Abstract: A liner for a combustor includes a support member, an intermediate member, and a liner member. The intermediate member is positioned intermediate the support member and the liner member and has a plurality of protrusions and a plurality of recesses. The support member is coupled to the intermediate member at a tangent of each protrusion. Additionally, the liner member is comprised of a ceramic matrix composite material. The liner member is coupled to the intermediate member at a tangent of each recess.

Abstract: A gas turbine engine composite vane assembly and method for making same are disclosed. The method includes providing at least two gas turbine engine airfoil composite preform components. The airfoil composite preform components are interlocked with a first locking component so that mating faces of the airfoil composite preform components face each other. A filler material is inserted between the mating surfaces of the airfoil composite preform components.

Abstract: A turbine vane assembly adapted for use in a gas turbine engine includes a vane support and a vane heat shield comprising ceramic matrix composite materials to insulate the metallic materials of the vane support. The turbine vane assembly further includes an aero-load transfer rib located outside of a primary gas path and configured to carry loads applied to the vane heat shield to the vane support.

Abstract: A combustor adapted for use in a gas turbine engine is disclosed. The combustor includes a metallic case forming a cavity and a ceramic liner arranged in the cavity of the metallic case. The ceramic liner defines a combustion chamber in which fuel is burned during operation of a gas turbine engine. The ceramic liner is located in the metallic case using a plurality of cross key connectors.

Abstract: One aspect of the present disclosure includes a turbine vane assembly comprising a vane made from ceramic matrix composite material having an outer wall extending between a leading edge and a trailing edge and between a first end and an opposing second end; an endwall made at least partially from a ceramic matrix composite material configured to engage the first end of the vane; and a retaining region including corresponding bi-cast grooves formed adjacent the first end of the vane and a receiving aperture formed in the endwall; wherein a bond is formed in the retaining region to join the vane and endwall together.

Abstract: A gas turbine engine composite vane assembly and method for making same are disclosed. The method includes providing at least two gas turbine engine airfoil composite preform components. The airfoil composite preform components are interlocked with a first locking component so that mating faces of the airfoil composite preform components face each other. A filler material is inserted between the mating surfaces of the airfoil composite preform components.

Abstract: A turbine vane assembly adapted for use in a gas turbine engine includes a vane support and a vane heat shield comprising ceramic matrix composite materials to insulate the metallic materials of the vane support. The turbine vane assembly further includes an aero-load transfer rib located outside of a primary gas path and configured to carry loads applied to the vane heat shield to the vane support.

Abstract: A blade for a gas turbine engine comprises a blade portion having a first end and a second end and an engagement portion including a first surface coupled to the second end of the blade portion and a second surface coupled to the second end of the blade portion, the first and second surfaces arranged to extend divergently away from one another. The engagement portion is adapted for coupling to a wheel included in a gas turbine engine wheel.

Abstract: A method of forming a composite preform containing multiple laminates is disclosed. The method may include providing a first sublaminate comprising stacked fibers woven into a fabric; providing a second sublaminate comprising stacked fibers woven into a fabric; joining the first sublaminate and the second sublaminate forming a component comprising a region of discontinuity sandwiched between the first sublaminate and the second sublaminate; rigidizing the component; and softening the region between the first sublaminate and the second sublaminate. In illustrative embodiments, the method may include manipulating the region of discontinuity between the first sublaminate and the second sublaminate to reduce the incoherence between the sublaminates by moving fibers from the sublaminates through at least part of the region between the first sublaminate and the second sublaminate.

Abstract: A ceramic matrix composite having improved operating characteristics includes a barrier layer.

Abstract: A gas turbine engine composite vane assembly and method for making same are disclosed. The method includes providing at least two gas turbine engine airfoil composite preform components. The airfoil composite preform components are interlocked with a first locking component so that mating faces of the airfoil composite preform components face each other. A filler material is inserted between the mating surfaces of the airfoil composite preform components.

Abstract: A turbine shroud for a gas turbine engine includes an annular metallic carrier, a blade track, and a cross-key connection formed between the annular metallic carrier and the ceramic blade track. The cross-key connection is formed between the annular metallic carrier and inserts included in the blade track. The inserts are bonded to an annular runner also included in the blade track by a braze layer.

Abstract: A method of making a fiber tow coating is provided. The method includes providing a fiber tow selected from the group consisting of carbon and silicon; and applying an oxide-based fiber interface coating onto the fiber tow using directed vapor deposition or other like deposition method.

Abstract: A method of making a fiber tow coating is provided. The method includes providing a fiber tow selected from the group consisting of carbon and silicon; and applying an oxide-based fiber interface coating onto the fiber tow using directed vapor deposition or other like deposition method.

Abstract: A barrier coating for isolating a metallic support component from a composite component in a gas turbine engine is provided. The barrier coating may be applied to the metallic support component so that when the ceramic component is mounted on the metallic support component the barrier coating is engaged.

Abstract: Creating a ceramic matrix structure may include alternating different layers of fiber fabric to achieve porosity, infiltration, and/or other objectives. Different layers of the ceramic matrix structure may be configured differently to achieve different objectives in different regions of the ceramic matrix structure. The fiber tow spacing, fiber count and/or other characteristics may be varied within individual layers and/or among different layers of the ceramic matrix structure.

Abstract: One aspect of the present disclosure includes a turbine vane assembly comprising a vane made from ceramic matrix composite material having an outer wall extending between a leading edge and a trailing edge and between a first end and an opposing second end; an endwall made at least partially from a ceramic matrix composite material configured to engage the first end of the vane; and a retaining region including corresponding bi-cast grooves formed adjacent the first end of the vane and a receiving aperture formed in the endwall; wherein a bond is formed in the retaining region to join the vane and endwall together.

Abstract: One aspect of the present disclosure includes a turbine vane assembly comprising a vane made from ceramic matrix composite material having an outer wall extending between a leading edge and a trailing edge and between a first end and an opposing second end; an endwall made at least partially from a ceramic matrix composite material configured to engage the first end of the vane; and a retaining region including corresponding bi-cast grooves formed adjacent the first end of the vane and a receiving aperture formed in the endwall; wherein a bond is formed in the retaining region to join the vane and endwall together.

Abstract: A multilayer interface coating for composite material fibers includes a first coating layer deposited onto a fiber and a second coating layer deposited onto the first coating layer.

Abstract: Ceramic matrix composite materials and a process for making said composite materials are disclosed.