Abstract: Aspects of the disclosure are directed to a seal configured to interface with at least a first component and a second component of a gas turbine engine. A method for forming the seal includes obtaining an ingot of a fine grained, or a coarse grained, or a columnar grained or a single crystal material from a precipitation hardened nickel base superalloy containing at least 40% by volume of the precipitate of the form Ni3(Al, X), where X is a metallic or refractory element, and processing the ingot to generate a sheet of the material, where the sheet has a thickness within a range of 0.010 inches and 0.050 inches inclusive.

Abstract: Gas turbine engines and turbines thereof including a stator section having a plurality of vanes, a rotating section having a plurality of blades, the rotating section being axially adjacent the stator section along an axis of the turbine, the stator section being aftward of the rotating section along the axis of the turbine, and a primary tangential onboard injector located radially inward from the stator section and configured to direct an airflow from the stator section in a forward direction toward the rotating section, the primary tangential onboard injector turning the airflow in a direction of rotation of the rotating section.

Abstract: A hollow turbine airfoil or a hollow turbine casting including a cooling passage partition. The cooling passage partition is formed from a single crystal grain structure nickel based super alloy, a cobalt based super alloy, a nickel-aluminum based alloy, or a coated refractory metal.

Abstract: An airfoil includes an airfoil section that defines an airfoil profile, and a first endwall section with which the airfoil section is attached. First and second airfoil pieces form different portions of the airfoil profile. The first and second airfoil pieces include respective first ends. The first ends interlock with the first endwall section such that the first and second airfoil pieces are retained with the first endwall section.

Abstract: A gas turbine engine with an adjustable vane includes a platform with a hole and an aperture. A vane is supported for rotation relative to the platform by a trunion that is received in the hole. The vane has an opening that is laterally spaced from the trunion and is in alignment with the aperture. The vane includes an airfoil with a cooling passage in fluid communication with the opening.

Abstract: A method for forming a passage in a ceramic matrix composite component includes forming a core for a ceramic matrix composite component; embedding a hollow member into the core at a desired location for a passage in the ceramic matrix composite component; wrapping the core with a ceramic material; and inserting a rod through the hollow member and into the core.

Abstract: An airfoil includes an airfoil section that defines an airfoil profile. The airfoil section includes a panel that forms a portion of the airfoil profile and a core structure to which the panel is secured.

Abstract: A vane platform cooling system may comprise a combustor shell and a combustor panel. A cavity may be located between the combustor shell and the combustor panel. A surface of the cavity may be angled toward the combustor shell. An orifice may be formed in a vane platform located aft of the combustor panel. A standoff located in the cavity may direct air toward the vane platform. An aperture may be formed in a surface of the vane platform. A channel formed through the vane platform may connect the orifice and aperture.

Abstract: A variable area turbine arrangement according to an exemplary aspect of the present disclosure includes, among other things, a first turbine section having at least a first variable vane row and a second turbine section downstream from the first turbine section and having at least a second variable vane row. A transition duct is disposed between the first turbine section and the second turbine section.

Abstract: An adjustable stator vane for a turbine engine includes a shaft, a flange and a stator vane body that pivots about a variable vane axis. The stator vane body extends axially between a first end and a second end. The stator vane body includes an airfoil, a cavity, and a body surface located at the first end. The cavity extends axially from an inlet in the body surface and into the airfoil. The shaft extends along the variable vane axis from the first end. The flange extends circumferentially around the inlet and the shaft, and radially from the stator vane body.

Abstract: A variable area turbine arrangement according to an exemplary aspect of the present disclosure includes, among other things, a variable vane assembly and a secondary flow system associated with the variable vane assembly. Flow modulation of a cooling fluid through the secondary flow system is changed simultaneously with actuation of the variable vane assembly.

Abstract: A variable-pitch vane assembly for a gas turbine engine includes a sync ring, a vane having a vane arm, and a pin installed through the sync ring and through the vane arm. The pin includes an anti-rotation notch located along a pin shaft. An anti-rotation spacer is engaged with the pin at the anti-rotation notch to prevent rotation of the pin. A turbine section of a gas turbine engine includes a turbine rotor and a turbine stator. The turbine stator includes one or more variable-pitch vane assemblies including a sync ring, a vane having a vane arm, and a pin installed through the sync ring and through the vane arm. The pin includes an anti-rotation notch located along a pin shaft. An anti-rotation spacer is engaged with the pin at the anti-rotation notch to prevent rotation of the pin.

