A turbomachine component, particularly a gas turbine engine component, has at least one part built in parts from a curved or planar panel, particularly a sheet metal, the part having a plurality of cooling channels via which a cooling fluid, particularly air, is guidable, wherein at least one of the plurality of cooling channels has a continuously tapered section. The at least one of the plurality of cooling channels has a single inlet port from a first surface of the panel and a single outlet port for the cooling fluid to another surface, particularly a surface opposite to the first surface, or to the first surface. Further the panel is built via laser sintering or laser melting or direct laser deposition. A gas turbine engine is equipped with such a component. A method of manufacturing includes incorporating cooling channels having a continuously tapered section.

The present invention relates to a ceramic or metallic component including a first region having a first porosity ranging between 1 and 30%. The component includes a second region having a second porosity that is less than the first porosity. The component includes at least one graded transition between the first and second regions.

Provided is a combustor assembly. The combust assembly includes: a plurality of swirlers through which a first fluid that is a part of a fluid discharged from a compressor passes; a base portion, in which the plurality of swirlers are provided, comprising a first through hole formed between one swirler and another swirler from among the plurality of swirlers so that a second fluid that is another part of the fluid discharged from the compressor and different from the first fluid passes through the first through hole; and a deflector provided in the base portion so as to face the first through hole for changing a moving direction of the second fluid.

An apparatus for a gas turbine engine can include an airfoil having an interior. The interior can be separated into one or more cooling air channels extending in a span-wise direction. An accelerator insert can be placed in one or more cooling air channels to define a reduced cross-sectional area within the cooling air channel to accelerate an airflow passing through the cooling air channel.

An apparatus and method of cooling an airfoil for a gas turbine engine includes a tip for the radially outer end of the airfoil with internal ribs defining cooling circuits within an interior of the airfoil. The ribs can be full-length, extending between a root and tip of the airfoil. A gap can be formed in the full-length ribs near the tip to form a thermal stress reduction structure for the full-length rib.

An example gas turbine engine component includes a component configured to separate a cooling air plenum from a heated gas environment. The component includes a substrate defining a surface, and a unitary structure. The unitary structure includes a cooling region and a cover layer. The cover layer defines a hot wall surface configured to face the heated gas environment. The cooling region is disposed between the cover surface and the substrate and includes a plurality of support structures extending between the cover layer and the surface of the substrate. At least some of the support structures define a respective bond surface bonded to the substrate at the surface of the substrate. An example technique for fabricating the gas turbine engine component includes additively depositing the unitary structure on the surface of the substrate.

In one embodiment, a gas turbine engine component includes a foam based core and a composite skin member. Both the foam based core and the composite skin member can be used to structurally support the gas turbine engine component. The composite skin member can be a CMC material and is used to partially encapsulate the foam core. The gas turbine engine component can take the form of an airfoil member such as a blade or a vane, a combustor liner, etc. A first portion of the composite skin member includes a first surface extending past an edge of the component creating a step approximate an edge section. In another embodiment, composite skin members can be used to form a continuous shape for the edge section such that the foam core forms part of a gas path surface.

The present disclosure concerns a heat exchanger, which may for example be utilised in a gas turbine engine or in other applications. Example embodiments include a heat exchanger comprising: an external surface for exchanging heat with an external fluid flow passing over the external surface; a first fluid passage extending through the heat exchanger from a first fluid inlet to a first fluid outlet, a first portion of the first fluid passage extending along the heat exchanger adjacent to the external surface for a first cooling fluid passing through the first fluid passage to exchange heat with the external fluid flow; and a second fluid passage extending through the heat exchanger from a second fluid inlet to a second fluid outlet located at the external surface for a second cooling fluid to pass from the second fluid inlet into the external fluid flow.

An airfoil for a gas turbine engine is provided that includes a first portion formed from a first plurality of plies of a ceramic matrix composite material and defining an inner surface of the airfoil, as well as a second portion formed from a second plurality of plies of a ceramic matrix composite material and defining an outer surface of the airfoil. The first portion and the second portion define a non-line of sight cooling aperture extending from the inner surface to the outer surface of the airfoil. In one embodiment, a surface angle that is less than 45° is defined between a second aperture and the outer surface. A method for forming an airfoil for a gas turbine engine also is provided.

A component includes an outer wall that includes an exterior surface, and at least one plenum defined interiorly to the outer wall and configured to receive a cooling fluid therein. The component also includes a coating system disposed on the exterior surface. The coating system has a thickness. The component further includes a plurality of adaptive cooling openings defined in the outer wall. Each of the adaptive cooling openings extends from a first end inflow communication with the at least one plenum, outward through the exterior surface and to a second end covered underneath at least a portion of the thickness of the coating system.