The present disclosure is directed to a rotor blade for a gas turbine system. The rotor blade includes a platform having a radially inner surface and a radially outer surface. A shank portion extends radially inwardly from the radially inner surface of the platform. The shank portion and the platform collectively define a shank pocket. An airfoil extends radially outwardly from the radially outer surface of the platform. The shank portion, the platform, and the airfoil collectively define a cooling passage extending from a cooling passage inlet defined by the shank portion or the platform and directly coupled to the shank pocket through the platform to a cooling passage outlet defined by the airfoil.

A seal assembly includes first, second, and third gas turbine engine components, such as vane assemblies, that are successively arranged around an axis. Each component has first and second mate faces such that the first mate face of the first component and the second mate face of the second component define a first mate face gap, and the first mate face of the second component and the second mate face of the third component to define a second mate face gap. A seal arc segment has first and second seal portions and a connector portion that joins the seal portions. The first seal portion bridges the first mate face gap to seal the first mate face gap, the second seal portion bridges the second mate face gap to seal the second mate face gap, and the connector portion spans circumferentially across the second component.

A manufacturing method is provided. During this method, a preform component is provided for a turbine engine. The preform component includes a substrate comprising electrically conductive material having an outer coating comprising non-electrically conductive material applied over a surface of the substrate. A preform aperture is formed in the preform component using an electrical discharge machining electrode. The preform aperture includes a meter section of a cooling aperture in the substrate. The preform aperture also includes a pilot hole in the outer coating. A diffuser section of the cooling aperture is formed in at least the outer coating using a second machining process.

A cooling structure for a turbomachine. In one embodiment, the cooling structure is for a seal slot of the turbomachine. The cooling structure includes a body coupled to a surface of the seal slot. The body includes a passageway on a first surface of the body for providing a cooling fluid to the seal slot. In an other embodiment, a apparatus includes a first component and a second component adjacent the first component. The apparatus also includes a seal slot extending between the first component and the second component, and a cooling structure positioned within the seal slot. The cooling structure includes a body coupled to a surface of the seal slot. The body has a passageway on a first surface of the body for providing a cooling fluid to the seal slot.

A turbine rotor blade includes an airfoil, root, and platform that is between the root and a proximate end portion of the airfoil. The blade defines a passage having a first leg, second leg, and arcuate portion. The arcuate portion is at least partially within the platform and connects the first and second legs. The first leg extends between a distal end portion of the airfoil and an inlet of the arcuate portion. The second leg extends from an outlet of the arcuate portion to the distal end portion of the airfoil. The platform includes a first feed passage and branch passages. The first feed passage is open through an extrados of the arcuate portion and is in fluid communication with the branch passages. The inlet of each branch passage is connected with the first feed passage while the outlet is open to an exterior of the platform.

A turbine blade that allows an improvement in torque and power, and a turbine and gas turbine including the same are provided. The turbine blade includes an airfoil having a suction side and a pressure side, a platform coupled to a bottom of the airfoil, and a root protruding downward from the platform and coupled to a rotor disk, wherein the airfoil includes a cooling passage formed therein and a discharge hole connected to an upper portion of the cooling passage to discharge cooling air, and the discharge hole is inclined toward a tip of the turbine blade while extending from an inside to an outside thereof.

A turbine stator vane includes; an airfoil portion; a shroud disposed on at least one of the side of a tip end portion or the side of a root end portion of the airfoil portion; and a protruding portion protruding toward the opposite side to the airfoil portion across a gas path surface in a radial direction. The shroud includes: a circumferential-direction passage disposed at the side of a trailing edge and extending in a circumferential direction; and a plurality of trailing edge end portion passages arranged in the circumferential direction at the side of the trailing edge, each trailing edge end portion passage having a first end portion connected to the circumferential-direction passage and a second end portion having an opening on a trailing edge end surface of the shroud. The circumferential-direction passage includes, in a cross-sectional view as seen in the circumferential direction, an oblique passage having a third end portion to protrude closer to the gas path surface toward the side of a leading edge, and a fourth end portion formed at the side of the trailing edge with respect to the third end portion, the oblique passage having an opening portion disposed on a trailing-edge side end surface of the protruding portion and closed by a lid portion, and a position, in an axial direction, of the first end portion where each trailing edge end portion passage connects to the circumferential-direction passage is positioned at the side of the leading edge with respect to a position of the trailing-edge side end surface of the protruding portion at a position where the protruding portion connects to the shroud.

A method includes forming a ceramic matrix composite component by infiltrating an array of ceramic-based fibers with a ceramic-based matrix; forming a plurality of cooling holes in the ceramic matrix composite component; applying a slurry of particles in a carrier fluid to the ceramic matrix composite component such that the slurry passes through the cooling holes and wicks into the ceramic matrix composite material; and processing the ceramic matrix composite component to remove the carrier fluid, thereby leaving a filler at a wall surface of the plurality of cooling holes. A component is also disclosed.

The present technique presents a gas turbine blade for re-using cooling air, a turbomachine assembly having the blade, and a gas turbine having the turbomachine assembly. The blade includes a platform and an airfoil extending from the platform. The airfoil includes a pressure surface, a suction surface, a leading edge and a trailing edge. The platform includes a pressure side, a suction side, a leading-edge side and a trailing-edge side, disposed towards the pressure surface, the suction surface, the leading edge and the trailing edge of the airfoil, respectively. The suction side of the platform includes a part of the upper surface and a suction-side lateral surface of the platform. At least a part of an edge between the suction-side lateral surface and the upper surface of the platform comprises a chamfer part.

A gas turbine engine includes a ceramic wall for bounding an engine core gas path. The ceramic wall has a ceramic wall first side that faces the engine core gas path and a ceramic wall second side that faces away from the engine core gas path. There is a metallic wall adjacent the ceramic wall second side. The metallic wall has a metallic wall first side that faces the ceramic wall and a metallic wall second side that faces away from the ceramic wall. The metallic wall and the ceramic wall are spaced apart such that there is a channel there between. There is a seal on the ceramic wall second side, and the metallic wall has at least one cooling hole adjacent the seal for emitting cooling air to cool the seal.