing section prevents or blocks the compressed working fluid from leaking from the first pressure chamber of the turbine or flowing into the second pressure chamber.

[0081] A method for sealing a first pressure chamber and a second pressure chamber of a turbine. The method includes pressurizing a sealing recess with a compressed working fluid at a first pressure, wherein the compressed working fluid applies a radially inward force relative to the radial direction to a linear sealing section disposed within the sealing recess, thereby causing the linear sealing section to abut against the lower surface of the sealing recess via the compressed working fluid.

[0082] According to the method for sealing the first and second pressure chambers of a turbine as described in the foregoing clause, the compressed working fluid applies a rearward axial force relative to the axial direction to the linear sealing section, thereby pressing the linear sealing section against the front of the secondary flange of the nozzle flange.

[0083] This written description uses examples to disclose the contents of this disclosure, including best practices, and to enable those skilled in the art to practice this disclosure, including making and using any device or system and performing any combined methods. The patentable scope of this disclosure is defined by the claims and may include other examples that would occur to those skilled in the art. Such other examples are intended to be within the scope of the claims if they include structural elements that are not indistinguishable from the literal language of the claims, or if they include equivalent structural elements that are not substantially indistinguishable from the literal language of the claims.

Claims

1. A turbine, characterized in that, include: A nozzle section, the nozzle section including an inner shroud and a nozzle flange, the inner shroud defining a bottom surface, and the nozzle flange defining a front side surface and a rear side surface; and A floating rotor seal, the floating rotor seal being connected to the nozzle flange via a carrier flange, wherein the carrier flange includes: A front wall and a rear wall, wherein the nozzle flange is positioned between the front wall and the rear wall, and wherein a flow path is defined between the front wall, the nozzle flange and the rear wall; A sealing recess, the sealing recess being defined in one of the front wall or the rear wall, wherein the sealing recess is in fluid communication with the flow path; and At least one linear sealing section is partially disposed within the sealing recess, wherein, in response to pressurization of the sealing recess by the working fluid in the flow path, the at least one linear sealing section forms a seal against the nozzle flange or the bottom surface.

2. The turbine according to claim 1, characterized in that, The front wall defines a rearward-facing surface oriented toward the front side surface of the nozzle flange, wherein the sealing recess is defined along the rearward-facing surface.

3. The turbine according to claim 2, characterized in that, The sealing recess is radially offset from the top surface of the front wall.

4. The turbine according to claim 1, characterized in that, The rear wall defines a forward-facing surface oriented toward the rear side surface of the nozzle flange, wherein the sealing recess is defined along the forward-facing surface.

5. The turbine according to claim 4, characterized in that, The sealing recess is radially offset from the top surface of the rear wall.

6. The turbine according to claim 1, characterized in that, The nozzle flange includes a secondary flange defining a front and a rear, the rear wall of the carrier flange defining a rear surface, the sealing recess defining a rear surface along the rear wall, and the sealing recess oriented toward the front of the secondary flange.

7. The turbine according to claim 1, characterized in that, The front wall of the carrier flange defines a top surface, and the sealing recess is defined along the top surface of the front wall and oriented toward the bottom surface of the inner shroud.

8. The turbine according to claim 1, characterized in that, The at least one linear sealing segment includes a first linear sealing segment and a second linear sealing segment that are at least partially disposed in the sealing recess.

9. The turbine according to claim 8, characterized in that, The end of the first linear sealing section overlaps with the adjacent end of the second linear sealing section.

10. The turbine according to claim 1, characterized in that, The at least one linear sealing section comprises a plurality of linear sealing sections arranged in a ring around the longitudinal centerline of the turbine.

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