Provided is a gas turbine vane and blade assembly in which lattice structures are installed between an impingement plate and an effusion plate. The gas turbine vane and blade assembly is capable of enhancing cooling efficiency in an impingement/effusion cooling technique.

In addition, the gas turbine vane and blade assembly can be manufactured using an additive manufacturing technique, and the lattice structures are capable of replacing supports that are used during an additive manufacturing process, and improving not only structural rigidity and stability but also cooling performance.

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 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.

An apparatus and method regarding an airfoil for a turbine engine, the airfoil comprising an outer wall defining an interior bound by a pressure side and a suction side extending axially between a leading edge and a trailing edge defining a chord-wise direction and extending radially between a root and a tip defining a span-wise direction. The airfoil further comprising at least one cooling passage extending radially within the interior and defining a primary cooling airflow; and at least one cooling hole having an inlet defining a first cross-sectional area, the inlet in communication with the cooling passage, an outlet defining a second cross-sectional area greater than the first cross-sectional area; wherein the primary cooling airflow enters the inlet in a first direction and exits the outlet in a second direction different than the first direction.

A manufacturing method is provided. During this method, a preform component is provided for a turbine engine. The preform component includes a substrate. A meter section of a cooling aperture is formed in the substrate. An internal coating is applied onto a surface of the meter section. An external coating is applied over the substrate. A diffuser section of the cooling aperture is formed in the external coating and the substrate to provide the cooling aperture.

A gas turbine engine article includes a silicon-containing ceramic wall that has an external combustion gaspath side and an internal side that borders a cooling air cavity. The external combustion gaspath side has an associated combustion gas flow direction there along. An array of cooling holes extends through the silicon-containing ceramic wall and connects the internal side with the external combustion gaspath side. The cooling holes are oriented to discharge cooling air to the external gaspath side in a direction counter to the combustion gas flow direction.

A turbine component includes a body having a pair of spaced walls, with at least one of the walls for facing a fluid flow when mounted in a gas turbine engine. There are a plurality of wall cooling passages having a generally boomerang shape such that a peak apex is spaced from the wall and an indent apex is adjacent to the wall, with the plurality of wall cooling passages having interior sides extending from the peak apex toward the wall to define a corner. Outer sides extend from the corners with a component away from the wall and to the indent apex. A gas turbine engine is also disclosed.

A manufacturing method is provided during which a preform component for a turbine engine is provided. A cooling aperture is formed in the preform component. The cooling aperture includes a centerline, an inlet and an outlet. The cooling aperture extends longitudinally along the centerline through a wall of the preform component from the inlet to the outlet. The forming of the cooling aperture includes forming a first portion of the cooling aperture using a machining tool implement with a first toolpath that is angularly offset from the centerline by a first angle between thirty-five degrees and ninety degrees.

A turbine blade having cooling holes formed therein and a turbine including the same are provided. The turbine blade includes an airfoil having a leading edge and a trailing edge formed thereon and a cooling passage defined for flow of a cooling fluid therethrough, and a cooling hole configured to communicate between the cooling passage and outside in the airfoil and having an inlet and an outlet, wherein the cooling hole includes an expanded portion and a grooved portion formed in the outlet, the grooved portion being recessed from the expanded portion toward the trailing edge.

A combustor component according to at least one embodiment of the present invention includes a cylindrical body which internally includes a combustion chamber, and includes a weld part where a plurality of through holes opening to the combustion chamber are formed, and a housing which is disposed on an outer circumferential side of the cylindrical body to cover a part of the weld part, and defines an acoustic damping space communicating with the combustion chamber via at least one of the through holes. The plurality of through holes in the weld part has a formation density which is higher in a first region of the weld part covered with the housing than in a second region of the weld part positioned outside the housing.