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.

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.

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 damper seal for a gas turbine engine includes a damper body extending in a first direction between a leading edge portion and a trailing edge portion, extending in a second direction between first and second sidewalls, and extending in a third direction between a convex outer damper face and a concave inner damper face. The inner damper face establishes a damper pocket. The leading and trailing edge portions slope inwardly from opposite ends of the damper body to bound the damper pocket in the first direction. The first and second sidewalls extend from the leading edge portion to the trailing edge portion and slope inwardly from opposite sides of the damper body to bound the damper pocket in the second direction. The outer damper face is pre-formed according to a first predetermined geometry that substantially corresponds to a second predetermined geometry of a platform undersurface bounding a neck pocket of an airfoil. A method of damping for a gas turbine engine is also disclosed.

A method for forming a component for a gas turbine engine may include forming a first portion of the component that includes a cast metal or metal alloy, forming a second portion of the component that includes presintered preform defining at least one support structure, positioning the second portion on the first portion to define an assembly such that the first portion and the second portion define at least one cooling channel therebetween, and heating the assembly to join the first portion and the second portion and form the component.

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 turbine blade for a gas turbine engine. The turbine blade includes an airfoil having a tip cavity. The tip cavity has a floor bounded by a wall. A pocket opening is formed in the wall proximate a trailing edge of the turbine blade. A passageway communicates cooling medium from an internal cooling circuit to a cooling hole formed in part through the floor and in part through the trailing edge.

A turbine blade is provided with an airfoil portion having a leading edge, a trailing edge, and a pressure surface and a suction surface which extend between the leading edge and the trailing edge, the airfoil portion internally forming a cooling passage. The cooling passage includes: a first cooling passage located closer to the pressure surface than the suction surface; and a second cooling passage located closer to the suction surface than the pressure surface. The first cooling passage and the second cooling passage are separated by a partition member disposed in the airfoil portion. At least one communication space connecting the first cooling passage and the second cooling passage is formed in the partition member.

A method of forming a ceramic matrix composite component with cooling channels includes embedding a plurality of wires into a preform structure, densifying the preform structure with embedded wires, and removing the plurality of wires to create a plurality of corresponding channels within the densified structure.

An endwall assembly disposed at one end of a vane assembly may comprise an endwall spar that includes an cold side, an hot side, a leading edge, a trailing edge, and an axis extending from the leading edge to the trailing edge perpendicular to the leading edge. The endwall assembly may include a coversheet on the hot side of the endwall spar and a cooling channel that includes a cooling fluid inlet disposed in the endwall spar, and a cooling fluid outlet. The endwall assembly may include a structure protruding from the cold side of the endwall spar, wherein the structure is located between the cooling fluid inlet and the cooling fluid outlet along the axis; and a cooling fluid source cavity on the cold side of the wall, the cooling fluid source cavity in fluid communication with the cooling fluid channel via the cooling fluid inlet.