Abstract: A variable area turbine arrangement according to an exemplary aspect of the present disclosure includes, among other things, a variable vane assembly and a secondary flow system associated with the variable vane assembly. Flow modulation of a cooling fluid through the secondary flow system is changed simultaneously with actuation of the variable vane assembly.

Abstract: Gas turbine engines and turbines thereof including a stator section having a plurality of vanes, a rotating section having a plurality of blades, the rotating section being axially adjacent the stator section along an axis of the turbine, the stator section being aftward of the rotating section along the axis of the turbine, and a primary tangential onboard injector located radially inward from the stator section and configured to direct an airflow from the stator section in a forward direction toward the rotating section, the primary tangential onboard injector turning the airflow in a direction of rotation of the rotating section.

Abstract: A gas turbine engine according to an example of the present disclosure include a compressor section, a combustor, and a turbine section. The combustor has a radially outer surface that defines a diffuser chamber radially outwardly of the combustor. The turbine section has a high pressure turbine first stage blade that has an outer tip, and a blade outer air seal positioned radially outwardly of the outer tip. A tap for tapping air has been compressed by the compressor and is passed through a heat exchanger. The air downstream of the heat exchanger passes through at least one pipe and into a manifold radially outward of the blade outer air seal, and then passes across the blade outer air seal to cool the blade outer air seal.

Abstract: Aspects of the disclosure are directed to a seal configured to interface with at least a first component and a second component of a gas turbine engine. A method for forming the seal includes obtaining an ingot of a fine grained, or a coarse grained, or a columnar grained or a single crystal material from a precipitation hardened nickel base superalloy containing at least 40% by volume of the precipitate of the form Ni3(Al, X), where X is a metallic or refractory element, and processing the ingot to generate a sheet of the material, where the sheet has a thickness within a range of 0.010 inches and 0.050 inches inclusive.

Abstract: An assembly for a turbine engine includes a plurality of vane segments. The vane segments are fastened together and form an adjustable stator vane that pivots about a variable vane axis. The adjustable stator vane includes a stator vane body, a shaft and a flange. The stator vane body extends axially between a first end and a second end, and includes an airfoil, a body surface and a cavity. The body surface is located at the first end. The cavity extends axially from an inlet in the body surface and into the airfoil. The shaft extends along the variable vane axis from the first end. The flange extends circumferentially at least partially around the inlet, and radially from the stator vane body. A first of the vane segments includes the flange. A second of the vane segments includes at least a portion of the airfoil.

Abstract: A component according to an exemplary aspect of the present disclosure includes, among other things, a shell defining an interior, a spar extending into the interior and a first flange attached to the spar. The spar is configured to pivot to change a positioning of the shell.

Abstract: A vane support system includes a frame and a vane. The frame has a first end configured to engage to a first platform and a second end configured to engage a second platform, so the frame can structurally support at least one of the first platform and the second platform. The first and second ends define a vane axis therebetween. The vane is mounted to the frame about the vane axis. A gas turbine engine includes a case defining a centerline axis of the engine, an inner housing and a plurality of variable vanes. The inner housing is radially inward of the case with respect to the centerline axis. At least one of the variable vanes structurally supports the case and the inner housing in response to at least one of radial, axial or tangential loads with respect to the centerline axis.

Abstract: Aspects of the disclosure are directed to a seal configured to interface with at least a first component and a second component of a gas turbine engine. A method for forming the seal includes obtaining an ingot of a fine grained, or a coarse grained, or a columnar grained or a single crystal material from a precipitation hardened nickel base superalloy containing at least 40% by volume of the precipitate of the form Ni3(Al, X), where X is a metallic or refractory element, and processing the ingot to generate a sheet of the material, where the sheet has a thickness within a range of 0.010 inches and 0.050 inches inclusive.