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: A method of forming a ceramic matrix composite includes fabricating a space filling insert by assembling the space filling insert from a fibrous ceramic material, placing the space filling insert into a tooling fixture, the tooling fixture having an internal shape complementary to a shape of the insert, and infiltrating the space filling insert with a first ceramic material. The method further includes forming a preform by laying up a plurality of fibrous ceramic plies and inserting the infiltrated space filling insert into a void defined by the plurality of plies. The method further includes infiltrating the preform with a second ceramic material.

Abstract: A ceramic matrix composite laminate comprises a ceramic matrix that encapsulates a plurality of layers. Each layer comprises fibers. Each layer comprises a plurality of fill fibers and a plurality of warp fibers or a plurality of bias fibers. The outermost layer contains a different concentration of fibers per unit volume than a layer located near an interior of the ceramic matrix composite laminate. A gradient in the number of fibers exists between the outermost layer and the layer located at the interior of the ceramic matrix composite laminate, or a combination thereof. A combined ceramic matrix composite comprises a plurality of composite laminates; wherein each laminate has a different fiber concentration gradient from another laminate that it is in contact with.

Abstract: A gas turbine engine comprising a ring having an inner ring and an outer ring that define a gas path therebetween, at least one of the inner ring and the outer ring include two or more rings; a vane disposed in the gas path and supported by the inner ring and the outer ring, the vane includes an inner platform and an outer platform, wherein at least one of the inner rings and outer rings include one or more flange designed to abut an opposing flange defined by at least one of the inner platform and the outer platform, a contact surface designed to abut an opposing contact surface defined by at least one of the inner platform and the outer platform, and an angled contact surface designed to abut an opposing angled contact surface defined by at least one of the inner platform and the outer platform.

Abstract: A mandrel system and method for may be used to produce a closed-end hollow ceramic matrix composite (CMC) part with a negative draft, with the CMC part having a proximal open end and a distal closed end. The system includes a first layer of mandrel configured for a first portion of the CMC part and having a first distal surface adjacent the closed-end of the CMC part, the first layer being formed of a first plurality of segments. The system also includes a second layer of mandrel configured for a second portion of the CMC part and having a second distal surface adjacent a proximal surface of the first layer, the second layer being formed of a second plurality of segments. The first and second plurality of segments are configured to be extractable from the proximal open end of the CMC part.

Abstract: An airfoil includes an airfoil section formed of a ceramic matrix composite having fiber plies in a ceramic matrix. The fiber plies include core fiber plies defining a radial tube circumscribing an internal cavity, and skin fiber plies defining an exterior of the airfoil section and wrapping around the core fiber plies from the pressure side of the trailing edge region, through the leading edge, and to the suction side of the trailing edge region. A filler element is disposed in the trailing edge region aft of the internal cavity, sandwiched between the skin fiber plies on the pressure and suction sides. At least one cooling passage includes an inlet orifice section opening to the internal cavity, a radially-elongated serpentine plenum winding chordally through the fiber plies, outlet orifice sections extending through the trailing edge, and intermediate passage sections connecting the serpentine plenum with the outlet orifice sections.

Abstract: A ceramic matrix composite (CMC) airfoil component can be made by forming a stack of fabric plies with a plurality of plies adhered together with a tackifier. The stack of fabric plies is machined into an airfoil shape to form an airfoil core with an orientation of the fabric plies selected based upon strength and/or performance requirements of the airfoil component. The airfoil core is then overwrapped with an outer mold layer formed of one or more plies of fabric to form a preform. The preform is then densified, such as via melt infiltration (MI) and/or chemical vapor infiltration (CVI), to form the CMC airfoil component such as a turbine vane or a turbine blade.

Abstract: A preform comprising a first sub-laminate comprising a plurality of layers and a second sub-laminate comprising a plurality of layers. The first sub-laminate comprises a first unit cell comprising a first volume fraction of tows, where the first volume fraction of tows comprise first tows having a first tow spacing between successive first tows. The second sub-laminate comprises a second unit cell comprising a second volume fraction of tows, where the second volume fraction of tows comprise second tows having a second tow spacing between successive second tows. The first volume fraction of tows in the first unit cell is equal to the second volume fraction of tows in the second unit cell. The second tow spacing is less than the first tow spacing.

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 CMC airfoil includes core fiber plies that define a radial tube that circumscribes an internal cavity, and skin fiber plies that define an exterior of the airfoil section. There is a filler element in the trailing edge region aft of the internal cavity and sandwiched between the skin fiber plies on the pressure side and the skin fiber plies on the suction side. At least one cooling passage includes at least one inlet orifice section that opens to the internal cavity and extends through the core fiber plies, a radially-elongated multi-dimensional plenum connected with the at least one inlet orifice section, a plurality of outlet orifice sections extending through the trailing edge, and a plurality of intermediate passage sections bound by at least one of the skin plies and the filler element and connecting the radially-elongated multi-dimensional plenum with the plurality of outlet orifice sections.

