Abstract: An airfoil includes a platform that has platform leading and trailing ends, lateral side faces, and inner and outer faces. An airfoil portion extends outwardly from the inner face of the platform. The airfoil portion includes airfoil leading and trailing ends and side walls that join the airfoil leading and trailing ends. The platform includes a cooling passage that has an inlet at a forward location, outlet slots at the platform trailing end, and an intermediate passage portion that extends from the inlet to the outlet slots. The intermediate passage portion includes a common manifold region that feeds the outlet slots. The platform includes a rib that is elongated in a length direction between the platform leading and trailing ends. The rib divides two of the cooling passages such that the two cooling passages are fluidly unconnected with each other in the platform.

Abstract: An impingement cooled component includes a first wall having a plurality of impingement holes and a second wall spaced apart from the first wall. The second wall is downstream of the first wall, relative to a cooling flow, the second wall has a contoured surface facing the first wall. The contoured surface includes a plurality of contours defined by at least one of a plurality of peaks and a plurality of valleys, and at least one of the contours in the plurality of contours is aligned with an axis defined by one of the impingement holes in the plurality of impingement holes.

Abstract: An impingement cooled component includes a first wall having a plurality of impingement holes and a second wall spaced apart from the first wall. The second wall is downstream of the first wall, relative to a cooling flow, the second wall has a contoured surface facing the first wall. The contoured surface includes a plurality of contours defined by at least one of a plurality of peaks and a plurality of valleys, and at least one of the contours in the plurality of contours is aligned with an axis defined by one of the impingement holes in the plurality of impingement holes.

Abstract: An airfoil for a gas turbine engine includes pressure and suction side walls joined to one another at leading and trailing edges. The pressure and suction side walls surround an airfoil cavity and provide an exterior airfoil surface. A baffle is arranged in the airfoil cavity and includes a supply hole. Ribs extend between at least one of the pressure and suction side walls into the airfoil cavity to support the baffle relative to the at least one of the pressure and suction side walls. The ribs are configured to provide a serpentine cooling passage between the baffle and at least one of the pressure and suction side walls. The serpentine cooling passage has first and second passes joined by a bend. The ribs form a film cooling cavity between the first and second passes. The supply hole fluidly connects the baffle to the film cooling cavity. Film cooling holes extend through the at least one of the pressure and suction side walls. The film cooling holes are in fluid communication with the film cooling cavity.

Abstract: An assembly for a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, an airfoil including a radial end, a first passageway having an outlet at the radial end, and a second passageway having an inlet at the radial end. The assembly further includes a cover having at least one turning cavity configured to direct fluid expelled from the outlet of the first passageway into the inlet of the second passageway.

Abstract: A flowpath component for a gas turbine engine includes a leading edge, a trailing edge connected to the leading edge via a first surface and a second surface, an impingement cavity internal to the flowpath component, the impingement cavity being aligned with one of the leading edge and the trailing edge, a cooling passage extending at least partially through the flowpath component, and a plurality of crossover holes connecting the cooling passage to the impingement cavity. At least one of the crossover holes is aligned normal to an expected direction of fluid flow through the cooling passage and is unaligned with an axial line drawn perpendicular to a stacking line of the flowpath component.

Abstract: An airfoil for a gas turbine engine includes an airfoil including spaced apart pressure and suction side walls joined at leading and trailing edges to provide an exterior airfoil surface that extends in a radial direction from a platform to a tip. A cavity is provided between the pressure and suction side walls near the trailing edge. The cavity includes an interior region bounded by first and second exit regions arranged at angle relative to one another. The first and second exit regions are respectively in low and high pressure regions relative to one another. First and second pedestal groups respectively are arranged at the first and second exit regions. The second pedestal group has first and second pedestals each terminating in an end. The ends of the second pedestals extend beyond the ends of first pedestals.

Abstract: A gas turbine engine component has an airfoil extending radially inwardly of an outer platform to an inner platform. A central passage is formed within the airfoil and has an inlet end for receiving cooling air. An outlet end in the inner platform delivers cooling air to a downstream use. The airfoil has a suction wall and a pressure wall, and extends in an axial direction from a leading edge to a trailing edge. A suction skin core is between the central passage and the suction wall. A pressure skin core is between the central passage and the pressure wall.

Abstract: a gas turbine engine structure includes a body radially spanning an inner portion and an outer portion. The structure body includes a first passage interior to the structure body. The first passage includes a first opening on one of a radially outward edge and a radially inward edge of the structure, and a second opening on the other of the radially outward edge and the radially inward edge of the structure. A second passage is also included interior to the structure. The second passage is approximately aligned with the first passage, and is disposed between the first passage and one of the first surface and the second surface of the structure. The second passage insulates the first passage from heat transfer through one of the first surface and the second surface.

Abstract: A film cooled component may comprise a cooling chamber and a first ligament centered about a first axis. The first ligament may be in fluid communication with the cooling chamber. A first meter may be disposed at an end of the first ligament. A first diffuser may extend from the first meter to a surface of the film cooled component. The first diffuser may comprise a first tapered sidewall oriented at an angle of between 5 degrees to 15 degrees relative to the first axis. The first diffuser may further comprise a first non-tapered sidewall oriented at an angle less than 5 degrees relative to the first axis.

