Ring guard assembly

Abstract

A shroud assembly is provided that includes a shroud support and an annular shroud. The shroud assembly includes one or more pins for securing the annular shroud to the shroud support. The pins have blocks that can translate radially to allow the shroud to expand and contract in the radial direction. A gas turbine engine is also provided that includes a compressor section, a combustion section, a turbine section, and a shroud assembly. The shroud assembly includes one or more pins for securing a continuous shroud to the shroud support. The pins have blocks that can translate radially to allow the shroud to expand and contract in the radial direction. A method for assembling a shroud assembly structure in a gas turbine engine is also provided.

Description

[0001] This invention was made with Government support under Award W58RGZ-16-C-0047 from the U.S. Army. The Government has certain rights in this invention. Technical Field

[0002] The present subject matter generally relates to gas turbine engines. More particularly, the present subject matter relates to a shroud assembly for a gas turbine engine. Background Art

[0003] The efficiency of a gas turbine engine depends on many factors, one of which is the radial clearance between adjacent rotating and non-rotating components, such as the rotor blade tips and the shrouds surrounding the outer tips of the rotor blades. If the clearance is too large, unacceptable levels of gas leakage can occur, resulting in a loss of efficiency. If the clearance is too small, there is a risk of contact between the components under certain conditions.

[0004] During operation, temperature differences across the engine often cause rotating and non-rotating components to expand and contract radially at different rates. Therefore, an improved shroud assembly is needed to maintain proper clearance between the rotor blade tips and the shroud during thermal expansion and contraction. Summary of the Invention

[0005] In one aspect, embodiments of the present disclosure relate to a shroud assembly. The shroud assembly includes a shroud support and an annular shroud, the annular shroud defining a radial centerline along a circumferential direction. The annular shroud includes one or more shroud retaining features having a circumferential surface. For example, the annular shroud may include three or more shroud retaining features. One or more pins are also provided for securing the annular shroud to the shroud support. For example, in an embodiment, three or more pins are provided for securing the annular shroud to the shroud support. Each pin has an outer pin coupled to an inner pin, the inner pin being arranged radially inward from the outer pin. The inner pin has a block thereon that is capable of radially translating along the inner pin. The block engages the circumferential surface of the one or more retaining features for securing the annular shroud to the shroud support.

[0006] In another aspect, an embodiment of the present disclosure relates to a gas turbine engine. The gas turbine engine includes a compressor section, a combustion section, and a turbine section that are in a series flow relationship and together define a core air flow path. The gas turbine engine includes a shroud assembly positioned in at least one of the compressor section or the turbine section and at least partially defining the core air flow path, the shroud assembly including a support structure and an annular shroud that defines a radial centerline along a circumferential direction. The annular shroud includes one or more retaining features having a circumferential surface that is configured to engage one or more pins to secure the annular shroud to the support, each of the one or more pins having an outer pin coupled to the support structure and an inner pin disposed radially inward from the outer pin, the inner pin having a block that can translate radially thereon along the inner pin. The block engages the circumferential surface of the one or more retaining features to secure the annular shroud to the support structure.

[0007] In another aspect, embodiments of the present disclosure relate to a method for assembling a shroud assembly structure in a gas turbine engine. The method includes providing at least a portion of a gas turbine engine having one or more shroud support structures, the one or more shroud supports having a radially inner surface; disposing an annular shroud clearance tool between the radially inner surface of the shroud support and an annular shroud to ensure clearance between the radially inner surface and the annular shroud, the annular shroud having one or more shroud retaining features; disposing a block in the one or more shroud retaining features, the block coupled to an inner pin coupled to an outer pin, the block being capable of translating in a radial direction relative to the inner pin; rotating the outer pin or the inner pin to translate the block so that the block frictionally engages a circumferential surface on the one or more shroud retaining features to couple the annular shroud to the block; optionally, adjusting the torque of the pin to further secure the annular shroud to the block; optionally, securing the outer pin to the shroud support structure; and removing the annular shroud clearance tool. BRIEF DESCRIPTION OF THE DRAWINGS

[0008] The above and other aspects, features and advantages of the present disclosure will become more apparent from the following detailed description when taken in conjunction with the accompanying drawings, in which:

[0009] Figure 1 is a schematic cross-sectional view of an exemplary gas turbine engine according to various embodiments of the present disclosure.

[0010] Figure 2 yes Figure 1 a side cross-sectional view of the compressor section, combustion section, and high-pressure turbine section of the illustrated gas turbine engine;

[0011] Figure 3 is a cross-sectional view of a shroud assembly according to various embodiments of the present disclosure;

[0012] Figure 4 is a perspective view of an annular shroud according to various embodiments of the present disclosure;

[0013] Figure 5 is an enlarged perspective view of an annular shroud according to various embodiments of the present disclosure;

[0014] Figure 6 is a perspective view of a portion of a shield assembly according to various embodiments of the present disclosure;

[0015] Figure 7 is a top view of a shield assembly according to various embodiments of the present disclosure.

[0016] Figure 8 is a top view of a shield assembly according to various embodiments of the present disclosure.

[0017] Figure 9 is a top view of a portion of a shroud assembly according to various embodiments of the present disclosure.

[0018] Figure 10 is an axial view of a portion of a shroud assembly according to various embodiments of the present disclosure;

[0019] Figure 11 is an axial view of a portion of a shroud assembly according to various embodiments of the present disclosure; and

[0020] Figure 12 is a cross-sectional view of a shroud assembly according to various embodiments of the present disclosure. DETAILED DESCRIPTION

[0021] Reference will now be made in detail to the present embodiments of the invention, 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 invention. As used herein, the terms "first," "second," and "third" are used interchangeably to distinguish one component from another and are not intended to indicate the position or importance of the various components. The terms "upstream" and "downstream" refer to the relative flow direction relative to the flow of a fluid in a fluid pathway. For example, "upstream" refers to the flow direction of an incoming fluid flow and "downstream" refers to the flow direction of an outgoing fluid flow. "HP" means high pressure and "LP" means low pressure.

