A method of forming a ceramic matrix composite includes depositing particles on a ceramic fabric formed from a plurality of ceramic tows, applying a binder to at least the particles to form a stabilized ceramic fabric, forming a preform using the stabilized ceramic fabric, and densifying the preform. The ceramic tows are formed from a first material and the particles are formed from at least a second material.

A method of preparing a ceramic fabric for use in a ceramic matrix composite includes transforming a ceramic tow from a first tow geometry to a second tow geometry, thereby reducing a first dimension of the ceramic tow and increasing a second dimension of the ceramic tow orthogonal to the first dimension to produce a flattened tow. The method includes weaving or braiding the flattened ceramic tow to form a ceramic fabric.

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.

A method of machining cooling holes includes providing a workpiece in which a cooling hole is to be formed. The cooling hole, once formed, defines distinct first and second sections. The workpiece is secured in a fixture that is mounted in a first machine. In the first machine, a laser is used to drill a through-hole in a wall of the workpiece. The through-hole is spatially common to the first and second sections of the cooling hole. After drilling the through-hole, the fixture with the workpiece secured therein is removed from the first machine and mounted in a second machine. In the second machine, ultrasonic machining is used to expand a portion of the through-hole to form the second section. An abrasive slurry used in the process is drained through the through-hole during the ultrasonic machining.

A method for producing a ceramic matrix composite component is disclosed. The method includes providing a plurality of first ceramic fiber plies including a plurality of interconnected tows and a plurality of first pores positioned between adjacent tows. The method includes applying a plurality of first ceramic particles within the plurality of first pores. Next, the method includes applying a plurality of second ceramic fiber plies onto an outer surface of the plurality of first ceramic fiber plies. The second ceramic fiber plies include a plurality of interconnected tows and a plurality of second pores positioned between adjacent tows. The method then includes applying a plurality of second ceramic particles within the plurality of second pores. Further, the plurality of second ceramic particles are larger than the plurality of first ceramic particles. Lastly, the method includes densifying the ceramic matrix composite preform to form the ceramic matrix composite component.

A composite layer system is presented. The composite layer system includes a metallic substrate, a structured surface, and a thermal protection system. The structured surface may be additively manufactured onto the metallic substrate and includes structured surface features formed to project above the metallic substrate. Each of the structured surface features are separated from adjacent structured surface features by grooves. The thermal protection coating may be thermally sprayed onto the structured surface and is bonded to each of the structured surface features.

The invention relates to a component of a gas turbine engine which has a fiber-composite material having a multiplicity of rovings, wherein at least one roving of the fiber-composite material along the spatial extent thereof has a variable cross section, at least two rovings have in each case a dissimilar cross section and/or the at least one roving follows a variable direction and the at least one roving is deposited by means of tailored fiber placement. The invention also relates to a method of a component.

A turbine vane assembly adapted for use in a gas turbine engine includes a plurality of turbine vanes, an outer vane support, and an inner vane support. The plurality of turbine vanes comprise ceramic matrix composite material and are adapted to interact with hot gases flowing through a gas path of the gas turbine engine during use of the turbine vane assembly. A spar system couples the inner vane support with the outer vane support.

A turbine shroud assembly for a gas turbine engine includes a shroud wall extending circumferentially partway around a central reference axis to define a gas path of the gas turbine engine. An attachment feature extends radially from the shroud wall. A foam is located at least on the shroud wall.

An airfoil assembly for a gas turbine engine includes an airfoil extending radially relative to a central reference axis. A spar is located within the airfoil and spaced from the airfoil at all radial locations along the airfoil such that a gap is maintained between the airfoil and the spar. A foam is located between the airfoil and the spar.