Abstract: A tooling assembly suitable for use in infiltrating a fibrous preform comprises an outer tooling fixture and an inner tooling fixture. The outer tooling fixture comprises a first portion removably attached to a second portion, and a first pair of oppositely disposed windows defined by the first portion and the second portion. Each of the first portion and the second portion comprise a base with a recess. The inner tooling fixture comprises a body defined by at least one portion, the body configured to at least partially surround the fibrous preform, a plurality of holes extending through a thickness of the at least one portion, each hole of the plurality of holes having a length (L) and a diameter (D), and a first neck region engageable with a recess of the first portion or the second portion.

Abstract: A vane multiplet includes first and second ceramic matrix composite (CMC) singlet vanes that are arranged circumferentially adjacent each other. Each of the CMC singlet vanes includes an airfoil section and a platform at one end of the airfoil section. The platform defines forward and trailing platform edges and first and second circumferential side edges. A CMC overwrap conjoins the CMC singlet vanes. The CMC overwrap includes fiber plies that are fused to the platforms of the CMC singlet vanes.

Abstract: An assembly for use in releasing a preform from a tooling fixture comprises a solid body engageable with the fixture via a seal, a manifold formed in the solid body and in flow communication with the preform via a plurality of holes in the tooling fixture, and a fluid source in flow communication with the manifold and configured to supply fluid to the manifold. The fluid comprises at least one of a gas and a slurry.

Abstract: An assembly for a gas turbine engine includes a plurality of vanes having a radially outer platform with a flange that extends radially outward therefrom and a plurality of notches in the flange. A ring is located radially outward from the radially outer platform of the plurality of vanes. An axially extending projection extends through and corresponds with each of the plurality of notches on the flange.

Abstract: A ceramic matrix composite airfoil includes a high-pressure surface and a low-pressure surface connected at a leading edge and a trailing edge. The high-pressure surface and the low-pressure surface extend from a first end to a second end. A leading edge cooling passage includes an inlet portion, a midspan portion and an outlet portion. A cross-sectional flow area of the midspan portion is less than a cross-sectional flow area of either the inlet portion or the outlet portion.

Abstract: An assembly for a gas turbine engine includes, among other things, an end wall including a main body extending between a first end portion and a second end portion to establish a seal face. The end wall includes a first attachment portion dimensioned to fixedly attach the main body to a static structure at a first attachment point. An airfoil extends radially inwardly from the end wall relative to an assembly axis. The airfoil includes an inner cavity extending between a first end portion and a second end portion, the first end portion adjacent the end wall of the airfoil. A spar member includes a spar body extending between a first end portion and a second end portion. The spar body extends at least partially through the inner cavity. The first end portion of the end wall is cantilevered from the first attachment point.

Abstract: A vane multiplet includes first and second ceramic matrix composite (CMC) singlet vanes that are arranged circumferentially adjacent each other. Each of the CMC singlet vanes includes an airfoil section and a platform at one end of the airfoil section. The platform defines forward and trailing platform edges and first and second circumferential side edges. A CMC overwrap conjoins the CMC singlet vanes. The CMC overwrap includes fiber plies that are fused to the platforms of the CMC singlet vanes.

Abstract: An assembly for a gas turbine engine includes a plurality of vanes having a radially outer platform with a flange that extends radially outward therefrom and a plurality of notches in the flange. A ring is located radially outward from the radially outer platform of the plurality of vanes. An axially extending projection extends through and corresponds with each of the plurality of notches on the flange.

Abstract: A ceramic matrix composite airfoil includes a high-pressure surface and a low-pressure surface connected at a leading edge and a trailing edge. The high-pressure surface and the low-pressure surface extend from a first end to a second end. A leading edge cooling passage includes an inlet portion, a midspan portion and an outlet portion. A cross-sectional flow area of the midspan portion is less than a cross-sectional flow area of either the inlet portion or the outlet portion.

Abstract: A seal system includes a ceramic component that has a non-core-gaspath surface region that defines a first surface roughness and a core gaspath surface region. A metallic component is situated adjacent the non-core-gaspath surface region. A coating system is disposed on the ceramic component. The coating system includes a silicon-containing layer on the non-core-gaspath surface region and a barrier layer that has a first section on the silicon-containing layer and a second section on the core-gaspath region and that is connected to the first section. The surface of the barrier layer has a second surface roughness that is less than the first surface roughness. The first section is in contact with the metallic component and the second section serves as an environmental barrier on the core-gaspath region.