[0022] In addition, as used herein, the terms "axial" or "axially" refer to a dimension along the longitudinal axis of the engine. The term "front" used in conjunction with "axial" or "axially" refers to a direction toward the engine inlet, or a component that is relatively closer to the engine inlet than another component. The term "rear" used in conjunction with "axial" or "axially" refers to a direction toward the engine exhaust nozzle, or a component that is relatively closer to the engine exhaust nozzle than another component. The term "radial" or "radially" refers to a dimension extending between the central longitudinal axis (or centerline) of the engine and the outer circumference of the engine. Radially inward toward the longitudinal axis and radially outward away from the longitudinal axis.

[0023] The present disclosure may include, consist essentially of, or consist of the components of the present disclosure and other materials described herein. As used herein, "consisting essentially of" means that a composition or component may include additional materials, but only if the additional materials do not materially alter the basic and novel characteristics of the claimed composition or method.

[0024] Each embodiment is provided for the purpose of explaining the invention and not for limitation thereof. In fact, it will be apparent to those skilled in the art that modifications and variations may be made to the invention without departing from the scope or spirit of the invention. For example, features shown or described as part of one embodiment may be used in another embodiment to produce yet another embodiment. Therefore, the present invention is intended to cover such modifications and variations that fall within the scope of the appended claims and their equivalents. Although exemplary embodiments of the invention will be generally described in the context of a turbine shroud incorporated into a turbofan jet engine for illustrative purposes, it will be readily understood by those skilled in the art that embodiments of the invention may be applied to any turbine incorporated into any turbine and are not limited to gas turbofan jet engines unless specifically recited in the claims.

[0025] Exemplary aspects of the present disclosure relate to a shroud assembly including an annular shroud coupled to a support structure. During operation of a gas turbine engine, the shroud experiences thermal expansion and contraction due to varying operating temperatures. As the shroud expands or contracts, it is important that the shroud remains centered along its radial centerline to provide adequate clearance between the blades and the shroud.

[0026] Furthermore, due to the size of the tip clearance, the turbine shroud directly impacts the overall efficiency or performance of the gas turbine engine. The turbine shroud further impacts engine performance because any compressor discharge and / or bleed air used to cool the turbine shroud is unused during the combustion process or during the work expansion process of the turbine blades and is therefore unavailable for producing useful work. Therefore, it is desirable to control or reduce the amount of bleed air used to cool the turbine shroud to maximize the overall efficiency of the engine. It is also desirable to use CMC materials in the shroud because they have higher temperature resistance than metal-based materials.

[0027] Therefore, in one aspect, a shroud assembly is provided that includes a shroud support and an annular shroud. The annular shroud is continuous in the circumferential direction, so that no splines are required to radially seal the shroud segments, effectively reducing the number of components in the shroud assembly, reducing the complexity of the assembly, and reducing the bleed air required to purge the end gap. In addition, the shroud assembly includes one or more pins for securing the annular shroud to the support structure. The pins include an outer pin secured to the shroud support and an inner pin having a block thereon. The inner pin is offset from the centerline of the outer pin, which allows the annular shroud to be securely held in the center of the engine centerline. In addition, the offset inner pin creates a cam mechanism, and rotation of the inner pin causes the block to translate. The block engages the annular shroud and can translate in a radial direction, thereby allowing thermal expansion or contraction of the annular shroud while maintaining the position of the annular shroud about the radial centerline. Improved sealing can therefore improve the performance and efficiency of the engine.

[0028] Referring now to the accompanying drawings, Figure 1 is a schematic cross-sectional view of a gas turbine engine 100 according to an exemplary embodiment of the present disclosure. More specifically, Figure 1 In an embodiment, the gas turbine engine 100 is an aerospace, high-bypass turbofan jet engine configured to be mounted to an aircraft, such as in an underwing or tail-mount configuration. Figure 1 As shown, gas turbine engine 100 defines an axial direction A1 (extending parallel or coaxially with longitudinal centerline 102 for reference) and a radial direction R1 . Generally, gas turbine engine 100 includes a fan section 104 and a core turbine engine 106 disposed downstream of fan section 104 .

[0029] The depicted exemplary core turbine engine 106 generally includes a substantially tubular outer casing 108 defining an annular inlet 110. Outer casing 108 encloses, in series flow relationship, a compressor section 112 including a first supercharger or LP compressor 114 and a second HP compressor 116; a combustion section 118; a turbine section 120 including a first HP turbine 122 and a second LP turbine 124; and an exhaust nozzle section 126. An HP shaft or spool 128 drivingly connects the HP turbine 122 to the HP compressor 116. An LP shaft or spool 130 drivingly connects the LP turbine 124 to the LP compressor 114. Together, the compressor section, combustion section 118, turbine section, and exhaust nozzle section 126 define a core air flow path 132 through the core turbine engine 106.

[0030] Still refer to Figure 1 In the embodiment of the fan section 104, the fan section 104 includes a variable pitch fan 134 having a plurality of fan blades 136 coupled to a disk 138 in a circumferentially spaced apart manner. As shown, the fan blades 136 extend generally outwardly from the disk 138 in a radial direction R. Each fan blade 136 is rotatable relative to the disk 138 about a pitch axis because the fan blades 136 are operably coupled to a suitable actuating member 140 that is configured to collectively change the pitch of the fan blades 136, for example, in unison. The fan blades 136, disk 138, and actuating member 140 are rotatable together about the longitudinal centerline 102 via the LP shaft 130 through a power gearbox 142. The power gearbox 142 includes a plurality of gears for reducing the rotational speed of the LP shaft 130 to a more efficient fan speed.