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 method of forming a ceramic matrix composite (CMC) vane including an inner diameter platform, an outer diameter platform, and an airfoil extending between the inner diameter platform and the outer diameter platform includes laying up a ceramic preform by arranging at least one core formed from a first ceramic material, and laying up, over the at least one ceramic core, a plurality of contoured plies, each of the plurality of contoured plies formed from a second ceramic material. Each of the plurality of contoured plies partially forms the inner diameter platform, the outer diameter platform, and the airfoil of the vane. The method further includes densifying the ceramic preform with a matrix.

Abstract: Systems and methods are disclosed for producing a hollow composite part, such as a CMC airfoil, using a segmented mandrel formed of individual mandrel segments. The individual mandrel segments may be dimensioned to permit extraction from highly twisted and/or curved parts.

Abstract: A method of reinforcing a ceramic sandwich structure includes forming a ceramic core, assembling the sandwich structure by disposing the core between and in physical contact with a first panel and a second panel, and weaving a plurality of z-fibers at least partially through a thickness of the sandwich structure.

Abstract: A method of fabricating a seal arc segment includes laying-up at least one radially inner-most ceramic fabric ply on a mandrel, laying-up the at least one discontinuous radially intermediate ceramic fabric ply on the at least one inner-most ceramic fabric ply, the at least one discontinuous radially intermediate ceramic fabric ply having cutout voids, laying-up at least one radially outer-most ceramic fabric ply on the at least one discontinuous radially intermediate ceramic fabric ply with the cutout voids, the radially inner-most ceramic fabric ply, the at least one discontinuous radially intermediate ceramic fabric ply, and the radially outer-most ceramic fabric ply together providing a seal arc segment fiber preform, and densifying the seal arc segment fiber preform with a ceramic matrix material to form a seal arc segment wall of a seal arc segment, wherein the cutout voids define cooling channels in the seal arc segment wall.

Abstract: A combined gas turbine engine vane and blade outer air seal assembly includes a vane having an airfoil extending from the leading edge to a trailing edge, and has an outer platform. The outer platform has a cooling channel that extends into the airfoil to receive cooling air. The outer platform extends to an integral blade outer air seal to be positioned radially outwardly of a turbine blade in a gas turbine engine. At least a portion of the vane and the blade outer air seal are formed of ceramic matrix composite materials. A gas turbine engine is also disclosed.

Abstract: A method of forming a ceramic matrix composite (CMC) vane including an inner diameter platform, an outer diameter platform, and an airfoil extending between the inner diameter platform and the outer diameter platform includes laying up a ceramic preform by arranging at least one core formed from a first ceramic material, and laying up, over the at least one ceramic core, a plurality of contoured plies, each of the plurality of contoured plies formed from a second ceramic material. Each of the plurality of contoured plies partially forms the inner diameter platform, the outer diameter platform, and the airfoil of the vane. The method further includes densifying the ceramic preform with a matrix.

Abstract: A method of fabricating a ceramic matrix composite component includes fabricating a ceramic preform, the preform comprising a hollow portion with an internal cavity extending along a first axis from a first end to a second end of the hollow portion, supporting the hollow portion with an auxetic mandrel disposed within and the internal cavity and coaxial with the hollow portion, the auxetic mandrel comprising a first mandrel end and a second mandrel end, at least partially densifying the preform with a matrix, and removing the auxetic mandrel from the internal cavity by simultaneously applying a compressive force to each of the first mandrel end and the second mandrel end creating a deformed auxetic mandrel to reduce a cross-sectional profile of the auxetic mandrel along a second axis orthogonal to the first axis, and extracting the deformed auxetic mandrel from the first end of the hollow portion.

Abstract: A combined gas turbine engine vane and blade outer air seal assembly includes a vane having an airfoil extending from the leading edge to a trailing edge, and has an outer platform. The outer platform has a cooling channel that extends into the airfoil to receive cooling air. The outer platform extends to an integral blade outer air seal to be positioned radially outwardly of a turbine blade in a gas turbine engine. At least a portion of the vane and the blade outer air seal are formed of ceramic matrix composite materials. A gas turbine engine is also disclosed.

Abstract: A gas turbine engine component includes an inner surface and an outer surface. The outer surface and the inner surface are formed of ceramic matrix composite layers, and mount member. The mount member is in contact with an insert that is inserted within outer layers of the outer platform to form a thickened portion having an end surface that will react against forces from the mount member. A gas turbine engine and a method are also disclosed.