Abstract: A nozzle includes an endwall connected to an airfoil at one of a tip and root of the airfoil. The endwall includes a first and second side slash faces at respective circumferential edges thereof. The endwall also includes an inner surface extending between the first and second side slash faces. The inner surface includes a first planar surface portion adjacent to the first side slash face, a second planar surface portion adjacent to the second side slash face, and an arcuate surface portion extending between the first and second planar surface portions. The planar surface portions reduce a chute height by reducing hot gas path curvature locally along the side slash faces, which reduces working fluid ingestion between endwalls of the adjacent nozzles. The planar surface portions can be applied to an inner and/or an outer endwall of a nozzle.

Abstract: A nozzle for a turbine system includes an airfoil, an inner sidewall, and an outer sidewall. Each of the inner sidewall and outer sidewall includes a peripheral edge defining a pressure side slash face, a suction side slash face, a leading edge face, and a trailing edge face. At least one of the inner sidewall pressure side slash face, the inner sidewall suction side slash face, the outer sidewall pressure side slash face, or the outer sidewall suction side slash face includes a first swept surface extending at a first angle relative to a nominal slash face angle and a second swept surface extending at a second angle relative to the nominal slash face angle. The first and second swept surfaces meet at an arc having a peak that is circumferentially aligned with a stiffening member extending circumferentially on a respective sidewall.

Abstract: A hollow CMC article, a mandrel for forming the article and a method for forming the article are disclosed. The article includes a ply-wrap layer defining a cavity. The ply-wrap layer includes a first face, a second face, a root portion bridging the faces, and a plurality of CMC wrap plies. The root portion defines a terminus of the ply-wrap layer including a cross-sectional conformation consisting of a curve having a single turning point. Each of the plurality of CMC wrap plies are disposed along the first face, wrap over the root portion, and extend along the second face. The hollow article further includes a plurality of CMC lateral plies disposed along the faces.

Abstract: A nozzle for a turbine system includes an airfoil, an inner sidewall, and an outer sidewall. Each of the inner sidewall and outer sidewall each includes a peripheral edge defining a pressure side slash face, a suction side slash face, a leading edge face, and a trailing edge face. At least one of the inner sidewall pressure side slash face, the inner sidewall suction side slash face, the outer sidewall pressure side slash face, or the outer sidewall suction side slash face includes a first swept surface extending at a first angle relative to a nominal slash face angle and a second swept surface extending at a second angle relative to the nominal slash face angle. The first and second swept surfaces meet at a joining line that is circumferentially aligned with a stiffening member extending circumferentially on a respective sidewall.

Abstract: A component includes an outer wall that includes an exterior surface, and at least one plenum defined interiorly to the outer wall and configured to receive a cooling fluid therein. The component also includes a coating system disposed on the exterior surface. The coating system has a thickness. The component further includes a plurality of adaptive cooling openings defined in the outer wall. Each of the adaptive cooling openings extends from a first end inflow communication with the at least one plenum, outward through the exterior surface and to a second end covered underneath at least a portion of the thickness of the coating system.

Abstract: A turbine nozzle for a gas turbine includes: an airfoil having an airfoil shape having a nominal profile substantially in accordance with at least a portion of Cartesian coordinate values of X, Y and Z set forth in TABLE I. The Cartesian coordinate values are non-dimensional values of from 0% to 100% convertible to distances by multiplying the values by a height of the airfoil expressed in units of distance. The X and Y values are connected by smooth continuing arcs that define airfoil profile sections at each distance Z along at least a portion of the airfoil, and the airfoil profile sections at the Z distances are joined smoothly with one another to form the nominal profile. A trailing edge profile for a turbine nozzle is also disclosed.

Abstract: A method of protecting a gas turbine component for operation in a high temperature environment that includes the gas turbine component including a substrate having a silicon-containing layer, wherein the gas turbine component has a curved surface; forming a flexible mask configured to cover the curved surface of the gas turbine component, the flexible mask including a plurality of slots disposed in a pattern; disposing the flexible mask in direct contact with the curved surface of the gas turbine component; applying a bondcoat onto the flexible mask and the gas turbine component, such that bondcoat fills the plurality of slots and contacts the curved surface; and removing the flexible mask by heat or chemical reaction, such that, after removing the flexible mask, the curved surface of the gas turbine component comprises a patterned bondcoat layer in the pattern defined by the flexible mask.