[0031] Still refer to Figure 1 In the exemplary embodiment, disk 138 is covered by a rotatable forward nacelle 144 having an aerodynamic profile to facilitate airflow through the plurality of fan blades 136. Furthermore, exemplary fan section 104 includes an annular fan case or outer nacelle 146 that circumferentially surrounds fan 134 and / or at least a portion of core turbine engine 106. Furthermore, for the depicted embodiment, nacelle 146 is supported relative to core turbine engine 106 by a plurality of circumferentially spaced outlet guide vanes 148. Furthermore, a downstream section 150 of nacelle 146 extends above the exterior of core turbine engine 106 to define a bypass airflow passage 152 therebetween.

[0032] During operation of gas turbine engine 100, a quantity of air 154 enters gas turbine engine 100 through nacelle 146 and / or an associated inlet 156 of fan section 104. As quantity of air 154 passes through fan blades 136, a first portion of air 154, as indicated by arrow 158, is directed or channeled into bypass airflow passage 152, and a second portion of air 154, as indicated by arrow 160, is directed or channeled into LP compressor 114. The pressure of second portion of air 160 increases as it passes through high pressure (HP) compressor 116 and enters combustion section 118.

[0033] Still refer to Figure 1 , a second portion of compressed air 160 from the compressor section is mixed with fuel and combusted within combustion section 118 to provide combustion gases 162. Combustion gases 162 are directed from combustion section 118 along a hot gas path 174 through HP turbine 122, wherein a portion of thermal and / or kinetic energy from combustion gases 162 is extracted via sequential stages of HP turbine rotor blades 164 coupled to outer casing 108 and HP turbine rotor blades 166 coupled to HP shaft or spool 128, thereby rotating HP shaft or spool 128 and thereby supporting operation of HP compressor 116. Combustion gases 162 are then directed through LP turbine 124, wherein a second portion of thermal and kinetic energy is extracted from combustion gases 162 via sequential stages of LP turbine stator vanes 168 coupled to outer casing 108 and LP turbine rotor blades 170 coupled to LP shaft or spool 130, thereby rotating LP shaft or spool 130 and thereby supporting operation of LP compressor 114 and / or rotation of fan 134.

[0034] Combustion gases 162 are then directed through exhaust nozzle section 126 of core turbine engine 106 to provide propulsive thrust. Simultaneously, the pressure of first portion air 158 is significantly increased as first portion air 158 is directed through bypass airflow passage 152 before being discharged from fan nozzle exhaust section 172 of gas turbine engine 100, also providing propulsive thrust. HP turbine 122, LP turbine 124, and exhaust nozzle section 126 at least partially define a hot gas path 174 for directing combustion gases 162 through core turbine engine 106.

[0035] It should be understood that Figure 1The exemplary gas turbine engine 100 depicted in FIG is merely an example, and in other exemplary embodiments, the gas turbine engine 100 may have any other suitable configuration. Additionally or alternatively, aspects of the present disclosure may be used with any other suitable aviation gas turbine engine, such as a turboshaft engine, a turboprop engine, a turbojet engine, a low-bypass turbofan engine, etc. Furthermore, aspects of the present disclosure may further be used with any other land-based gas turbine engine (e.g., a power generation gas turbine engine) or any aeroderivative gas turbine engine (e.g., a marine gas turbine engine).

[0036] It should be understood that in other exemplary embodiments, aspects of the present disclosure may be incorporated into any other exemplary gas turbine engine. For example, in other exemplary embodiments, aspects of the present disclosure may be incorporated into a turboshaft engine. In such cases, the gas turbine engine may not include a fan section, an outer nacelle, or an array of outlet guide vanes extending between the cowling of the gas turbine engine and the nacelle of the gas turbine engine (e.g., through a bypass airflow channel). Furthermore, with such a configuration, the gas turbine engine may be coupled to a load, such as a generator, a rotor assembly (e.g., when incorporated into a helicopter), etc., via the engine's low-pressure shaft.

[0037] Figure 2 supply Figure 1 FIG2 is a side cross-sectional view of the compressor section 112, combustion section 118, and turbine section 120 of the core turbine engine 106. More specifically, the aft end of the HP compressor 116, the combustor section 118, and the forward end of the HP turbine 122 are shown.

[0038] Compressed air 176 exits the HP compressor 116 through a diffuser 178 located at the rear end or outlet of the HP compressor 116 and diffuses into the combustion section 118. The combustion section 118 of the core turbine engine 106 is annularly surrounded by a radially inner combustor outer casing 180 and an outer combustor outer casing 182. The radially inner combustor outer casing 180 and the radially outer combustor outer casing 182 both extend generally in the axial direction A1 and surround a combustor assembly 184 in an annular ring. The inner combustor outer casing 180 and the outer combustor outer casing 182 are joined together at the annular diffuser 178 at the front end of the combustion section 118.

[0039] As shown, combustor assembly 184 generally includes an inner liner 186 extending generally in an axial direction A1 between an aft end 188 and a forward end 190, and an outer liner 192 also extending generally in axial direction A1 between an aft end 194 and a forward end 196. Together, inner liner 186 and outer liner 192 at least partially define a combustion chamber 198 therebetween. Inner liner 186 and outer liner 192 are each attached to or integrally formed with an annular dome. More specifically, the annular dome includes an inner dome segment 200 integrally formed with forward end 190 of inner liner 186 and an outer dome segment 202 generally formed with forward end 196 of outer liner 192. Furthermore, inner dome segment 200 and outer dome segment 202 may each be integrally formed (or alternatively, may be formed from multiple components attached in any suitable manner) and may each extend circumferentially to define an annular shape. However, it should be understood that in other embodiments, the combustor assembly 184 may not include the inner dome segment 200 and / or the outer dome segment 202; may include separately formed inner dome segments 200 and / or outer dome segments 202 that are attached to the respective inner and outer liners 186 and 192; or may have any other suitable configuration.

