A method of manufacturing a component for a gas turbine engine includes applying a thermoplastic polymer sheet over a composite body for the component; applying a shield over part of the composite body, the shield terminating at an end which overlies the thermoplastic polymer sheet and defines an interface between shielded and unshielded regions of the component; and pressing the shield into the thermoplastic polymer sheet so that the thermoplastic polymer sheet deforms around the end of the shield, such that the exterior profile of the component at the interface between the shielded and unshielded regions is flush.

An airfoil includes an airfoil section defining suction and pressure sides and leading and a trailing ends, the airfoil section being formed of a fiber-reinforced composite comprised of fiber plies, the fiber plies including at least one core fiber ply defining a tube that circumscribes an internal cavity and an overwrap fiber ply that wraps around the at least one core fiber ply, and aft of the internal cavity the overwrap fiber ply on the pressure side including first fingers and the overwrap ply on the suction side including second fingers, the first fingers being interdigitated with the second fingers to form an interlocked joint in the trailing end.

A turbine vane assembly adapted for use in a gas turbine engine includes a spar, a turbine vane, and load transfer pins. The spar comprises metallic materials and is configured to support other components of the turbine vane assembly relative to an associated turbine case. The turbine vane comprises ceramic matrix composite materials and is shaped to include an airfoil configured to direct the flow of hot gasses through a primary gas path of the turbine vane assembly.

In one example, a method for forming an environmental barrier coating (EBC) and/or abradable coating on a substrate. The method may include depositing a coating on a ceramic or ceramic matrix composite (CMC) substrate to form an as-deposited coating, wherein the coating includes at least one of an environmental barrier coating (EBC) and an abradable coating. The method further comprises heat treating the as-deposited coating at or above a first temperature for a first period of time following the deposition of the as-deposited coating on the substrate, wherein heat treating the as-deposited coating includes heating the as-deposited coating to or above the first temperature at a controlled rate. The heat treatment may be configured to at least one of decrease the open pores and/or microcracks of the heat-treated coating compared to the as-deposited coating or control a grain size of the heat-treated coating.

A method for manufacturing a composite part reinforced with nanotubes, includes stacking a plurality of composite plies of prepreg and at least one composite ply integrating nanotubes, the at least one composite ply integrating nanotubes being positioned in an inter-ply space between two composite plies of prepreg, wherein the at least one composite ply integrating nanotubes is a ply having a thermoplastic component, the nanotubes being integrated in the thermoplastic component.

An airfoil assembly includes first and second fiber-reinforced composite airfoil rings that each have inner and outer platform sections, a suction side wall extending between the inner and outer platforms, a pressure side wall extending between the inner and outer platforms, and suction and pressure side mate faces along, respectively, edges of the suction and pressure side walls. The suction side mate face of the first fiber-reinforced composite airfoil ring and the pressure side mate face of the second fiber-reinforced composite airfoil ring mate at an interface to form an airfoil that circumscribes an internal cavity. A least one of the suction or pressure side mate faces includes protrusions along a trailing edge of the airfoil. The protrusions define a toothed exit slot for emitting cooling air from the internal cavity.

A fan platform for gas turbine engine is provided. The fan platform incudes a body portion and a flow path surface coupled to the body portion. The body portion and the flow path surface define at least a portion of a flow path extending through the engine. The body portion and/or the flow path surface include an impact region including hybrid composite plies including one or more metallic tows. A gas turbine engine including the fan platform and methods for forming the fan platform are also disclosed.

A gas turbine engine component includes a ceramic matrix composite (CMC) body that includes an interior surface that defines a blind cavity. A barrier coating is disposed on an exterior of the CMC body. A ceramic body in the blind cavity lines at least a portion of the interior surface of the CMC body. The ceramic body defines an open, interconnected network of pores.

A method to produce a ceramic matrix composite with controlled surface characteristics includes: applying a scrim ply to a surface of a fiber preform, where the fiber preform includes silicon carbide fibers coated with boron nitride; infiltrating the fiber preform and the scrim ply with a slurry, thereby forming an impregnated ply on an impregnated fiber preform; infiltrating the impregnated fiber preform and the impregnated ply with a melt comprising silicon, and then cooling, thereby forming a ceramic matrix composite having a ceramic surface layer thereon, where the ceramic surface layer has a predetermined thickness and is devoid of boron; machining or grit blasting the ceramic surface layer to form an intermediate layer suitable for coating; and depositing an environmental barrier coating on the intermediate layer. Thus, a ceramic matrix composite coated with the environmental barrier coating is formed with the intermediate layer in between.

A turbomachine turbine nozzle extending around a central axis, including at least one radially outer shroud, at least one radially inner shroud, and at least one blade made of ceramic matrix composite material, distinct from the radially inner shroud and from the radially outer shroud, and extending radially between the radially inner shroud and the radially outer shroud, the blade being hollow and including a cavity opening at a radially inner end and at a radially outer end of the blade, the nozzle including at least one tubular mast arranged in the cavity of the blade and allowing routing the ventilation air passing through the cavity of the blade, the mast including a radially outer end attached to the radially outer shroud, and a radially inner end cooperating with a radial flange for positioning the radially inner shroud.