Abstract: A turbine section for a gas turbine engine includes inner and outer diameter ceramic matrix composite (CMC) support rings that define a gaspath there between. Each of the inner and outer diameter CMC support rings are monolithic and continuous. CMC vane arc segments are disposed in the gaspath and supported by the inner and outer diameter CMC support rings. Each of the CMC vane arc segments includes inner and outer platforms and an airfoil section there between. At least one retainer engages the inner or outer diameter CMC support rings with the CMC vane arc segments to retain the CMC vane arc segments between the inner and outer diameter CMC support rings.

Abstract: A method of improving surface roughness of ceramic matrix composites (CMCs) is provided. The method includes completing a formation of the CMCs and a chemical vapor infiltration (CVI) process to initially coat the CMCs, inspecting a CMC surface, identifying, from a result of the inspecting, a defect in the CMC surface that negatively impacts a surface roughness characteristic thereof, locally targeting and ablating the defect and re-inspecting the CMC surface to ensure that the defect is correct.

Abstract: An article includes a ceramic matrix composite (CMC) vane arc segment for disposal about an engine central axis. The CMC vane arc segment includes first and second platforms and an airfoil section that extends therebetween. The airfoil section defines a radial stacking axis that is oblique to a Z-axis that is perpendicular to the engine central axis such that the radial stacking axis forms a non-zero angle with the Z-axis.

Abstract: A CMC component for a gas turbine engine includes an airfoil including a leading edge, a trailing edge, a suction sidewall, and a pressure sidewall having a locally-thickened region, the locally-thickened region being a maximum thickness region of the pressure sidewall. The locally-thickened region has a first thickness, a maximum thickness region of the suction sidewall has a second thickness, and the first thickness is greater than the second thickness.

Abstract: A gas turbine engine includes a ceramic matrix composite (CMC) vane arc segment that has first and second platforms and an airfoil section that extends radially therebetween. The airfoil section includes an internal through-cavity, and the first and second platforms include, respectively, platform inlet and outlet ports connected to the internal through-cavity for conveying cooling air. The CMC vane arc segment is radially mounted between first and second metallic vane supports. The second metallic vane support includes a plenum and a plenum inlet port connected with the platform outlet port for receiving the cooling air from the internal through-cavity. There is a seal located radially between the second platform and the second metallic vane support. The seal circumscribes the platform outlet port to limit leakage of the cooling air between the second platform and the second metallic vane support.

Abstract: An article includes a ceramic matrix composite (CMC) vane arc segment for disposal about an engine central axis. The CMC vane arc segment includes first and second platforms and an airfoil section that extends therebetween. The airfoil section defines a radial stacking axis that is oblique to a Z-axis that is perpendicular to the engine central axis such that the radial stacking axis forms a non-zero angle with the Z-axis.

Abstract: A component for a gas turbine engine includes a matrix composite component having a radially outer end and a radially inner end. The ceramic matrix component having an internal chamber defined by an inner surface. A spar is received within the internal cavity, and spaced from an inner surface of the matrix component defining a chamber with the inner surface. Flow guides are formed on one of an outer surface of the spar and the inner surface of the matrix component. The flow guides direct airflow towards a portion of the inner surface. An air inlet chamber is formed at one radial end of the spar and an air outlet chamber formed at an opposed radial end of the spar. The air inlet chamber is defined such that air will flow into the internal chamber, outwardly of the spar, and inwardly of the inner surface of the matrix component. A gas turbine engine is also disclosed.

Abstract: A gas turbine engine includes a circumferential row of vanes arranged about a central engine axis. The vanes include respective internal cavities and vane outlet ports for conveying cooling air. A tangential onboard injector (TOBI) radially supports the vanes at an inner diameter of the circumferential row. The TOBI includes fore and aft annular walls and an outer diameter annular wall. The fore and aft annular walls and the outer diameter annular wall define an annular plenum. The outer diameter annular wall includes TOBI inlet ports that are connected, respectively, with the vane outlet ports to receive the cooling air from each of the vanes into the plenum. The TOBI includes axially-oriented nozzles for discharging the cooling air from the plenum in an aft direction.

Abstract: A method of improving surface roughness of ceramic matrix composites (CMCs) is provided. The method includes completing a formation of the CMCs and a chemical vapor infiltration (CVI) process to initially coat the CMCs, inspecting a CMC surface, identifying, from a result of the inspecting, a defect in the CMC surface that negatively impacts a surface roughness characteristic thereof, locally targeting and ablating the defect and re-inspecting the CMC surface to ensure that the defect is correct.