[0040] Still refer to Figure 2 The combustor assembly 184 further includes a plurality of fuel-air mixers 204 that are spaced circumferentially and at least partially positioned within the annular dome. More specifically, the plurality of fuel-air mixers 204 are at least partially disposed between the outer dome section 202 and the inner dome section 200 along the radial direction R1. Compressed air 176 from the compressor section 112 of the gas turbine engine 100 flows into or through the fuel-air mixers 204, where the compressed air 176 is mixed with fuel and ignited to produce combustion gases 162 within the combustion chamber 198. The inner dome section 200 and the outer dome section 202 are configured to facilitate providing such a flow of compressed air 176 from the compressor section 112 into or through the fuel-air mixers 204.

[0041] As described above, combustion gases 162 flow from combustion chamber 198 and through turbine section 120 of gas turbine engine 100, wherein a portion of the thermal and / or kinetic energy from the combustion gases 162 is extracted via successive stages of turbine stator vanes and turbine rotor blades within HP turbine 122 and LP turbine 124. More specifically, as Figure 2 As shown, combustion gases 162 from a combustion chamber 198 flow into the HP turbine 122 immediately downstream of the combustion chamber 198 , where thermal and / or kinetic energy from the combustion gases 162 is extracted via successive stages of HP turbine rotor blades 164 , 166 .

[0042] like Figure 2As shown, not all of the compressed air 176 flows into or directly passes through the fuel-air mixer 204 and into the combustion chamber 198. Some of the compressed air 176 is discharged into a plenum 206 surrounding the combustor assembly 184. The plenum 206 is generally defined between the outer combustor casings 180, 182 and the liners 186, 192. The outer combustor casing 182 and the outer liner 192 define an outer plenum 208, which is generally disposed radially outward from the combustion chamber 198. The inner combustor casing 180 and the inner liner 186 define an inner plenum 210, which is generally disposed radially inward relative to the combustion chamber 198. As the compressed air 176 is diffused by the diffuser 178, some of the compressed air 176 flows radially outward into the outer plenum 208, and some of the compressed air 176 flows radially inward into the inner plenum 210.

[0043] The compressed air 176 flowing radially outward into the outer plenum 208 generally flows axially toward the turbine section 120. Specifically, the compressed air 176 flows over the HP turbine 122 stator vanes 164 and rotor blades 166. The outer plenum 208 may also extend to the LP turbine 124 ( Figure 1 ).

[0044] like Figure 2 As further shown, the HP turbine 122 includes one or more shroud assemblies 300, each shroud assembly 300 forming an annular shroud ring around the annular array of HP turbine rotor blades 164 and 166. In this example, the annular shroud ring is circumferentially disposed around the annular array of rotor blades 164 of the first stage 212 of the HP turbine 122, and the annular ring is circumferentially disposed around the annular array of turbine rotor blades 166 of the second stage 214. Typically, the annular shrouds of the shroud assemblies 300 are spaced apart from the blade tips 216 of each of the rotor blades 164 and 166. The shroud assemblies 300 generally reduce radial leakage into and out of the core air flow path 132 and may also reduce axial leakage.

[0045] It should be noted that the shroud assembly 300 may also be used in a similar manner in the LP compressor 114, the HP compressor 116, and / or the LP turbine 124. Thus, the shroud assembly 300 disclosed herein is not limited to use in the HP turbine 122, but may be used in any suitable section of the gas turbine engine 100, or a turbine engine more generally. The shroud assembly 300 includes an annular shroud 302 coupled to a shroud support 304. The shroud support 304 may be a hanger. The shroud support 304 couples to and supports the annular shroud 302 in the gas turbine engine 100 and is itself supported by various other components in the gas turbine engine 100. The shroud support 304 may be a multi-piece hanger or may be formed from a single piece. In the exemplary embodiment, the shroud support 304 is a single-piece hanger. The annular shroud 302 includes a radially inner side 308 that faces one or more blade tips 216.

[0046] The compressed air 176 flowing through the outer plenum 208 has a pressure of either P1 or P3, which exerts a radially inward force on the outer side 306 of the annular shroud 302. The combustion gases 162 flowing through the hot gas path 174 of the HP turbine have a pressure of either P2 or P4, which exerts a radially outward force on the inner side 308 of the shroud segment 302. It should be understood that during operation of the gas turbine engine 100, relative to this portion of the core turbine engine 106, P1 is generally greater than P2 and P3 is generally greater than P4. It should also be understood that in some circumstances, such as when the gas turbine engine 100 is not operating or when the gas turbine engine 100 is experiencing a stall, P1 may not be greater than P2 and / or P3 may not be greater than P4.

[0047] In some embodiments, components of the gas turbine engine 100, particularly components within the hot gas path, such as components within or downstream of the combustion section, may include ceramic matrix composite (CMC) materials, which are non-metallic materials with high temperature resistance capabilities. For example, the annular shroud 302 may be formed from a CMC material. Generally speaking, the performance and efficiency of a turbine can be improved by increasing the combustion gas temperature; therefore, non-traditional high temperature materials, such as CMC materials, are more commonly used for various components within the gas turbine engine, including components within the combustion gas flow path. Exemplary CMC materials for gas turbine engine components may include silicon carbide (SiC), silicon, silicon dioxide, or alumina matrix materials, and combinations thereof. Ceramic fibers may be embedded in a matrix, such as oxidation-stabilized reinforcing fibers, including monofilaments such as sapphire and silicon carbide (e.g., Textron's SCS-6), and rovings and yarns including silicon carbide (e.g., Nippon Carbon's Ube Industries and Dow Corning ), aluminum silicates (such as Nextel's 440 and 480), and chopped whiskers and fibers (such as Nextel's 440 and ), and optionally ceramic particles (e.g., oxides of silicon, aluminum, zirconium, yttrium, and combinations thereof) and inorganic fillers (e.g., pyrophyllite, wollastonite, mica, talc, kyanite, and montmorillonite). For example, in certain embodiments, the fiber bundles that may include a ceramic refractory coating are formed into reinforcement tapes, such as unidirectional reinforcement tapes. Multiple tapes can be stacked together (e.g., as plies) to form a preformed component. The fiber bundles can be impregnated with the slurry composition before or after forming the preform. The preform can then be subjected to a heat treatment, such as curing or burnout to produce a high carbon residue in the preform, and subsequent chemical treatment, such as infiltration with silicon melt, to obtain a component formed of a CMC material having the desired chemical composition. In other embodiments, the CMC material can be formed, for example, as carbon fiber cloth instead of a tape.