Abstract: A turbine nozzle for a gas turbine includes: an airfoil having an airfoil shape having a nominal profile substantially in accordance with at least a portion of Cartesian coordinate values of X, Y and Z set forth in TABLE I. The Cartesian coordinate values are non-dimensional values of from 0% to 100% convertible to distances by multiplying the values by a height of the airfoil expressed in units of distance. The X and Y values are connected by smooth continuing arcs that define airfoil profile sections at each distance Z along at least a portion of the airfoil, and the airfoil profile sections at the Z distances are joined smoothly with one another to form the nominal profile. A trailing edge profile for a turbine nozzle is also disclosed.

Abstract: A nozzle for a turbine system includes an airfoil, an inner sidewall, and an outer sidewall. Each of the inner sidewall and outer sidewall includes a peripheral edge defining a pressure side slash face, a suction side slash face, a leading edge face, and a trailing edge face. At least one of the inner sidewall pressure side slash face, the inner sidewall suction side slash face, the outer sidewall pressure side slash face, or the outer sidewall suction side slash face includes a first swept surface extending at a first angle relative to a nominal slash face angle and a second swept surface extending at a second angle relative to the nominal slash face angle. The first and second swept surfaces meet at an arc having a peak that is circumferentially aligned with a stiffening member extending circumferentially on a respective sidewall.

Abstract: A nozzle for a turbine system includes an airfoil, an inner sidewall, and an outer sidewall. Each of the inner sidewall and outer sidewall each includes a peripheral edge defining a pressure side slash face, a suction side slash face, a leading edge face, and a trailing edge face. At least one of the inner sidewall pressure side slash face, the inner sidewall suction side slash face, the outer sidewall pressure side slash face, or the outer sidewall suction side slash face includes a first swept surface extending at a first angle relative to a nominal slash face angle and a second swept surface extending at a second angle relative to the nominal slash face angle. The first and second swept surfaces meet at a joining line that is circumferentially aligned with a stiffening member extending circumferentially on a respective sidewall.

Abstract: A component includes an outer wall that includes an exterior surface, and at least one plenum defined interiorly to the outer wall and configured to receive a cooling fluid therein. The component also includes a coating system disposed on the exterior surface. The coating system has a thickness. The component further includes a plurality of adaptive cooling openings defined in the outer wall. Each of the adaptive cooling openings extends from a first end inflow communication with the at least one plenum, outward through the exterior surface and to a second end covered underneath at least a portion of the thickness of the coating system.

Abstract: A method of forming a cooling assembly in a turbomachine part is provided. The method includes placing an encapsulated diffuser insert partially into a hole in the turbomachine part. The encapsulated diffuser insert has an unobstructed central passageway with a generally circular cross-section at a first end and an elongated rectangular cross-section at a second end opposing the first end. The second end has a sacrificial cap. A coating step coats the turbomachine part to at least partially encapsulate the encapsulated diffuser insert in a coating. A removing step removes the sacrificial cap to enable air flow through the central passageway. The encapsulated diffuser insert remains in the hole of the turbomachine part and the coating, thereby providing the unobstructed central passageway with a generally circular first end and an elongated rectangular second end adjacent to an outer surface of the turbomachine part.

Abstract: A ceramic matrix composite (CMC) turbine blade assembly includes a rotor, a CMC turbine blade, and at least one dovetail sleeve. The rotor has a blade slot with at least one slot surface. The slot surface is at a slot angle. The CMC turbine blade is received in the blade slot. The CMC turbine blade includes a dovetail root having at least one root surface. The root surface is at a root angle. The root angle is at least 5 degrees greater than the slot angle. The dovetail sleeve is received in the blade slot of the rotor. The dovetail sleeve has at least one inner surface contacting at least one root surface and at least one outer surface contacting at least one slot surface to radially retain the CMC turbine blade in the blade slot. A dovetail sleeve and a method of mounting a CMC turbine blade are also disclosed.

