Abstract: A gas turbine engine stator vane combination includes a stator vane body having an airfoil with a leading edge, a trailing edge, a suction side and a pressure side, and having at least one internal cooling channel. The stator vane body is formed of ceramic matrix composite materials. A spar is received within the at least one cooling air channel and formed of a metal. The spar has an outer peripheral surface spaced from an inner peripheral surface of the cooling air channel to define a cooling air path, with the cooling air path having a varying cross-sectional area through the cooling air channel. A gas turbine engine is also disclosed.

Abstract: A flange stiffener support for a duct in a gas turbine engine includes a circumferential flange segment that has a plurality of flange fastener openings. An axial body segment extends from a radially inner side of the circumferential segment and has at least one body fastener opening. The axial body segment extends at least partially along a circumferential distance of the circumferential flange segment.

Abstract: A gas turbine engine stator vane combination includes a stator vane body having an airfoil with a leading edge, a trailing edge, a suction side and a pressure side, and having at least one internal cooling channel. The stator vane body is formed of ceramic matrix composite materials. A spar is received within the at least one cooling air channel and formed of a metal. The spar has an outer peripheral surface spaced from an inner peripheral surface of the cooling air channel to define a cooling air path, with the cooling air path having a varying cross-sectional area through the cooling air channel. A gas turbine engine is also disclosed.

Abstract: An airfoil includes an airfoil wall that has a span between first and second radial ends and that defines leading and trailing edges, pressure and suction sides each joining the leading and trailing edges, and an internal cavity. The internal cavity is divided into at least first and second sub-cavities extending over the span. The first sub-cavity defines a venturi tube.

Abstract: A method of forming a ceramic matrix composite airfoil includes forming a fibrous ceramic preform by laying up a plurality of plies to form a hollow core cavity, overwrapping the plurality of plies of the core cavity with a first overwrap layer, weaving together a tows of first free portion and tows of a second free portion of the first overwrap layer to form a cross brace, overwrapping the plurality of plies of the core cavity and the first overwrap layer with a second overwrap layer, and consolidating the first free portion and the second free portion of the first overwrap layer and a third free portion and a fourth free portion of the second overwrap layer to form a trailing edge.

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 according to an aspect of the present disclosure includes a metallic damper including a first contact surface and a gas turbine engine component. The gas turbine engine component includes a main body extending in a first direction between a gaspath surface and a second contact surface. The first and second contact surfaces oppose each other along an interface extending in a second direction. The first and second contact surfaces are dimensioned to contact each other along the interface in a hot assembly state. The main body is established by a composite including fibers in a matrix material. At least some of the fibers are arranged to establish a plurality of cooling passages aligned with the interface relative to the second direction. A method of damping 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 method of forming a ceramic matrix composite airfoil includes forming a fibrous ceramic preform by laying up a plurality of plies to form a hollow core cavity, overwrapping the plurality of plies of the core cavity with a first overwrap layer, weaving together a tows of first free portion and tows of a second free portion of the first overwrap layer to form a cross brace, overwrapping the plurality of plies of the core cavity and the first overwrap layer with a second overwrap layer, and consolidating the first free portion and the second free portion of the first overwrap layer and a third free portion and a fourth free portion of the second overwrap layer to form a trailing edge.

Abstract: Disclosed herein is a method of reinforcing a composite comprising determining a location of a first cooling hole in a plurality of plies; where a cooling gas is transported through the cooling hole; disposing a z-fiber in the plurality of plies at a location proximate to where the first cooling hole will be located; where the z-fiber enters the plurality of plies at either an upper surface or a lower surface; and where the z-fiber traverses a portion of the plurality of plies in the z-direction proximate to the first cooling hole; and traverses the plurality of plies in an x or y direction further away from the first cooling hole; where the z-direction is in the thickness direction of the plurality of plies and where the x and y-direction are perpendicular to the z-direction.