[0048] As described above, components comprising CMC materials may be used within the hot gas path, such as within the combustion and / or turbine sections of engine 100. However, CMC components may also be used in other sections, such as the compressor and / or fan sections. As a specific example, described in more detail below, a turbine shroud may be formed from a CMC material to better withstand increased combustion gas temperatures.

[0049] Figure 3 A cross-sectional view of an exemplary shroud assembly 300 is provided. The shroud assembly 300 is positioned in at least one of the compressor section 112 and / or the turbine section 120 and at least partially defines the core air flow path 132. As an example, Figure 2 As shown, the shroud assembly 300 may be positioned to circumferentially surround the rotor blades 164 and 166 of the HP turbine 122. In other embodiments, the shroud assembly 300 may be located in other portions or locations within the gas turbine engine 100.

[0050] For this embodiment, the shroud assembly 300 includes an annular shroud 302 coupled to a shroud support 304. The shroud support 304 includes a radially inner side 310 that faces an outer side 306 of the annular shroud 302. The annular shroud 302 includes a radially inner side 308 that faces one or more blade tips 216 of the rotor blades 164. The annular shroud includes one or more shroud retention features 500.

[0051] One or more pins 400 are used to secure the annular shroud 302 to the support structure 304. The pins 400 include an outer pin 402 and an inner pin 404 having a block 406 thereon. The outer pin 402 may be secured to the support structure 304 by spot welding or any other suitable securing means. The block 406 fits loosely over the inner pin 404, enabling the block 406 to rotate about the inner pin 404 and also to translate the inner pin 404 radially upward and downward. In one embodiment, the outer pin 402 and the inner pin 404 may be integrally formed. However, in other embodiments, the inner pin 404 and the outer pin 402 may be two separate components that are subsequently connected together, for example, by a press fit. The inner pin 404 is offset from the centerline of the outer pin 402. Offsetting the inner pin 404 from the centerline of the outer pin 402 allows the annular shroud 302 to be securely held centered on the engine centerline. The offset inner pin 404 creates a cam mechanism, wherein rotation of the pin 400 causes the block 406 to translate. Although the term "block" is used, the present disclosure is not limited to cuboid or rectangular blocks. For example, the block 406 can be any suitable shape as long as it is capable of engaging the annular shroud 302 and the pin 400 according to various embodiments of the present disclosure. For example, suitable block shapes may include rectangular, cuboid, spherical, and / or cylindrical blocks. Additional details regarding the installation of the annular shroud 302 using one or more pins 400 disclosed herein will be further discussed below. In addition to radially positioning the annular shroud 302, the pin 400 prevents the shroud from rotating about the engine centerline while allowing radial and axial thermal growth of the annular shroud 302.

[0052] Referring now to FIG4 , a perspective view of an exemplary annular shroud 302 is shown. A circumferential direction C1 is shown. The annular shroud 302 includes a radially inner side 308 and one or more shroud retention features 500 disposed on the outer side 306 of the annular shroud 302. The shroud retention features 500 include one or more of cavities, recesses, slots, and the like. The shroud retention features 500 may be any shape suitable for positioning the block 406 within the shroud retention features 500. For example, the shroud retention features 500 may be square, rectangular, triangular, circular, oval, and the like. As shown, the annular shroud 302 may include a plurality of shroud retention features 500 disposed in a circumferential direction around the annular shroud 302. For example, in an exemplary embodiment, the annular shroud 302 includes at least one shroud retention structure 500, such as at least two shroud retention structures 500, such as at least three shroud retention structures 500, such as at least four shroud retention structures 500, and such as at least five shroud retention structures 500. The shield retention features 500 can be spaced apart from each other around the circumferential direction C1 of the annular shield 302. However, in an embodiment, the positions of the shield retention features 500 can be spaced apart from each other around the circumferential direction C1 of the annular shield 302. In certain embodiments, the shield retention features 500 can be arranged equidistant from each other around the circumferential direction C1 of the annular shield 302. However, in other embodiments, one or more shield retention features 500 can be spaced unequally from each other. For example, it is envisioned that one or more shield retention features 500 can be spaced closer to each other around the circumferential direction C1 of the annular shield 302, while other shield retention features 500 can be spaced farther from each other around the circumferential direction C1 of the annular shield 302.

[0053] refer to Figure 5 , the outer side 306 of the annular shroud 302 includes one or more shroud retention features 500. As shown, the shroud retention feature 500 may include one or more circumferential surfaces 502a, 502b, at least one rear surface 504, and at least one front surface 506. The one or more circumferential surfaces 502a, 502b are positioned along the circumferential direction C1 of the annular shroud 302. The rear surface 504 is spaced apart from the front surface 506. The one or more circumferential surfaces 502a, 502b, the rear surface 504, and the front surface 506 form the side surfaces of the shroud retention feature 500 that extend radially into the annular shroud 302. Although Figure 5 The shield retaining feature 500 shown is essentially rectangular, but the present disclosure is not limited thereto. In fact, the shield retaining member 500 can be formed by other shapes as long as the shape is suitable for engaging with the pin 400 to fix the annular shield 302 to the shield support 304, as will be discussed further below.