Abstract: An assembly for a gas turbine engine according to an example of the present disclosure includes at least one platform having a main body extending between a first mate face and a second mate face to establish a gas path surface. The main body has an internal passage extending circumferentially between a first opening along the first mate face and a second opening along the second mate face relative to an assembly axis. A perimeter of the first mate face establishes a first area, a perimeter of the second opening establishes a second area, and the second area is greater than the first area. An airfoil section extends radially from the at least one platform.

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: An airfoil includes an airfoil wall that defines a leading end, a trailing end, and pressure and suction sides that join the leading end and the trailing end. The airfoil wall includes a wishbone-shaped fiber layer structure that has a pair of arms that merge into a single leg. The pair of arms are formed by first and second S-shaped fiber layers each of which is comprised of a network of fiber tows. The first and second S-shaped fiber layers merge to form the single leg. The single leg comprises fiber tows from each of the first and second S-shaped fiber layers that are interwoven, and the single leg forms at least a portion of the trailing end of the airfoil wall.

Abstract: An airfoil for a gas turbine engine according to an example of the present disclosure includes an airfoil section extending in a radial direction from a root section to a tip portion. The airfoil section has an external wall and an internal wall. The airfoil section establishes an internal cooling arrangement including a first cooling passage having a first section and a tip flag section. The tip flag section extends in the chordwise direction along the tip portion from the first section. The first section includes a plurality of branched paths established by at least one turning vane.

Abstract: A seal system includes a ceramic component that has a non-core-gaspath surface region that defines a first surface roughness and a core gaspath surface region. A metallic component is situated adjacent the non-core-gaspath surface region. A coating system is disposed on the ceramic component. The coating system includes a silicon-containing layer on the non-core-gaspath surface region and a barrier layer that has a first section on the silicon-containing layer and a second section on the core-gaspath region and that is connected to the first section. The surface of the barrier layer has a second surface roughness that is less than the first surface roughness. The first section is in contact with the metallic component and the second section serves as an environmental barrier on the core-gaspath region.

Abstract: A seal system includes a ceramic component, a metallic component, a silicon-containing layer, and a barrier layer. The ceramic component has a first surface region that defines a first surface roughness. The metallic component is situated adjacent to the first surface region and has a second surface region facing the first surface region. The silicon-containing layer is on the first surface region of the ceramic component and has a contact surface that defines a second surface roughness which is less than the first surface roughness. The barrier layer is on the metallic component and in contact with the silicon-containing layer and serves to limit interaction between silicon of the silicon-containing layer and the metallic component. The barrier layer includes at least one of alumina or MCrAlY.

Abstract: 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.

Abstract: Disclosed is a method of reducing play in a vane arc segment. The vane arc segment includes an airfoil piece that defines first and second platforms and a hollow airfoil section that has an internal cavity and that extends between the first and second platforms. The first platform defines a gaspath side, a non-gaspath side, and a radial flange that projects from the non-gaspath side. Support hardware supports the airfoil piece via the radial flange, and a thermal insulation element is located adjacent the radial flange. The method includes performing a light scan of the radial flange to produce a digital three-dimensional model of the radial flange, and then machining the thermal insulation element in accordance with the digital three-dimensional model to provide a low-tolerance fit between the radial flange and the thermal insulation element that limits play between the airfoil piece and the thermal insulation element.

Abstract: A vane assembly includes an airfoil fairing that has first and second fairing platforms and a hollow airfoil section that extends there between. A spar has a spar leg that extends through the hollow airfoil section. The spar leg has a through-passage and an end portion that protrudes from the second fairing platform. A support is secured with the end portion of the spar leg. The support has a platform that includes first and second sides, an opening that extends between the first and second sides, and a nozzle tube that extends from the second side. The first side is adjacent the second fairing platform. The end portion of the spar leg is disposed in the opening so as to fluidly connect the through-passage with the nozzle tube.

Abstract: A flowpath component for a gas turbine engine includes a body having a leading edge and a trailing edge. A first exterior wall connects the leading edge to the trailing edge and a second exterior wall connects the leading edge to the trailing edge. At least one first internal cooling passage has a first inlet at a first end of the body. At least one second internal cooling passage has a second inlet at a second end of the body. The at least one first internal cooling passage is isolated from the at least one second internal cooling passage.