Abstract: A vane arc segment 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 platform adjacent the first fairing platform and a spar leg that extends from the spar platform and through the hollow airfoil section. The spar leg has an end portion that is distal from the spar platform and that protrudes from the second fairing platform. The end portion has a clevis mount. There is a support platform adjacent the second fairing platform. The support platform has a through-hole through which the end portion of the spar leg extends such that the clevis mount protrudes from the support platform. A pin extends though the clevis mount and locks the support platform to the spar leg such that the airfoil fairing is trapped between the spar platform and the support platform.

Abstract: An airfoil comprises a wall that defines a leading edge and a trailing edge and one or more cavities located within the wall along with a rib that separates the cavities. The rib or the wall comprises a first split ply that comprises a consolidated section and two or more split sections; wherein the split sections emanate from the consolidated section; and where the split sections define the wall and the cavities of the airfoil.

Abstract: A mandrel suitable for supporting an airfoil preform includes a first piece extending along a longitudinal axis of the mandrel, and a second piece in physical contact with the first piece at first interface region and extending along the longitudinal axis of the mandrel. The first piece is formed from a first material, and the second piece is formed from a second material.

Abstract: An airfoil comprises a wall that defines a leading edge and a trailing edge and one or more cavities located within the wall along with a rib that separates the cavities. The rib or the wall comprises a first split ply that comprises a consolidated section and two or more split sections; wherein the split sections emanate from the consolidated section; and where the split sections define the wall and the cavities of the airfoil.

Abstract: A gas turbine engine includes a turbine section disposed about an engine axis. The turbine section includes inner and outer diameter ceramic support rings that define a gaspath there between. Each of the inner and outer diameter ceramic support rings is monolithic and continuous. Ceramic vane arc segments are disposed in the gaspath and supported by the inner and outer diameter ceramic support rings. Each of the ceramic vane arc segments includes inner and outer platforms and an airfoil section there between. At least one retainer engages the inner or outer diameter ceramic support ring with the ceramic vane arc segments to retain the ceramic vane arc segments between the inner and outer diameter ceramic support rings.

Abstract: A vane arc segment includes a ceramic matrix composite fairing that has first and second platforms and an airfoil section that extends in a radial direction there between. Each of the first and second platforms includes axially-facing leading and trailing sides, a core gaspath side, and a non-core gaspath side. The non-core gaspath side of the first platform has a flange that projects radially there from and extends adjacent the trailing side. The flange has a radial face that is sloped with respect to the radial direction.

Abstract: A turbine section for a gas turbine engine includes inner and outer diameter ceramic matrix composite (CMC) support rings that define a gaspath there between. Each of the inner and outer diameter CMC support rings are monolithic and continuous. CMC vane arc segments are disposed in the gaspath and supported by the inner and outer diameter CMC support rings. Each of the CMC vane arc segments includes inner and outer platforms and an airfoil section there between. At least one retainer engages the inner or outer diameter CMC support rings with the CMC vane arc segments to retain the CMC vane arc segments between the inner and outer diameter CMC support rings.

Abstract: An airfoil includes an airfoil section that has an airfoil wall that circumscribes an internal airfoil cavity and defines leading and trailing ends and pressure and suction sides. The airfoil wall is formed of a ceramic matrix composite (CMC) laminate comprised of fiber plies disposed in a ceramic matrix. The fiber plies include a skin fiber ply that defines an exterior of the airfoil section and at least one inner fiber ply that lines the skin fiber ply. At least one inner fiber ply locally terminates at one of the leading end or the trailing end such that the airfoil wall has fewer of the fiber plies in the leading end or the trailing end than in each of the pressure side and the suction side.

Abstract: A gas turbine engine component includes a supply cavity and a manifold cavity that shares a common divider wall with the supply cavity. The common divider wall includes inlet metering holes that connect the supply cavity and the manifold cavity. An exterior wall has an exterior surface and an opposed interior surface that bounds portions of the supply cavity and of the manifold cavity. Outlet cooling holes extend through the exterior wall and connect the manifold cavity with the exterior surface. The number of the inlet metering holes is equal to or less than the number of the outlet cooling holes, and at least one of the inlet holes is coaxial with at least one of the outlet holes.

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 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.