[0054] like Figure 6As shown, the pin 400 can be disposed in a shroud retention feature 500 located on the outer side 306 of the annular shroud 302. The pin 400 is disposed in the shroud retention feature 500 such that a portion of the block 406 disposed on the inner pin 404 contacts and frictionally engages one of the circumferential surfaces 502b of the shroud retention feature 500. As shown, a portion of the inner pin 404 can also be disposed within the shroud retention feature 500 without any portion of the outer pin 402 being disposed within the shroud retention feature 500.

[0055] Figure 7-8 A schematic top view of the pin 400 disposed within the shield retention feature 500 is shown. The placement of the outer pin 402, the inner pin 404, and the block 406 are shown within the shield retention feature 500. Specifically, the placement of the block 406 relative to the circumferential surfaces 502a, 502b, the rear surface 504, and the front surface 506 is shown. For example, Figure 7 The block 406 is shown in a rearward position relative to the outer pin 402 and the inner pin 404. However, Figure 8 The block 406 is shown in a forward position relative to the outer pin 402 and the inner pin 404. As shown, the shield retention feature 500 is sized so that the block 406 can be positioned therein in various rearward or forward positions without contacting the rear surface 504 or the front surface 506 of the shield retention feature 500.

[0056] Now refer to Figure 9 , rotation of the internal pin 404 causes the block 406 to translate along an arc within the shroud retention feature 500 in a circumferential direction as indicated by arrow C. Assuming the block is placed within the shroud retention feature 500, the block 406 does not contact the rear surface 504 or the front surface 506 of the shroud retention feature 500. Instead, the pin 400 can be rotated to a desired torque such that the shroud engagement surface 410 of the block 406 contacts one of the circumferential shroud surfaces 502a, 502b. Once engaged, the pin 400 can be tightened such that the friction fit between the shroud engagement surface 410 and the circumferential surfaces 502a, 502b is secured.

[0057] Now refer to Figure 10-11, showing an axial view along the circumferential direction C1. An annular shroud 302 having an outer side 306 with a shroud retention feature 500 is secured to the shroud support 304 by pins 400. The inner side 308 of the annular shroud 302 faces the blade tips 216 of the turbine blades 164. The outer side 306 of the annular shroud 302 faces the radial inner side 310 of the shroud support 304. The outer pins 402 are secured to the shroud support 304, while the inner pins 404 and the blocks 406 are disposed within the shroud retention feature 500. The blocks 406 are capable of radial movement about the inner pins 404, thereby allowing for thermal expansion or contraction of the annular shroud 302. For example, the shroud engagement surface 410 of the blocks 406 is frictionally secured to the circumferential surface 502b of the shroud retention feature 500. Although the block 406 is shown as being fixed to the circumferential surface 502b, in other embodiments, it would be feasible to fix the shroud engaging surface 410 of the block 406 to the circumferential surface 502a. As the annular shroud 302 expands in the radial direction R1, the block 406 can translate in the radial direction R1 about the inner pin 404, so that both the annular shroud 302 and the block 406 move radially outward, as shown in FIG. Figure 11 Similarly, during thermal contraction, both the annular shroud 302 and the block 406 can translate radially inwardly back to the position of the block 406, more similar to Figure 10 The annular shroud 302 is positioned as shown. In other words, securing the annular shroud 302 to the shroud support 304 using the pin 400 including the block 406 allows the annular shroud 302 to uniformly contract and expand in the radial direction R1 while maintaining the annular shroud 302 centered about the axial centerline. Furthermore, assuming that the pin 400 is secured to the circumferential surfaces 502a, 502b via the block 406, the pin 400 prevents the annular shroud 302 from moving or rotating in the circumferential direction. Furthermore, although not shown in the figures, it is contemplated that the annular shroud 302, or more specifically one of the circumferential surfaces 502a, 502b that engages the block 406, can translate in the circumferential direction relative to the shroud engagement surface 410 of the block 406. That is, the circumferential surfaces 502a, 502b can translate in the circumferential direction relative to the block 406, while the block 406 does not translate about the inner pin 404.

[0058] Although the provided figures illustrate a single embodiment of the pins 400 and the shroud retention features 500, the present disclosure is not limited thereto. For example, in an exemplary embodiment, the number of pins 400 and the shroud retention features 500 on the annular shroud 302 can be equal. For example, the number of pins 400 used to secure the annular shroud 302 to the shroud support 304 can correspond to the number of shroud retention features 500 present on the annular shroud 302. In an embodiment, assuming that the annular shroud 302 is a continuous annular shroud and it is desired to maintain the radial position of the annular shroud 302, at least three shroud retention features 500 and at least three pins 400 can be used. In other embodiments, at least five shroud retention features 500 and at least five pins 400 can be used. It should be understood that any number of shroud retention features 500 and pins 400 can be used to radially secure the annular shroud 302, and the disclosure herein is not limited thereto.

[0059] In embodiments where multiple pins 400 are used to secure the annular shroud 302, the pins 400 can be rotated in either a clockwise or counterclockwise direction to secure the block 406 within the shroud retention feature 500, as disclosed herein. Thus, in embodiments where an even number of pins 400 are used, half of the pins can be secured in a counterclockwise manner, while the other half can be secured in a clockwise manner. In embodiments where an odd number of pins 400 are used, half of the pins 400 can be secured in a counterclockwise manner, while the other half can be secured in a clockwise manner, with the remaining pins secured in either a clockwise or counterclockwise manner.

