Abstract: A shroud assembly of a turbine system is presented. The shroud assembly includes a plurality of outer shroud segments disposed annularly and defining a first gap between a pair of outer shroud segments of the plurality of outer shroud segments; a plurality of inner shroud segments disposed annularly and radially inward of the plurality of outer shroud segments, and defining a second gap between a pair of inner shroud segments of the plurality of inner shroud segments, wherein the plurality of inner shroud segments is same in number to the plurality of outer shroud segments, and wherein the first gap and the second gap are offset; and an impingement plate coupled to an outer shroud segment of the pair of outer shroud segments and interposed between the outer shroud segment and the pair of inner shroud segments.

Abstract: The present disclosure relates to a turbomachine that includes a first component and a second component coupled to another and a seal assembly between the two components. The seal assembly may include a first intersegment seal, a second intersegment seal, and a rope seal that are configured to block a flow of gas between the two components.

Abstract: A system includes a web plate axially disposed between a transition piece and a turbine nozzle. The web plate includes a radial arm extending in a radial direction between an inner surface and an outer surface of the web plate. The radial arm, the inner surface, and the outer surface are disposed about an axial passage configured to facilitate a flow of combustion products from the transition piece to the turbine nozzle. The transition piece is disposed within a compressor discharge cavity configured to receive an oxidant. The radial arm includes an upstream face in fluid communication with the compressor discharge cavity. The radial arm also includes an arm passage that extends an axial length in an axial direction from the upstream face through at least an axial depth of the radial arm. The arm passage is configured to receive a portion of the oxidant through the upstream face.

Abstract: A system includes a transition piece configured to receive combustion products from a combustion chamber, and the transition piece is coupled to a first support of the system. The system further includes a turbine nozzle configured to receive combustion products from the transition piece, and the turbine nozzle is coupled to a second support of the system. The system further includes a web plate disposed at a radial position about a turbine axis between the transition piece and the turbine nozzle. The web plate is configured to form a first seal between the web plate and the transition piece and the web plate is configured to form a second seal between the web plate and the turbine nozzle. In addition, the web plate extends in a circumferential direction about the turbine axis and the web plate is coupled to a third support of the system. Further, the web plate includes an inner surface, an outer surface, and an arm extending in a radial direction between the inner surface and the outer surface.

Abstract: The present application provides seal assemblies having improved flexibility for reducing leakages between adjacent misaligned components of turbomachinery. The seal assemblies include a first outer shim formed of a flexible permeable material and a second outer shim formed of a substantially impervious material. At least the second outer shim is configured for sealing engagement with seal slots of the adjacent components. The seal assemblies may also include at least one of an inner shim and a filler layer positioned between the first and second outer shims. The seal assemblies may be sufficiently flexible to account for misalignment between the adjacent components, sufficiently stiff to meet assembly requirements, and sufficiently robust to meet operating requirements associated with turbomachinery. A turbomachine including the seal assembly is provided.

Abstract: Various embodiments include gas turbine seals and methods of forming such seals. In some cases, a turbine includes: a first arcuate component adjacent to a second arcuate component, each arcuate component including one or more slots having a seal assembly disposed therein. The seal assembly including an intersegment seal including a plurality of seal segments defining one or more end regions. One or more of the plurality of seal segments including at the one or more end regions a plurality of jet holes and a channel having a wire disposed therein, wherein the intersegment seal provides sealing of one or more end gaps defined proximate the one or more end regions in response to the thrust of a flow of pressurized air through the plurality of jet holes.

Abstract: An airfoil profile-shaped seal is disposable between an inside surface of a hollow airfoil body and a seal support. The seal includes an elongated member defining a generally closed curve having a generally straight portion, a generally curved portion, and having a solid cross-section across and along the length of the elongated member. When the seal is disposed in an airfoil profile-shaped groove of a seal support adjacent to the inner surface of the hollow airfoil body, and a fluid pressure on one side of the seal is greater than a fluid pressure on the other side of the seal, portions of the seal engage the inner surface of the hollow airfoil body and a surface of the groove to form a seal to inhibit leakage of fluid from the one side to the other side.

Abstract: Coating systems for components of a gas turbine engine, such as a compressor blade tip, are provided. The coating system can include an abradable material disposed along the compressor blade tip and may be used with a bare compressor casing. The abradable coating is softer than the compressor casing and can reduce the overall rub ratio thereby increasing the lifetime of the compressor blade and casing. Methods are also provided for applying the coating system onto a compressor blade.

Abstract: Coating systems for components of a gas turbine engine, such as a compressor casing, are provided. The coating system can include a ceramic material disposed along the compressor casing on a surface to be adjacent to a rotating compressor blade. The coating system is harder than the compressor blade and can reduce the rub ratio between the casing and blade. The coating system can thereby increase the lifetime of the compressor casing and blades. Methods are also provided for applying the coating system onto a compressor casing.

