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.

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.

A tip for a turbine component, turbine rotor blade or tip of a blade, is disclosed. The tip includes a tip plate configured to be coupled at a tip end of an airfoil chamber; and a tip rail extending radially from the tip plate, the tip rail disposed near or at a periphery of the tip plate. The trip rail includes a cooling structure constituting at least a portion of the tip rail. The cooling structure is in fluid communication with the airfoil chamber, and includes a plurality of repeating, three dimensional unit cells. Each unit cell defines a flow passage that is in fluid communication with the flow passage of at least one other unit cell. The flow passages of the 3D unit cells create a tortuous cooling passage in at least a portion of the tip rail.

An internally cooled component for a gas turbine engine includes a component having one or more exterior walls defining an internal component cavity configured for a cooling airflow to flow therethrough. An internal component rib extends into the internal component cavity from the one or more exterior walls. An insert is positioned in the internal component cavity, and a flow discourager is positioned at the insert and is configured to prevent the cooling airflow from flowing past the internal component 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.

One aspect of the disclosure provides for a turbine airfoil. The turbine airfoil may include a trailing edge having: a set of cooling channels having a first cooling channel fluidly connected to a second cooling channel; a first section having a first pin bank cooling arrangement, the first section fluidly connected to the first cooling channel; a second section having a second pin bank cooling arrangement, the second section fluidly connected to the second cooling channel and being radially inward of the first section; and a pressure side panel having a third pin bank cooling arrangement, the pressure side panel fluidly connected to the first cooling channel.

An apparatus and method for cooling an engine component such as a turbine engine airfoil, including a wall bounding an interior extending axially between a leading edge and a trailing edge and radially between a root and a tip. A cooling circuit it located within the interior of the airfoil can include a porous section having a porosity permitting a volume of fluid, such as air, to pass through the porous section.

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.

The invention relates to a thrust chamber device comprising a thrust chamber with a thrust space having a first portion, a second portion adjacent thereto, and a third portion adjacent to the second portion, the thrust space being delimited in all three portions by an outer nozzle wall having an outer thrust space surface, which outer thrust space surface tapers in the first and second portion toward the third portion, widens in the third portion away from the second portion, and has a narrowest point at the transition from the second portion to the third portion, the first portion being delimited by an inner nozzle wall with an inner thrust space surface, which tapers toward the second portion, an annular combustion chamber being formed between the inner thrust space surface and the outer thrust space surface and extending over the first portion.

An apparatus and method for cooling an engine component, such as an airfoil, including a wall to separate a hot flow from a cooling fluid flow. The component can include at least one trench disposed in a hot surface. The trench can be fed with the cooling fluid flow to cool the engine component along the hot surface with the cooling fluid flow.