[0060] A method for assembling a shroud assembly structure in a gas turbine engine is also provided. The method comprises utilizing Figure 12 The annular shroud clearance tool 600 is shown. For example, the annular shroud clearance tool 600 can be placed between the radial inner side 310 of the shroud support 304 and the radial outer side 306 of the annular shroud 302. The annular shroud 302 is placed between the radial inner side 310 of the shroud support 304. The radial inner side 308 of the annular shroud 302 faces the blade tips 216 of the turbine blades 164. One or more pins 400, including the outer pin 402, the inner pin 404, and the block 406, can then be placed within the shroud retention feature 500. The block 406 is disposed on the inner pin so that it can rotate when the outer pin 402 rotates. Therefore, the rotation of the outer pin 402 translates the block 406 in the circumferential direction so that the block 406 can engage one or more circumferential surfaces 502a, 502b on the shroud retention feature 500 to secure the block 406 to the annular shroud 302. Thereafter, the annular shroud 302 is properly positioned relative to the shroud support 304 , the pin 400 may be torqued and tightened and the annular shroud clearance tool 600 may be removed.

[0061] Thus, a method for assembling a shroud assembly structure in a gas turbine engine includes providing at least a portion of a gas turbine engine having one or more shroud support structures, the one or more shroud supports having a radial inner surface; providing an annular shroud gap tool between the radial inner surface of the shroud support structure and an annular shroud to ensure a gap between the radial inner surface and the annular shroud, the annular shroud having one or more shroud retaining features; providing a block in the one or more shroud retaining features, the block being coupled to an inner pin, the inner pin being coupled to an outer pin, the block being capable of translating in a radial direction relative to the inner pin; rotating the outer pin or the inner pin to translate the block so that the block frictionally engages a circumferential surface on the one or more shroud retaining features, thereby coupling the annular shroud to the block; optionally, adjusting the torque of the pin to further secure the annular shroud to the block; optionally, securing the outer pin to the shroud support structure; and removing the annular shroud gap tool.

[0062] Further aspects of the invention are provided by the subject matter of the following clauses:

[0063] 1. A shroud assembly defining an axial direction, a radial direction, and a circumferential direction, comprising: a shroud support; an annular shroud defining a radial centerline along the circumferential direction, the annular shroud including one or more retaining features having a circumferential surface; one or more pins, each pin having an outer pin connected to the shroud support and an inner pin set radially inward from the outer pin, the inner pin having a block capable of radially translating along the inner pin, wherein the block engages a circumferential surface for fixing the annular shroud to the one or more retaining features of the shroud support.

[0064] 2. A shroud assembly according to any of the preceding clauses, wherein the inner pin is offset from the centre line of the outer pin.

[0065] 3. A shield assembly according to any of the preceding clauses, wherein the one or more retaining features include a cavity formed on the outer surface of the annular shield, the cavity having a first circumferential surface, a second circumferential surface, a front surface and a rear surface, wherein the block engages one of the first circumferential surface or the second circumferential surface.

[0066] 4. The shroud assembly of any preceding clause, wherein the block comprises a shroud engaging surface for engaging a circumferential surface of the one or more retention features.

[0067] 5. The shroud assembly of any preceding clause, wherein the annular shroud comprises a plurality of retaining features spaced apart from one another for alignment with one or more pins in a circumferential direction around the annular shroud.

[0068] 6. A shroud assembly according to any preceding clause, wherein the block is capable of translating in a radial direction relative to the first pin due to thermal expansion or contraction of the annular shroud.

[0069] 7. The shroud assembly of any preceding clause, wherein the annular shroud comprises a ceramic matrix composite material.

[0070] 8. The shroud assembly according to any of the preceding clauses, wherein the annular shroud is provided around one or more blades in a turbine section of a gas turbine engine.

[0071] 9. The shroud assembly of any preceding clause, wherein the annular shroud is provided around one or more blades in a compressor section of a gas turbine engine.

[0072] 10. The shroud assembly of any of the preceding clauses, wherein the first pin and the second pin are integrally formed.

[0073] 11. The shroud assembly according to any of the preceding clauses, wherein the annular shroud is capable of uniform expansion and contraction in a radial direction.

[0074] 12. A gas turbine engine defining an axial direction, a radial direction, and a circumferential direction, comprising: a compressor section, a combustion section, and a turbine section in a series flow relationship and together defining a core air flow path; and a shroud assembly positioned in at least one of the compressor section or the turbine section and at least partially defining the core air flow path, the shroud assembly comprising a shroud support and an annular shroud, the annular shroud defining a radial centerline along a circumferential direction; the annular shroud comprising one or more retaining features having a circumferential surface configured to engage one or more pins to secure the annular shroud to the shroud support, each of the one or more multiple pins having an outer pin coupled to the shroud support and an inner pin radially inwardly disposed from the outer pin, the inner pin having a block capable of radially translating along the inner pin, wherein the block engages the circumferential surface of the one or more retaining features for securing the annular shroud to the shroud support.

[0075] 13. The gas turbine engine according to any of the preceding clauses, wherein the inner pin is offset from the centerline of the outer pin.

[0076] 14. A gas turbine engine according to any of the preceding clauses, wherein the one or more retaining features include a cavity formed in the outer surface of the annular shroud, the cavity having a first circumferential surface, a second circumferential surface, a front surface and a rear surface, wherein the block engages one of the first circumferential surface or the second circumferential surface.

[0077] 15. A gas turbine engine according to any preceding clause, wherein the block comprises a shroud engaging surface for engaging a circumferential surface of one or more retention features.

[0078] 16. The gas turbine engine according to any of the preceding clauses, wherein the annular shroud comprises a plurality of retaining features spaced apart from one another to align with one or more pins in a circumferential direction around the annular shroud.

[0079] 17. The gas turbine engine according to any of the preceding clauses, wherein the block is translatable in radial direction relative to the first pin due to thermal expansion or compression of the annular shroud.

[0080] 18. The gas turbine engine according to any of the preceding clauses, wherein the annular shroud comprises a composite matrix material.