Abstract: A seal for a rotary machine includes a flexible element extending circumferentially about a rotor and extending generally radially from a first end to a free second end. The flexible element is coupled proximate the first end for rotation with the rotor. The flexible element extends at substantially a first angle between at least an intermediate portion of the flexible element and the free second end when the rotor is operating at less than a critical speed, such that a clearance gap is defined between the free second end and a stationary portion. The seal also includes a retaining plate having a stop that orients the flexible element at a second angle proximate the free second end when the rotor is operating at equal to or greater than the critical speed, such that the flexible element forms a dynamic seal between the rotor and the stationary portion.

Abstract: A turbine system including a sealing component is presented. The sealing component is positioned in a gap between adjacent turbine components of the turbine system. The sealing component includes a metallic shim including a high-temperature-resistant alloy in a single crystal form. A turbine shroud assembly including the sealing component is also presented.

Abstract: Coating systems for components of a gas turbine engine, such as a compressor blade tip, are provided. The coating system can include a ceramic material disposed along the compressor blade tip and may be used with a bare compressor casing. The ceramic coating is harder than the bare compressor casing and can reduce the rub ratio thereby increasing the lifetime of the compressor blades. Methods are also provided for applying the coating system onto a compressor blade.

Abstract: A seal for a rotary machine includes a flexible element extending circumferentially about a rotating component and extending generally radially from a first end to a free second end. The flexible element is coupled proximate the first end for rotation with the rotating component. The seal also includes a retaining plate having a stop face that orients the flexible element at a first angle when the rotating component is operating at less than a critical speed and under the including of pressure loading, such that a clearance gap is defined between the free second end and a stationary component. The flexible element extends at substantially a second angle when the rotating component is operating at a speed greater than the critical speed, such that the flexible element forms a dynamic seal between the rotating component and the stationary component.

Abstract: Various embodiments include gas turbine seals and methods of forming such seals. In some cases, a turbine includes: a first arcuate component adjacent to a second arcuate component, each arcuate component including one or more slots located in an end face, each of the one or more slots having a plurality of axial surfaces and radially facing surfaces extending from opposite ends of the axial surfaces and a seal assembly disposed in the slot. The seal assembly including an intersegment seal including a plurality of seal segments defining one or more corner regions. The intersegment seal disposed in a slot defining a high-pressure slot side and a low-pressure slot side, wherein the slot includes a plurality of slot segments. Each of the plurality of seal segments including at least one extended portion to either seal the one or more corner regions on a low-pressure slot side, form a recess in the intersegment seal, or both.

Abstract: The present application provides a seal for use between adjacent turbine components and with a cooling flow. The seal may include an impingement baffle top plate, a base plate, and one or more spacer elements therebetween. The cooling flow provides cooling through the impingement baffle top plate.

Abstract: A turbine system including a sealing component is presented. The sealing component includes a ceramic material. The ceramic material includes grains having an average grain size of less than 10 microns. A turbine shroud assembly including the sealing component is also presented.

Abstract: Seal assemblies for reducing leakage between components of turbomachinery include a metallic shim, at least a pair of non-metallic end blocks, and ceramic fiber positioned between the shim and the end blocks. The shim may be mechanically coupled with the end blocks such that the metallic shim, end blocks and ceramic fiber are coupled. The end blocks account for misalignment of turbine components by ensuring sealing engagement of the seal to the components. The end blocks may be a ceramic or glass material, and the ceramic fiber may be a high temperature woven ceramic fiber. The ceramic fiber and/or the end blocks protect the metallic shim from reaching harmful temperatures during use of the seal, such as use in high temperature turbines including CMC components.

Abstract: Various embodiments include gas turbine seals and methods of forming such seals. In some cases, a turbine includes: a first arcuate component adjacent to a second arcuate component, each arcuate component including one or more slots located in an end face, each of the one or more slots having a plurality of axial surfaces and radially facing surfaces extending from opposite ends of the axial surfaces and a seal assembly disposed in the slot. The seal assembly including an intersegment seal including a plurality of seal segments defining one or more corner regions. The intersegment seal disposed in a slot defining a high-pressure slot side and a low-pressure slot side, wherein the slot includes a plurality of slot segments. Each of the plurality of seal segments including at least one extended portion to either seal the one or more corner regions on a low-pressure slot side, form a recess in the intersegment seal, or both.

Abstract: Various embodiments include gas turbine seals and methods of forming such seals. In some cases, a turbine includes: a first arcuate component adjacent to a second arcuate component, each arcuate component including one or more slots located in an end face, each of the one or more slots having a plurality of axial surfaces and radially facing surfaces extending from opposite ends of the axial surfaces and a seal assembly disposed in the slot. The seal assembly including a backup intersegment seal disposed in the slot on a high-pressure side of the slot and a shim seal disposed in the slot over the backup intersegment seal and covering the backup intersegment seal on a low-pressure side of the slot. The shim seal including one or more shim seal segments and having a gap formed between each of the one or more shim seal segments.

Abstract: A method and system for sealing between components within a gas turbine is provided. A first recess defined in a first component receives a seal member. A second recess defined in a second component adjacent the first component also receives the seal member. The first and second recesses are located proximate a hot gas path defined through the gas turbine, and define circumferential paths about the turbine axis. The seal member includes a sealing face that extends in a direction substantially parallel to the turbine axis. The seal member also includes a plurality of seal layers, wherein at least one of the seal layers includes at least one stress relief region for facilitating flexing of the first seal member.