A method for processing a CMC airfoil includes nesting an airfoil fiber preform in a cavity of a fixture that has first and second tool segments, closing the fixture by rotating a first tool segment about a hinge, the closing causing the tool segments to clamp on a tail portion of the fiber preform and thereby conform the tail portion to the fixture. While in the fixture, the fiber preform is then partially densified with an interface coating material to form a partially densified fiber preform. While still in the fixture, one or more cooling holes are drilled into the trailing edge of the partially densified fiber preform. After the drilling, the partially densified fiber preform is removed from the fixture and further densified with a ceramic matrix material to form a fully densified CMC airfoil.

A tooling fixture assembly for processing a component of a gas turbine engine is provided. The tooling fixture assembly includes a component nesting fixture for precisely positioning the component within the fixture using a plurality of datum locators. A transfer block is also positioned at a known location and orientation relative to the component and an adhesive is used to fix the relative position of the component and the transfer block. The component and the transfer block are then removed from the component nesting fixture and installed on a machine for drilling, machining, or otherwise processing the component. After the machining is complete, the adhesive may be removed, such that the component may be detached from the transfer block.

Methods for forming a hole in a coated component are provided. The method may include forming a sacrificial layer over a ceramic barrier coating of a substrate, drilling a hole into the coated component such that any spatter formed during drilling deposits onto the sacrificial layer, and removing the sacrificial layer along with the spatter deposited thereon. The sacrificial layer may include a rare earth oxide (e.g., rare earth oxide particles). Intermediate ceramic matrix composite (CMC) component are also provided. The intermediate CMC may include a CMC body, an environmental barrier coating on the bond coating, and a sacrificial layer on the environmental barrier coating, with the sacrificial layer including particles of a rare earth oxide dispersed in a polymeric matrix.

Introducing a through hole into a substrate (4) before coating, and performing the removal thereafter, shortens the machining times for producing a through hole (18) with a diffuser (13) and also subjects the intermediate layers to less stress.

The invention relates to a seal (10) for sealing a gap between a stationary component and a moving component, in particular for sealing a radial gap between a rotor and a stator of a turbomachine, comprising at least one sealing segment (12) with an edge zone (14) facing the gap, whereby the seal (10) is produced layer-by-layer by a free-forming method, in particular a generative or additive method. A plurality of pre-defined weak regions (16) is formed in the edge zone (14) of the sealing segment (12). In addition, the invention relates to a method for producing a seal (10) as well as a turbomachine.

A method for coating a turbine engine component, said method includes the steps of: placing the component into a chamber; injecting a non-reactive carrier gas containing a coating material into the chamber; and forming a coating on a desired portion of the component by locally heating the desired portion of the component by redirecting a directed energy beam onto the desired portion of the component.

In a rotary assembly for an exhaust gas turbocharger, with a turbine wheel, a compressor wheel which is non-rotatably connected with the turbine wheel by means of a shaft, and an intermediate component of the rotary assembly, which is arranged on the shaft between the compressor wheel and the turbine wheel, at least one of the surfaces of the intermediate component and at least one of a third surface of the compressor wheel and a fourth surface of the shaft comprises projections which are harder than a base of the surface arranged on the opposite side with which they are engaged thereby to increase a permissible operating torque of the exhaust gas turbocharger.

Methods for preparing slotted ceramic coatings and the resulting components comprising the same are provided. The methods and products include the incorporation of a coating system comprising a ceramic coating with cooling holes disposed throughout the ceramic coating and slots defined in the thermal barrier coating and disposed in relation to the cooling holes. The resulting ceramic coating has improved resistance to CMAS infiltration and improved compliance resulting in an increased life of the coated component.

A coated ceramic matrix composite or metallic component and a gas turbine assembly is provided. The component comprises a substrate comprising a first surface and a hot gas path surface. The hot gas path surface is arranged and disposed to contact a hot gas flow when the component is installed in the gas turbine. The first surface is disposed at an angle to the hot gas path surface and opposes at least one adjacent component when the component is installed in the gas turbine. The component further comprises an angled or rounded feature extending from the first surface to the hot gas path surface. The component further comprises an environmental barrier coating or thermal barrier coating on at least a portion of the hot gas path surface. The angled or rounded feature reduces an incidence angle of the hot gas flow onto the first surface. The gas turbine assembly comprises a plurality of the coated components.

A gas turbine engine blade includes a platform, an airfoil section extending from the platform in a first direction, a mount extending from the platform in a second direction opposite the first direction, and a damper deposited on one of the platform, the airfoil section, and the mount. A method of sizing a damper for a gas turbine engine blade is also disclosed.