Abstract: An assembly for a gas turbine engine according to an aspect of the present disclosure includes a metallic damper including a first contact surface and a gas turbine engine component. The gas turbine engine component includes a main body extending in a first direction between a gaspath surface and a second contact surface. The first and second contact surfaces oppose each other along an interface extending in a second direction. The first and second contact surfaces are dimensioned to contact each other along the interface in a hot assembly state. The main body is established by a composite including fibers in a matrix material. At least some of the fibers are arranged to establish a plurality of cooling passages aligned with the interface relative to the second direction. A method of damping for a gas turbine engine is also disclosed.

Abstract: Disclosed herein is a composite comprising a ceramic matrix composite comprising a ceramic matrix and ceramic fibers; and one or more sacrificial fibers woven into the ceramic matrix composite; where the sacrificial fibers are operative to undergo oxidation or melting upon being subjected to an elevated temperature; and wherein the sacrificial fibers leave cooling holes in the composite preform upon being subjected to oxidation or melting.

Abstract: A vane arc segment includes an airfoil fairing that has a fairing platform and an airfoil section that extends there from. The fairing platform defines a gaspath side and a non-gaspath side and includes a flange that projects from the non-gaspath side. The airfoil fairing is formed of a fiber-reinforced composite that includes a wishbone-shaped fiber layer structure that has first and second arms that converge and merge into a single leg. The first and second arms are formed of fiber plies comprised of a network of fiber tows. The single leg comprises fiber tows from each of the fiber plies of the first and second arms. The fiber tows of the first arm are interwoven in the single leg with the fiber tows of the second arm. The first arm, the second arm, or the single leg forms at least a portion of the flange.

Abstract: Plated polymeric gas turbine engine parts and methods for fabricating lightweight plated polymeric gas turbine engine parts are disclosed. The parts include a polymeric substrate plated with one or more metal layers. The polymeric material of the polymeric substrate may be structurally reinforced with materials that may include carbon, metal, or glass. The polymeric substrate may also include a plurality of layers to form a composite layup structure.

Abstract: Plated polymeric gas turbine engine parts and methods for fabricating lightweight plated polymeric gas turbine engine parts are disclosed. The parts include a polymeric substrate plated with one or more metal layers. The polymeric material of the polymeric substrate may be structurally reinforced with materials that may include carbon, metal, or glass. The polymeric substrate may also include a plurality of layers to form a composite layup structure.

Abstract: A rotor assembly for a gas turbine engine is disclosed. The assembly includes: a composite fan blade, the fan blade including a root; a metallic rotor including a slot for receiving the root; the root being at least partially coated with a metal to form a metal-coated portion; the metal-coated portion of the root being at least partially covered with an intermediate material; and the root, metal-coated portion and intermediate material being received in the slot and bonded to the rotor.

Abstract: Plated polymeric gas turbine engine parts and methods for fabricating lightweight plated polymeric gas turbine engine parts are disclosed. The parts include a polymeric substrate plated with one or more metal layers. The polymeric material of the polymeric substrate may be structurally reinforced with materials that may include carbon, metal, or glass. The polymeric substrate may also include a plurality of layers to form a composite layup structure.

Abstract: A nosecone for a turbine engine includes a nosecone body and a nosecone mount. The nosecone body extends along an axis between a tip end and a base end. The nosecone body is configured from or otherwise includes thermoplastic material. The nosecone body includes a shell and an arrangement of ribs, which structurally support at least a portion of the shell. A thickness of the arrangement of ribs is greater than or substantially equal to approximately one half of a thickness of the shell. The nosecone mount is adapted to connect the nosecone body to a component of the turbine engine.

Abstract: The present disclosure relates generally to a turbine engine case assembly, the turbine engine case assembly including a fan case including an outer frame encircling an axis, the outer frame including an outer frame exterior surface and an outer frame interior surface, at least two mounts disposed on the outer frame exterior surface, wherein the at least two mounts are circumferentially spaced along the outer frame exterior surface, and a compliant attachment device operably coupled to the at least two mounts.

Abstract: The present disclosure relates to a method of manufacturing composite airfoils bonded to a metallic root. A composite body may be formed with a metallic co-molded member. The co-molded member may be transient liquid phase (TLP) bonded to a metallic root. The metallic root may allow the composite body to be attached to a rotor. The airfoil may also have a metallic edge which is TLP bonded to the composite body via a co-molded edge.