Abstract: The present application and the resultant patent provide improved gas turbine component sealing. In one example embodiment, a gas turbine segment seal assembly may include a first tapered segment seal with a first tapered portion having a first tapered surface and a first taper angle. The gas turbine segment seal assembly may include a second tapered segment seal with a second tapered portion having a second tapered surface and a second taper angle. The gas turbine segment seal assembly may include a seal pin positioned in between the first tapered segment seal and the second segment seal and adjacent to the first tapered surface and the second tapered surface.

Abstract: A sealing arrangement for sealing between a first stage nozzle and a plurality of aft frames includes a first inner seal and a second inner seal which are circumferentially oriented and circumferentially aligned. A side seal is radially disposed between the first inner seal and the second inner seal. The side seal includes a first portion that is axially offset from the first inner seal and the second inner seal and a second portion abutting the first inner seal and the second inner seal. Also disclosed is an aft frame having a forward face and an aft face. The aft frame includes at least one side seal slot that extends along a side portion. A first portion of the side seal slot is axially offset from the aft face. A second portion of the side seal slot is axially and radially offset from the first portion.

Abstract: A turbine shroud assembly includes a plurality of arcuate shroud block assemblies annularly arranged to form a shroud segment. The plurality of shroud block assemblies includes a first shroud block assembly having a shroud block and a second shroud block assembly having a shroud block. The first shroud block assembly includes a seal interface member and a shroud seal. The seal interface member has a side portion that is adjacent to a radial side surface of the first shroud block. The second shroud block assembly includes a seal interface member and a shroud seal. The seal interface member has a side portion that is adjacent to a radial side surface of the second shroud block.

Abstract: A turbine component for use in a hot gas path of a gas turbine engine. The turbine component may include an outer surface, an internal cooling circuit, an adaptive cooling pathway in communication with the internal cooling circuit and extending through the outer surface, and a cooling plug having two or more materials positioned within the adaptive cooling pathway.

Abstract: Various embodiments include gas turbine seals and methods of forming such seals. In some cases, a turbine includes: a first arc segment adjacent to a second arc segment, each arc segment including an end surface and radially facing surfaces extending from opposite ends of the end surface; a slot located between the end surfaces of the first arc segment and the second arc segment; and a first seal disposed in the slot, the first seal contacting the first arc segment at the end surface and extending over the radially facing surfaces of the first arc segment, the first seal including: a shim contacting the first arc segment; a laminate material over the shim and covering the shim; and a conforming material coupling the laminate material to the shim.

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.

Abstract: A process of producing a ceramic matrix composite gas turbine component and a ceramic matrix composite gas turbine component are provided. The process includes modifying a surface of the ceramic matrix composite gas turbine component to produce a modified surface with a surface roughness of less than 6 micrometers. The modifying is selected from the group of techniques consisting of applying unreinforced matrix plies to the surface, vapor depositing silicon on the surface, honing the surface, applying braze paste to the surface, and combinations thereof. The component includes a modified surface including a surface roughness of less than 6 micrometers. The modified surface being selected from the group consisting of unreinforced matrix plies applied to a surface of the ceramic matrix composite gas turbine component, silicon vapor deposited on the surface, a honed surface, a braze paste applied to the surface, and combinations thereof.

Abstract: Components are disclosed which include a CMC substrate having a first surface and a second surface. The first surface is in fluid communication with a compressed, dry fluid, and the second surface is in fluid communication with a wet fluid stream and includes a hermetic coating. The components further include at least one opening extending from the first surface through a portion of the CMC substrate, wherein, upon removal of a fragment of one or both of the hermetic coating and the CMC substrate, the at least one opening selectively permits a flow of the compressed, dry fluid to the second surface. In one embodiment, the component is a gas turbine component, the wet fluid stream is a hot combustion stream, the hermetic coating is an environmental barrier coating, and the flow reduces or eliminates volatilization of the CMC substrate. Methods for forming the components are also disclosed.

Abstract: Various embodiments include gas turbine seals and methods of forming such seals. In some cases, a turbine includes: a first arc segment adjacent to a second arc segment, each arc segment including an end surface and radially facing surfaces extending from opposite ends of the end surface; a slot located between the end surfaces of the first arc segment and the second arc segment; and a first seal disposed in the slot, the first seal contacting the first arc segment at the end surface and extending over the radially facing surfaces of the first arc segment, the first seal including: a shim contacting the first arc segment; a laminate material over the shim and covering the shim; and a conforming material coupling the laminate material to the shim.