[0081] 19. A method for assembling a shroud assembly structure in a gas turbine engine, the method comprising: providing at least a portion of a gas turbine engine having one or more shroud supports, the one or more shroud supports having a radial inner surface; setting an annular shroud gap tool between the radial inner surface of the shroud support and an annular shroud to ensure a gap between the radial inner surface and the annular shroud, the annular shroud having one or more shroud retaining features; setting a block in the one or more shroud retaining features, the block being coupled to an inner pin, the inner pin being coupled to an outer pin, the block being capable of translating in a radial direction relative to the inner pin; rotating the outer pin or the inner pin to translate the block so that the block frictionally engages a circumferential surface on the one or more shroud retaining features, thereby coupling the annular shroud to the block; optionally, adjusting the torque of the pin to further secure the annular shroud to the block; optionally, securing the outer pin to the shroud support; and removing the annular shroud gap tool.

[0082] 20. The method according to any of the preceding clauses, wherein the annular shroud is capable of uniform expansion and contraction in the radial direction.

[0083] This written description uses examples to describe the disclosure, including the best mode, and also to enable any person skilled in the art to practice the disclosure, including making and using any devices or systems and performing any incorporated methods. The patentable scope of the disclosure is defined by the claims, and may include other examples that occur to those skilled in the art. Such other examples are intended to be within the scope of the claims if they have structural elements that do not differ from the literal language of the claims, or if they include equivalent structural elements with insubstantial differences from the literal language of the claims.

Claims

1. A shield assembly defining an axial direction, a radial direction and a circumferential direction, characterized in that: include: Shield support; an annular shroud defining a radial centerline along the circumferential direction, the annular shroud including one or more retention features having a circumferential surface; and One or more pins, each pin having an outer pin connected to the shroud support and an inner pin disposed radially inward from the outer pin, the inner pin having a block thereon capable of translating radially along the inner pin, wherein the inner pin is offset from a radial centerline of the outer pin; wherein the block engages the circumferential surface of the one or more retention features to secure the annular shroud to the shroud support.

2. The shield assembly according to claim 1, wherein: The one or more retention features include a cavity formed on an outer surface of the annular shroud, the cavity having a first circumferential surface, a second circumferential surface, a front surface, and a rear surface, wherein the block engages one of the first circumferential surface or the second circumferential surface.

3. The shield assembly according to claim 1, wherein: Wherein the block includes a shroud engaging surface for engaging the circumferential surface of the one or more retention features.

4. The shield assembly according to claim 1, wherein: The annular shroud includes a plurality of retaining features spaced apart from one another to align with the one or more pins in the circumferential direction around the annular shroud.

5. The shield assembly according to claim 1, wherein: Wherein due to thermal expansion or contraction of the annular shroud, the block is capable of translating in the radial direction relative to the inner pin.

6. The shield assembly according to claim 1, wherein: The annular shroud comprises a ceramic matrix composite material.

7. The shield assembly according to claim 1, wherein: The annular shroud is disposed around one or more blades in a turbine section of a gas turbine engine.

8. The shield assembly according to claim 1, wherein: The annular shroud is disposed around one or more blades in a compressor section of a gas turbine engine.

9. The shield assembly according to claim 1, wherein: Wherein the domestic sales and the foreign sales are formed integrally.

10. The shield assembly according to claim 1, wherein: The annular shroud is capable of uniformly expanding and contracting in the radial direction.

11. A gas turbine engine defining an axial direction, a radial direction and a circumferential direction, characterized in that: The gas turbine engine comprises: a compressor section, a combustion section, and a turbine section, the compressor section, the combustion section, and the turbine section being in series flow relationship and collectively defining a core air flow path; and a shroud assembly positioned in at least one of the compressor section or the turbine section and at least partially defining the core air flow path, the shroud assembly comprising a shroud support and an annular shroud defining a radial centerline along the circumferential direction; The annular shroud includes one or more retaining features having a circumferential surface, the one or more retaining features being configured to engage one or more pins to secure the annular shroud to the shroud support, each of the one or more pins having an outer pin coupled to the shroud support and an inner pin disposed radially inward from the outer pin, the inner pin having a block thereon capable of radially translating along the inner pin, wherein the inner pin is offset from a radial centerline of the outer pin, wherein the block engages the circumferential surface of the one or more retaining features for securing the annular shroud to the shroud support.

12. The gas turbine engine according to claim 11, wherein The one or more retention features include a cavity formed on an outer surface of the annular shroud, the cavity having a first circumferential surface, a second circumferential surface, a front surface, and a rear surface, wherein the block engages one of the first circumferential surface or the second circumferential surface.

13. The gas turbine engine according to claim 11, wherein: Wherein the block includes a shroud engaging surface for engaging the circumferential surface of the one or more retention features.

14. The gas turbine engine according to claim 11, wherein: The annular shroud includes a plurality of retaining features spaced apart from one another to align with the one or more pins in the circumferential direction around the annular shroud.

15. The gas turbine engine according to claim 11, wherein The block is capable of translating in the radial direction relative to the inner pin due to thermal expansion or compression of the annular shroud.

16. The gas turbine engine according to claim 11, wherein: The annular shroud comprises a composite matrix material.

17. A method for assembling a shroud assembly structure in a gas turbine engine, characterized in that: The method comprises: providing at least a portion of a gas turbine engine having one or more shroud supports, the one or more shroud supports having a radially inner surface; providing an annular shroud clearance tool between the radially inner surface of the shroud support and an annular shroud to ensure a clearance between the radially inner surface and the annular shroud, the annular shroud having one or more shroud retention features; providing a block in the one or more shroud retention features, the block coupled to an inner pin coupled to an outer pin, the block being translatable in a radial direction relative to the inner pin, wherein the inner pin is offset from a radial centerline of the outer pin; rotating the outer pin or the inner pin to translate the block so that the block frictionally engages a circumferential surface on the one or more shroud retention features to couple the annular shroud to the block; Optionally, adjusting the torque of the pins to further secure the annular shroud to the block; Optionally, securing the outer pin to the shroud support; and Remove the annular shroud clearance tool.

18. The method according to claim 17, characterized in that The annular shroud is capable of uniformly expanding and contracting in the radial direction.

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