A cooling device for a gas turbine engine component comprises a gas turbine engine component having an upstream channel and a downstream channel that define a cooling flow path. A meter feature includes at least one hole to meter flow from the upstream channel to the downstream channel, and has an upstream side and a downstream side. An exit diffuser extends outwardly from the downstream side of the meter feature to control flow in a desired direction into the downstream channel. A gas turbine engine is also disclosed.

An augmentor vane assembly of a gas turbine engine with an additively manufactured augmentor vane, the additively manufactured augmentor vane having a vane wall with integral longitudinal wall passages formed therein, at least one of the integral longitudinal wall passages including at least one entrance aperture and at least one exit aperture, the at least one exit aperture transverse to the integral longitudinal wall passage.

A cooled gas turbine engine component includes a wall having a plurality of effusion cooling apertures extending there-through from a first surface to a second surface. Each aperture has an inlet in the first surface and an outlet in the second surface. Each aperture includes an inlet portion, a collection chamber, a metering portion, a U-shaped bend portion and a diffusing portion arranged in flow series from the inlet to the outlet. The inlet portion of each aperture is arranged substantially perpendicularly to a surface of the collection chamber. The metering portion of each aperture is arranged to extend longitudinally from a first lateral side of the collection chamber and the diffusing portion of each aperture is arranged at an angle to the second surface. Each outlet has a quadrilateral shape in the second surface of the wall and each outlet is displaced laterally from the metering portion.

An aviation turbine blade extending in the radial direction and presenting a pressure side and a suction side, including a plurality of pressure side cavities extending radially at the pressure side of the blade, a plurality of suction side cavities extending radially at the suction side of the blade, and at least one central cavity located in the central portion of the blade and surrounded by pressure side cavities and by suction side cavities, the blade further including a plurality of cooling circuits, in which at least a first cooling circuit comprises: a first cavity and a second cavity, the first and second cavities communicating with each other at a radially inner end and at a radially outer end of the blade.

A method according for the manufacturing of a component in which walls surround a cavity and the cavity is accessible through at least one aperture formed in one of the walls, according to the following steps: manufacturing of the component by an additive method, in which metallic powder particles are applied to a support layer by layer in a process chamber, and the walls are each manufactured after the application of a layer of the metallic powder particles by melting by means of an energy beam along a predetermined path, connection of the aperture to a flushing device, supply of a liquid etchant into the cavity by means of the flushing device, selective dissolution by the etchant of power particles connected to each other only via sinter necks and/or fusible links and flushing of the etchant and the dissolved powder particles out of the cavity.

A turbine shroud segment comprises a body having an upstream end portion and a downstream end portion relative to a flow of gases through the gas path. A core cavity is defined in the body and extends axially from the upstream end portion to the downstream end portion. A plurality of cooling inlets is defined in the upstream end portion of the body for feeding coolant in the core cavity. A plurality of cooling outlets is defined in the downstream end portion of the body for discharging coolant from the core cavity. Pedestals are provided in the core cavity.

A turbine shroud segment has a body extending axially between a leading edge and a trailing edge and circumferentially between a first and a second lateral edge. A core cavity is defined in the body and extends axially from a front end adjacent the leading edge to a rear end adjacent to the trailing edge. A plurality of cooling inlets and outlets are respectively provided along the front end and the rear end of the core cavity. A crossover wall extends across the core cavity and defines a row of crossover holes configured to accelerate the flow of coolant directed into the core cavity via the cooling inlets. The crossover wall is positioned to accelerate the coolant flow at the beginning of the cooling scheme where the shroud segment is the most thermally solicited.

Disclosed is an aero-engine turbine assembly, including a turbine assembly body and a cooling component. The turbine assembly body is provided with an internal flow passage, and the turbine assembly body includes a turbine rotor disk, a blade end wall and a turbine rotor blade, which are successively fixedly connected with each other. The internal flow passage passes through the turbine rotor disk, the blade end wall and the turbine motor blade, and the internal flow passage is provided with an inlet and an outlet. The cooling component is fixed on the turbine rotor disk, and the cooling component includes an electromagnetic pump system, an expansion joint and a radiator, which are successively communicated with each other. The electromagnetic pump system is communicated with the inlet, to inject liquid metal to into the internal flow passage.

A component for a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a body portion, a cooling circuit disposed within the body portion and including at least a first cavity, a core in fluid communication with the first cavity, and an exit surface that extends through an exterior surface of the body portion. At least one surface indicator is visible near the exit surface.

A turbine blade in an industrial gas turbine includes a blade surface to be cooled by a film of cooling fluid, a plurality of cooling holes on the blade surface through which cooling fluid flows, each cooling hole including an inlet portion and an outlet portion, and a trench on the blade surface surrounding at least one outlet portion of the cooling hole, the trench extending in an axial direction and a radial direction from the outlet portion of the cooling hole, wherein the outlet portion of the cooling hole has a shape configured to generate a first stage diffusion of the cooling fluid and a wall of the trench is positioned in the axial direction from the outlet portion of the cooling hole to generate a second stage diffusion of the cooling fluid, thereby forming the film of cooling fluid.