Abstract: A process of producing a ceramic matrix composite component. The process includes positioning a plurality of ceramic matrix composite plies on top of one another and forming a cavity therein. At least a portion of the cavity includes a terminal diameter sufficiently small to permit infiltration of a densifying material. The plurality of ceramic matrix composite plies are densified to form a densified body. The densifying results in the portion of the cavity including the terminal diameter being filled with densifying material and the cavity is present in the densified body. A ceramic matrix composite having cavities therein is also disclosed.

Abstract: A process of producing a hot gas path turbine component. The process includes forming a void in a first ceramic matrix composite ply and forming a void in a second ceramic matrix composite ply. The second ceramic matrix composite ply is positioned on the first ceramic matrix composite ply such that the positioning aligns the voids to at least partially define a cavity in the component. A third ceramic matrix composite ply is positioned on the first ceramic matrix composite ply and the first ceramic matrix composite ply, the second ceramic matrix composite ply and the third ceramic matrix composite ply are densified to form a densified body. The cavity is present in the densified body. A ceramic matrix composite having cavities therein is also disclosed.

Abstract: The present application provides a turbine component for use in a hot gas path of a gas turbine. The turbine component may include an outer surface, an internal cooling circuit, a number of cooling pathways in communication with the internal cooling circuit and extending through the outer surface, and a number of adaptive cooling pathways in communication with the internal cooling circuit and extending through the outer surface. The adaptive cooling pathways may include a high temperature compound therein.

Abstract: The present application provide a seal for use between components engine facing a high pressure cooling air flow and a hot gas path in a gas turbine. The seal may include a first shim, a second shim with an air exit hole, one or more middle layers positioned between the first shim and the second shim, and one or more cooling pathways extending through the middle layers for the high pressure cooling air flow to pass therethrough and exit via the air exit hole into the hot gas path.

Abstract: A turbine shroud assembly is disclosed including an inner shroud having a surface adjacent to a hot gas path, an outer shroud, a damper block disposed between the inner shroud and the outer shroud, a first biasing apparatus, and a second biasing apparatus. The first biasing apparatus provides a first biasing force to the inner shroud, biasing the inner shroud a first deflection distance in a direction toward the hot gas path and away from the outer shroud. The second biasing apparatus provides a second biasing force to the damper block, biasing the damper block a second deflection distance in a direction toward the hot gas path and away from the outer shroud. The second deflection distance is greater than the first deflection distance.

Abstract: A turbine shroud assembly is disclosed including an inner shroud having a surface adjacent to a hot gas path, an outer shroud, and a biasing apparatus. The biasing apparatus is arranged and disposed to bias the inner shroud in a direction away from the hot gas path, loading the inner shroud to the outer shroud. In another embodiment, the biasing apparatus is a springless biasing apparatus including at least one bellows, at least one thrust piston, or a combination of at least one bellows and at least one thrust piston. A method for loading the turbine shroud assembly is disclosed including biasing the inner shroud having a surface adjacent to a hot gas path in a direction away from the hot gas path toward the outer shroud, wherein biasing the inner shroud includes a biasing force exerted by the biasing apparatus.

Abstract: Embodiments of the present disclosure provide components for hot gas path (HGP) components and methods of forming the same. A structure according to the present disclosure can include: an HGP component extending radially from a rotor axis of a turbomachine, the HGP component including a tapered edge; a plurality of first passages in fluid communication with a preliminary cooling zone of the HGP component, and extending through a sidewall positioned between the preliminary cooling zone and the tapered edge; and a plurality of second passages extending through at least the tapered edge, wherein each of the plurality of second passages is in fluid communication with the flow path for the operative fluid and at least one passage of the plurality of first passages, and wherein at least one of the plurality of second passages is radially displaced from each passage of the plurality of first passages.

Abstract: Exemplary embodiments include a multi-layer, modular and replaceable heat shield for gas turbines. The heat shield apparatus can include a base layer adjacent the airfoil and a thermal layer coupled to the base layer, wherein the base layer and the thermal layer match a contour of the airfoil.

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: The present application provide a seal for use between components engine facing a high pressure cooling air flow and a hot gas path in a gas turbine. The seal may include a first shim, a second shim with an air exit hole, one or more middle layers positioned between the first shim and the second shim, and one or more cooling pathways extending through the middle layers for the high pressure cooling air flow to pass therethrough and exit via the air exit hole into the hot gas path.

Abstract: A three-dimensional printing process, a swirling device, and a thermal management process are disclosed. The three-dimensional printing process includes distributing a material to a selected region, selectively laser melting the material, and forming a swirling device from the material. The swirling device is printed by selective laser melting. The thermal management process includes providing an article having a swirling device printed by selective laser melting, and cooling a portion of the article by transporting air through the swirling device.