Abstract: According to one aspect of the present disclosure, an engine component for a gas turbine engine is disclosed that includes an internal cavity and a baffle. The internal cavity is defined by the engine component, and the internal cavity is in fluid communication with a source of cooling air. The baffle is disposed within the internal cavity, and includes at least four longitudinal sidewalls that include at least two first sidewalls, one of which is upstream and one of which is downstream, and at least two second sidewalls, at least one of which faces towards a core flow path of the gas turbine engine. The baffle includes a plurality of openings for directing cooling air from within the baffle into the cavity. A majority of the openings are on at least one of the first sidewalls.

Abstract: An airfoil includes a platform that has platform leading and trailing ends, lateral side faces, and inner and outer faces. An airfoil portion extends outwardly from the inner face of the platform. The airfoil portion includes airfoil leading and trailing ends and side walls that join the airfoil leading and trailing ends. The platform includes a cooling passage that has an inlet at a forward location, outlet slots at the platform trailing end, and an intermediate passage portion that extends from the inlet to the outlet slots. The intermediate passage portion includes a common manifold region that feeds the outlet slots. The platform includes a rib that is elongated in a length direction between the platform leading and trailing ends. The rib divides two of the cooling passages such that the two cooling passages are fluidly unconnected with each other in the platform.

Abstract: An airfoil includes a platform that has platform leading and trailing ends, lateral side faces, and inner and outer faces. An airfoil portion extends outwardly from the inner face of the platform. The airfoil portion includes airfoil leading and trailing ends and side walls that join the airfoil leading and trailing ends. The platform includes a cooling passage that has an inlet at a forward location, outlet slots at the platform trailing end, and an intermediate passage portion that extends from the inlet to the outlet slots. The intermediate passage portion includes a common manifold region that feeds the outlet slots.

Abstract: What is described is a transfer tube for use with an airfoil of a gas turbine engine coupled to a platform. The transfer tube includes a main body having a first end defining an inlet configured to receive a flow of fluid and a second end defining an outlet for the flow of fluid, the main body having a curved section. The transfer tube also includes a first mating face coupled to the first end of the main body. The transfer tube also includes a second mating face coupled to the second end of the main body. At least one of the first mating face or the second mating face is configured to be coupled to a platform body of the outer diameter platform.

Abstract: A heat shield is provided for use in an inner cavity of a gas turbine engine component. The heat shield comprises a heat shield body comprising opposed side portions defining an inner channel. An inner section of a side portion of the opposed side portions is inwardly sloped such that a cross-section of the heat shield body is smaller in an inner body portion relative to a cross-section of an outer body portion thereof. An aperture is disposed through the side portion in a selected radial position. A vane assembly including the heat shield is also provided.

Abstract: A flow path component for a gas powered turbine includes a flow path component body having cross sectional profile having a leading edge and a trailing edge. The leading edge is connected to the trailing edge by a first side and by a second side opposite the first side. A serpentine cooling passage includes a plurality of segments, each of the segments being generally radially aligned. A first subset of the segments is disposed along one of the first side and the second side of the cross sectional profile, and a second subset of the segments spans the flow path component body from the first side to the second side. A cavity is positioned internal to the flow path component body. The cavity is at least partially shielded from one of the first side and the second side by at least one of the plurality of segments.

Abstract: A flow splitting baffle for separating a main cooling flow through an inner channel of a component includes a tubular structure defining a tubular cavity and having a longitudinal axis. The flow splitting baffle also includes an inner ring coupled to the tubular structure and extending away from the tubular structure along a plane that is perpendicular to the longitudinal axis. The flow splitting baffle also includes an outer ring parallel to the plane, positioned a first distance from the tubular structure, extending away from the tubular structure and positioned a second distance from the inner ring. The flow splitting baffle also includes a strut perpendicular to the plane, extending from the inner ring to the outer ring and coupled to the inner ring and the outer ring.

Abstract: An assembly for a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, an airfoil including a radial end, a first passageway having an outlet at the radial end, and a second passageway having an inlet at the radial end. The assembly further includes a cover having at least one turning cavity configured to direct fluid expelled from the outlet of the first passageway into the inlet of the second passageway.

Abstract: An airfoil for a gas turbine engine includes pressure and suction side walls joined to one another at leading and trailing edges. The pressure and suction side walls surround an airfoil cavity and provide an exterior airfoil surface. A baffle is arranged in the airfoil cavity and includes a supply hole. Ribs extend between at least one of the pressure and suction side walls into the airfoil cavity to support the baffle relative to the at least one of the pressure and suction side walls. The ribs are configured to provide a serpentine cooling passage between the baffle and at least one of the pressure and suction side walls. The serpentine cooling passage has first and second passes joined by a bend. The ribs form a film cooling cavity between the first and second passes. The supply hole fluidly connects the baffle to the film cooling cavity. Film cooling holes extend through the at least one of the pressure and suction side walls. The film cooling holes are in fluid communication with the film cooling cavity.

Abstract: An assembly for a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, an airfoil including a radial end, a first passageway having an outlet at the radial end, and a second passageway having an inlet at the radial end. The assembly further includes a cover having at least one turning cavity configured to direct fluid expelled from the outlet of the first passageway into the inlet of the second passageway.