Abstract: A casting core assembly for a gas turbine engine component according to an implementation includes a skin core corresponding to a first cooling passage of an airfoil. The first cooling passage includes a first section and a tip flag section joined at a first bend. The skin core includes a first portion corresponding to the first section and a tip flag portion corresponding to the tip flag section. The skin core includes at least one arcuate slot corresponding to at least one turning vane of the airfoil. A method of forming an airfoil for a gas turbine engine is also disclosed.

Abstract: A combustor includes a combustion chamber and a liner that bounds at least a portion of the combustion chamber. The liner includes a first side facing the combustion chamber and a second side facing away from the combustion chamber. The liner is formed of a lay-up of ceramic matrix composite (CMC) plies that have a first CMC ply on the first side, a second CMC ply on the second side, and intermediate CMC plies between the first and second CMC plies. At least one of the intermediate CMC plies has a pattern of voids that define cooling channels in the combustor panel. The cooling channels are bound on lateral channel sides by the at least one of the intermediate CMC plies.

Abstract: A method of fabricating a variable geometry space filling insert for a ceramic matrix composite component includes arranging a plurality of fiber bodies into a preform insert, applying a polymer binder to the plurality of fiber bodies, trimming at least a first subset of the plurality of fiber bodies such that each of the first subset of fiber bodies is shorter than at least a second subset of fiber bodies, shaping the insert with a forming tool to form a shaped insert.

Abstract: An airfoil includes an airfoil section that defines a trailing edge region that includes a trailing edge. The airfoil section is formed of a ceramic matrix composite that includes core fiber plies and skin fiber plies. The core fiber plies define a radial tube that has a radiused end in the trailing edge region. The skin fiber plies wrap around the core fiber plies and a filler element in the trailing edge region is aft of the internal cavity and is sandwiched between the skin fiber plies on the pressure side and the skin fiber plies on the suction side. There is at least one cooling passage that includes a first, inlet orifice section that opens to the internal cavity at a location forward of the radiused end and that extends through the core fiber plies, and a second, outlet orifice section that extends through the trailing edge.

Abstract: A gas turbine engine combustor panel assembly includes a shell defining an outer periphery of a combustion chamber with an outer skin with a first plurality of holes, a combustor panel including a first end, a second plurality of holes extending through an inner face to an inner void, to an outer face with a third plurality of holes to allow fluid flow from the first plurality of holes to the combustion chamber, a plurality of bolt holes, a second end with an internal combustion chamber section extending axially from a central axis and from a panel wall of the combustor panel. The assembly includes a sheet metal seal plate abutting the plurality of bolt holes and the interface, and a plurality of bolts extending along a central axis and interfacing with the plurality of bolt holes to connect the combustor panel and shell.

Abstract: A combustor includes a combustion chamber and a liner that bounds at least a portion of the combustion chamber. The liner includes a first side facing the combustion chamber and a second side facing away from the combustion chamber. The liner is formed of a lay-up of ceramic matrix composite (CMC) plies that have a first CMC ply on the first side, a second CMC ply on the second side, and intermediate CMC plies between the first and second CMC plies. At least one of the intermediate CMC plies has a pattern of voids that define cooling channels in the combustor panel. The cooling channels are bound on lateral channel sides by the at least one of the intermediate CMC plies.

Abstract: An airfoil for a gas turbine engine includes an airfoil section and a platform that has a non-gaspath side, a flange that extends from the non-gaspath side, and a gaspath side from which the airfoil section extends. The flange is comprised of a sandwich composite that includes first and second ceramic matrix composite (CMC) skins that each have at least one 2-D ceramic fiber ply, and a cellular core disposed between the first and second CMC skins.

Abstract: A method for fabricating an airfoil for a gas turbine engine includes providing a core blank made of a cellular material, shaping the core blank into a cellular core, forming a fiber preform that has an airfoil section and a platform by laying-up first and second ceramic fiber ply skins on the cellular core such that in the platform the cellular core is sandwiched radially between the first and second ceramic fiber ply skins, the first and second ceramic fiber ply skins each include at least one 2-D ceramic fiber ply, and densifying the fiber preform with a ceramic matrix.

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 casting core assembly for a gas turbine engine component according to an implementation includes a skin core corresponding to a first cooling passage of an airfoil. The first cooling passage includes a first section and a tip flag section joined at a first bend. The skin core includes a first portion corresponding to the first section and a tip flag portion corresponding to the tip flag section. The skin core includes at least one arcuate slot corresponding to at least one turning vane of the airfoil. A method of forming an airfoil for a gas turbine engine is also disclosed.

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 section of a gas turbine engine includes a ceramic component, a metallic component, a seal situated between the metallic component and the ceramic component at a sealing interface, a silicon-based coating disposed on the ceramic component at the sealing interface, and a non-reactive coating disposed on the metallic component at the sealing interface. The non-reactive coating provides thermochemical protection against interaction between the ceramic component, the silicon-based coating, and the metallic component. A method of providing a seal in a gas turbine engine is also disclosed.

Abstract: A method of assembly includes locating a plurality of vanes about an inner circumference of a ring, the ring has a plurality of axially-extending projections circumferentially-spaced around the ring, each of the vanes has a radially outer platform, a flange that projects radially outwardly from the radially outer platform, and a notch that extends axially through the flange such that the notch opens at forward and aft flange faces of the flange, aligning the notch of each of the vanes with a corresponding one of the axially-extending projections on the ring, and inserting the projection into the notch to thereby prevent the vanes from rotating relative to the ring.

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: 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: An airfoil for a gas turbine engine includes an airfoil section and a platform that has a non-gaspath side and a gaspath side from which the airfoil section extends. The platform is comprised of a sandwich composite that has first and second ceramic matrix composite (CMC) skins that each include at least one 2-D ceramic fiber ply. The first and second CMC skins are disposed, respectively, on the gaspath side and the non-gaspath side, and there is a cellular core disposed radially between the first and second CMC skins.

Abstract: A fixture for processing a ceramic matrix composite article includes first and second tool segments that each have an exterior side and an interior side. The tool segments define there between a working region for holding a fiber preform that is to be densified with a ceramic material. Each of the interior sides has grooves there along for delivering a precursor gas to the working region.

Abstract: A method of coating a seal slot includes applying a coating to a slot in a component and machining the coating to provide an open height configured to receive a seal. After the machining a surface roughness of the coating is less than about 100 ra. A scaling assembly is also disclosed.

Abstract: A plurality of static stator vane are circumferentially spaced in a row axially intermediate the rows of turbine blades. Each stator vane has at least an outer platform with mount structure. Each of the stator vanes are formed of composite materials. The mount structure is provided with sacrificial material. At least a first of the stator vanes is circumferentially adjacent to a second of the stator vanes. An orientation of the first of the stator vanes is re-staggered relative to the second of the stator vanes. The sacrificial material of the first of the stator vanes is machined away such that a final trailing edge of the first of the stator vanes mount structure is now better axially aligned with the trailing edge of the second of the stator vanes mount structure trailing edge.

Abstract: A plurality of static stator vane are circumferentially spaced in a row axially intermediate the rows of turbine blades. Each stator vane has at least an outer platform with mount structure. Each of the stator vanes are formed of composite materials. The mount structure is provided with sacrificial material. At least a first of the stator vanes is circumferentially adjacent to a second of the stator vanes. An orientation of the first of the stator vanes is re-staggered relative to the second of the stator vanes. The sacrificial material of the first of the stator vanes is machined away such that a final trailing edge of the first of the stator vanes mount structure is now better axially aligned with the trailing edge of the second of the stator vanes mount structure trailing edge.