Abstract: An airfoil stage of a turbine engine includes an upstream airfoil assembly, a downstream airfoil assembly in rotational relationship to the upstream airfoil assembly and a rim seal assembly integrated therebetween. The rim seal assembly may include a sloped downstream portion of a platform of the upstream airfoil assembly, an upstream segment of a platform of the downstream airfoil assembly and a nub that projects radially outward from the upstream segment. The downstream portion and the upstream segment are spaced from one-another defining a cooling cavity therebetween for the flow of cooling air. The portion and segment overlap axially such that the nub is axially aligned to the downstream portion for improved cooling effectiveness and a reduction of core airflow into the cooling cavity.

Abstract: A stator vane rail is provided. The stator vane rail may comprise a forward rail, an aft rail located axially opposite the forward rail, a first axial surface extending between the forward rail and the aft rail, a first feather seal slot disposed in the first axial surface, wherein a terminus of the first feather seal slot is radially spaced from a radial surface of the aft rail, a second axial surface extending between the forward rail and the aft rail, and a second feather seal slot disposed in the second axial surface, wherein the second feather seal slot extends from the radial surface of the aft rail.

Abstract: A cooled component for a gas turbine engine includes a plurality of internal ribs extending substantially parallel to a longitudinal axis of the gas turbine engine. The internal ribs are disposed within an internal cavity defining cooling air passages within the cooled component. A plurality of cooling holes are arranged in rows with axial orientations alternating between a radially outboard bias directing cooling air radially outward and a radially inboard bias directing cooling air radially inward. Each of the cooling holes includes an internal opening in communication with one of the cooling air passages and an external opening open to an outer surface of the cooled component. The external opening of each of the plurality of cooling holes is disposed on a side opposite the internal rib relative to a corresponding internal opening. A gas turbine engine and a method of fabricating a turbine airfoil are also disclosed.

Abstract: A flow path component includes a platform having at least one radially aligned face. A chordal seal extends axially from the radially aligned face. The chordal seal includes a first curved face configured to prevent edge line contact under deflection conditions while the flow path component is installed in an engine.

Abstract: An article includes a body that has a coating thereon. The coating has a first portion disposed on a first section of the body and a second portion disposed on a second, different section of the body. The first portion has a first microstructure and the second portion has a second, different microstructure.

Abstract: A vane includes a pair of airfoils that have a plurality of film cooling holes that extend through an exterior surface of the airfoils. Each plurality of film cooling holes break through the exterior surface at geometric coordinates in accordance with Cartesian coordinate values of X, Y and Z as set forth in Table 1. Each geometric coordinates is measured from a reference point on a leading edge rail of a platform of the vane.

Abstract: A section of a gas turbine engine according to an exemplary aspect of this disclosure includes, among other things, a vane assembly including a featherseal slot. The section further includes a featherseal at least partially received in the featherseal slot. The featherseal includes a radial portion having at least one tapered side.

Abstract: A component for a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, an body portion that extends between a leading edge and a trailing edge. At least one of the leading edge and the trailing edge includes at least one discharge slot having a first portion that includes an oval geometry.

Abstract: A component for a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a mate face and a feather seal slot axially extending along the mate face, the feather seal slot having a variable width along a portion of its axial length.

Abstract: An airfoil includes an airfoil wall including an exterior airfoil surface and at least partially defines an airfoil cavity. A fillet is on the exterior airfoil surface. A recess is in an interior surface of the airfoil wall adjacent the fillet. A baffle tube is located in the airfoil cavity spaced from the recess.

Abstract: An airfoil platform includes a platform body having an arcuate base defining a centerline axis. A rail extends radially outboard from a side of the arcuate base. A feed orifice is defined at least partially in the rail. The inner diameter edge of an axially facing opening of the feed orifice is spaced apart radially outboard of an outer diameter surface of the arcuate base. A method for manufacturing an airfoil platform includes providing an airfoil platform body and forming the feed orifice through the rail. Forming the feed orifice through the rail includes forming the feed orifice radially outboard of and spaced apart from the outer diameter surface of the arcuate base.

Abstract: A vane includes a pair of airfoils that have a plurality of film cooling holes that extend through an exterior surface of the airfoils. Each plurality of film cooling holes break through the exterior surface at geometric coordinates in accordance with Cartesian coordinate values of X, Y and Z as set forth in Table 1. Each geometric coordinates is measured from a reference point on a leading edge rail of a platform of the vane.

Abstract: A stator vane rail is provided. The stator vane rail may comprise a forward rail, an aft rail located axially opposite the forward rail, a first axial surface extending between the forward rail and the aft rail, a first feather seal slot disposed in the first axial surface, wherein a terminus of the first feather seal slot is radially spaced from a radial surface of the aft rail, a second axial surface extending between the forward rail and the aft rail, and a second feather seal slot disposed in the second axial surface, wherein the second feather seal slot extends from the radial surface of the aft rail.

Abstract: A flow path component includes a platform having at least one radially aligned face. A chordal seal extends axially from the radially aligned face. The chordal seal includes a first curved face configured to prevent edge line contact under deflection conditions while the flow path component is installed in an engine.

Abstract: A cooled component for a gas turbine engine includes a plurality of internal ribs extending substantially parallel to a longitudinal axis of the gas turbine engine. The internal ribs are disposed within an internal cavity defining cooling air passages within the cooled component. A plurality of cooling holes are arranged in rows with axial orientations alternating between a radially outboard bias directing cooling air radially outward and a radially inboard bias directing cooling air radially inward. Each of the cooling holes includes an internal opening in communication with one of the cooling air passages and an external opening open to an outer surface of the cooled component. The external opening of each of the plurality of cooling holes is disposed on a side opposite the internal rib relative to a corresponding internal opening. A gas turbine engine and a method of fabricating a turbine airfoil are also disclosed.

Abstract: An airfoil for a gas turbine engine includes a first airfoil. A first chordal seal is located adjacent a first end of the airfoil. A second chordal seal is located adjacent a second end of the airfoil. The first chordal seal includes a first edge parallel to a first edge on the second chordal seal.

Abstract: An airfoil of a turbine engine may have internal passages to permit the travel of cooling air through the airfoil. Multiple airfoils may be formed together in a stator vane assembly, and cooling air may be directed through the stator vane assembly. The stator vane assembly may be mounted to an engine structure to carry structural loads experienced by the stator vane assembly. Buttresses may be formed in the stator vane assembly, such as beneath each airfoil of the stator vane assembly, and extending into a passage permitting the travel of cooling air through the airfoil. The buttresses may enhance the strength and stability of the stator vane assembly by facilitating the transmission of structural loads to an engine structure. A channel may be defined through one or more buttress to enable the passage of cooling air into the airfoils, and yet allow the buttresses to carry structural loads.

Abstract: A gas turbine engine component comprising a nozzle segment, the nozzle segment comprising at least one substrate having a surface. A metallic bondcoat is coupled to the surface of the substrate. An yttria-stabilized zirconia thermal barrier coating is coupled to the metallic bondcoat opposite the surface.

Abstract: An airfoil for a gas turbine engine includes a first airfoil. A first chordal seal is located adjacent a first end of the airfoil. A second chordal seal is located adjacent a second end of the airfoil. The first chordal seal includes a first edge parallel to a first edge on the second chordal seal.

Abstract: An airfoil for a gas turbine engine includes a first platform located at a first end of a first airfoil. A cooling passage extends through the first platform and includes a first portion that has a first thickness and a second portion that has a second thickness and surrounds opposing ends of the first portion.