Abstract: A ceramic matrix composite (CMC) turbine blade includes an airfoil, a hub extending from the airfoil, and a shank extending from the hub. The airfoil includes a leading edge, a trailing edge, a pressure side, and a suction side. The shank includes a dovetail root having a dovetail path curved in a radial plane. In some embodiments, a leading shank length of the shank at the leading edge and a trailing shank length of the shank at the trailing edge are greater than an intermediate shank length at an intermediate location between the leading edge and the trailing edge. At least one of the airfoil, the hub, and the shank is formed from a CMC. A method of forming the CMC turbine blade includes forming the dovetail root to have a dovetail path curved in a radial plane.

Abstract: A ceramic matrix composite (CMC) turbine blade includes a root region, a narrowed neck region extending from the root region, and a hub region extending from the narrowed neck region. At least one central CMC ply extending from the root region is interwoven with at least one insert extending toward the narrowed neck region from the hub region. A method of forming a CMC turbine blade includes interweaving at least one central CMC ply extending from the root region with at least one insert extending toward the narrowed neck region from the hub region. The CMC turbine blade includes a root region, a narrowed neck region extending from the root region, and a hub region extending from the narrowed neck region.

Abstract: A preform CMC article is disclosed comprising an interior ply structure having at least one interior CMC ply including at least one longitudinal CMC ply disposed along the article length, an exterior shell ply forming an article surface and having at least one exterior CMC ply substantially surrounding the interior ply structure, and at least one wicking portion in which the interior ply structure penetrates the exterior ply shell with an exposed edge of the longitudinal CMC ply disposed at the article surface. A CMC article is disclosed including the interior ply structure and the exterior ply shell, wherein the longitudinal CMC ply includes an exposed edge disposed at the surface of the CMC article. A method for forming the CMC article is disclosed including wicking a melt infiltration agent into the article through the wicking portion into the interior ply structure along the longitudinal CMC ply.

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: A ceramic matrix composite (CMC) hollow blade includes a CMC airfoil, which includes at least one airfoil CMC ply, at least one cavity CMC ply, and an insert. The airfoil CMC ply defines the contour of the CMC airfoil including a first edge, a second edge opposite the first edge, a first side extending from the first edge to the second edge, and a second side opposite the first side. The cavity CMC ply defines a cavity within the CMC airfoil. The insert is located between the first edge and the cavity. The insert is wrapped by a CMC ply such that the CMC ply extends along the insert from the first side of the CMC airfoil across the mean camber line of the CMC airfoil and to the second side of the CMC airfoil. The CMC ply terminates on the second side of the CMC airfoil.

Abstract: A method of forming a cooling assembly in a turbomachine part is provided. The method includes placing an encapsulated diffuser insert partially into a hole in the turbomachine part. The encapsulated diffuser insert has an unobstructed central passageway with a generally circular cross-section at a first end and an elongated rectangular cross-section at a second end opposing the first end. The second end has a sacrificial cap. A coating step coats the turbomachine part to at least partially encapsulate the encapsulated diffuser insert in a coating. A removing step removes the sacrificial cap to enable air flow through the central passageway. The encapsulated diffuser insert remains in the hole of the turbomachine part and the coating, thereby providing the unobstructed central passageway with a generally circular first end and an elongated rectangular second end adjacent to an outer surface of the turbomachine part.

Abstract: A hollow CMC article, a mandrel for forming the article and a method for forming the article are disclosed. The article includes a ply-wrap layer defining a cavity. The ply-wrap layer includes a first face, a second face, a root portion bridging the faces, and a plurality of CMC wrap plies. The root portion defines a terminus of the ply-wrap layer including a cross-sectional conformation consisting of a curve having a single turning point. Each of the plurality of CMC wrap plies are disposed along the first face, wrap over the root portion, and extend along the second face. The hollow article further includes a plurality of CMC lateral plies disposed along the faces.