Abstract: A hollow airfoil body includes a baffle that is arranged in the airfoil body that extends in a radial direction and provides a fluid flow direction. The baffle has multiple holes that include first and second holes that are configured to conduct the cooling airflow in a chordwise direction toward the trailing edge of the airfoil body. The airfoil has a first standoff that extends from the airfoil body to support the baffle. The first axial standoff has a first length that defines an axial passage including the multiple holes. The airfoil includes a second standoff that has a second length less than the first length. The second standoff is arranged radially between the first and second holes. The first hole is smaller than the second hole. The second hole is downstream from the first hole relative to the fluid flow direction.

Abstract: A casting core assembly for a gas turbine engine component according to an implementation includes a skin core corresponding to a first cooling passage of an airfoil. The first cooling passage includes a first section and a tip flag section joined at a first bend. The skin core includes a first portion corresponding to the first section and a tip flag portion corresponding to the tip flag section. The skin core includes at least one arcuate slot corresponding to at least one turning vane of the airfoil. A method of forming an airfoil for a gas turbine engine is also disclosed.

Abstract: A turbine blade includes a platform, a root section, and an airfoil section extending from the platform to a tip. The airfoil section includes a leading edge and a trailing edge extending from the platform to the tip. A tip wall is at the tip and extends from the leading edge to the trailing edge. A first core passage extends from the root section to the tip wall between the leading edge and the trailing edge. A first tip flag passage extends adjacent to the tip wall from the first core passage to a first flag outlet on the trailing edge. A second tip flag passage extends toward the leading edge from a second flag outlet on the trailing edge and is between the first tip flag passage and the root section. A second core passage extends from the root section to the second tip flag passage.

Abstract: Vane assemblies for gas turbine engines are described. The vane assemblies include a platform having an interior platform surface, a forward rail, and an aft rail, wherein the interior platform surface, the forward rail, and the aft rail define a plenum, an airfoil extending radially inward from the platform on a side opposite the forward and aft rails, the airfoil having a leading edge cavity and a baffle installed within the leading edge cavity, and platform feed structure arranged on the platform in the plenum and defining a fluid path through the forward rail and into the baffle of the leading edge cavity.

Abstract: Baffles for installation within airfoils include a baffle body defining a feed cavity and extending between inner and outer diameter ends. A forward standoff shelf is formed along an exterior surface of the baffle and defined by a depression, bend, or channel in a material of the baffle body extending between the inner and outer diameter ends. The forward standoff shelf is configured to engage with a forward rail of the airfoil body, and an aft standoff shelf is formed along an exterior surface of the baffle body and configured to engage with an aft rail of the airfoil body. A surface of the baffle body between the forward standoff shelf and the aft standoff shelf defines a side channel surface extending in a radial direction along the baffle body between the outer diameter end and the inner diameter end.

Abstract: A manufacturing method is provided. During this method, a preform component for a turbine engine is provided that includes a substrate. A meter section of a cooling aperture is formed in the substrate. An external coating is applied over the substrate. At least a portion of the substrate and the external coating is scanned with an imaging system to provide scan data indicative of an internal structure of the portion of the substrate and the external coating. A diffuser section of the cooling aperture is formed in the external coating and the substrate based on the scan data.

Abstract: A turbine blade includes a platform, a root section, and an airfoil section extending from the platform to a tip. The airfoil section includes a leading edge and a trailing edge extending from the platform to the tip. A tip wall is at the tip and extends from the leading edge to the trailing edge. A first core passage extends from the root section to the tip wall between the leading edge and the trailing edge. A first tip flag passage extends adjacent to the tip wall from the first core passage to a first flag outlet on the trailing edge. A second tip flag passage extends toward the leading edge from a second flag outlet on the trailing edge and is between the first tip flag passage and the root section. A second core passage extends from the root section to the second tip flag passage.

Abstract: A turbine vane assembly for a gas turbine engine is disclosed herein. The turbine vane assembly includes a turbine vane including a leading edge, a pressure edge, a suction edge, and a trailing edge, a core defined by the turbine vane, an outer platform end wall connected to the turbine vane, the outer platform end wall defining an interior space, the interior space being in fluid communication with the core, and a plurality of cooling holes formed in the turbine vane, the plurality of cooling holes being in fluid communication with the core.

Abstract: An airfoil for a gas turbine engine includes a first wall disposed in an interior cavity, the first wall extending in the spanwise direction from a base region to a tip wall and adjoining a pressure side wall and a suction side wall to form a first cooling channel; a second wall disposed in the interior cavity and extending in a chordwise direction from the first wall toward a trailing edge, the second wall adjoining the pressure side wall and the suction side wall to form a second cooling channel; a first hole through the first wall, the first hole connecting the first cooling channel to the second cooling channel; and a second hole though the second wall connecting the second cooling channel to a third cooling channel.

Abstract: An airfoil includes a platform and an airfoil section that extends from the platform. The airfoil defines leading and trailing ends and suction and pressure sides. The airfoil section has a transition region through which the airfoil section blends into the platform. The trailing end of the airfoil section has an axial cooling slot opening through the transition region and defines a circumferentially diverging ramp in the transition region.

Abstract: A manufacturing method is provided. During this method, a preform component for a turbine engine is provided that includes a substrate. A meter section of a cooling aperture is formed in the substrate. An external coating is applied over the substrate. At least a portion of the substrate and the external coating is scanned with an imaging system to provide scan data indicative of an internal structure of the portion of the substrate and the external coating. A diffuser section of the cooling aperture is formed in the external coating and the substrate based on the scan data.

Abstract: A turbine engine airfoil element has: a plurality of spanwise main body passageways along a camber line; and a plurality of spanwise skin passageways along the pressure side. The spanwise skin passageways include: first skin passageways each having a plurality of film cooling outlets to the pressure side; and second skin passageways each lacking film cooling outlets to the pressure side. Linking passageways extend between the first skin passageways and the second skin passageways. The first skin passageways and second skin passageways are directly fed from one or more inlets of the airfoil element.

Abstract: A casting core assembly for a gas turbine engine component according to an implementation includes a skin core corresponding to a first cooling passage of an airfoil. The first cooling passage includes a first section and a tip flag section joined at a first bend. The skin core includes a first portion corresponding to the first section and a tip flag portion corresponding to the tip flag section. The skin core includes at least one arcuate slot corresponding to at least one turning vane of the airfoil. A method of forming an airfoil for a gas turbine engine is also disclosed.

Abstract: A turbine engine airfoil element has: a plurality of spanwise main body passageways along a camber line; and a plurality of spanwise skin passageways along the pressure side. The spanwise skin passageways include: first skin passageways each having a plurality of film cooling outlets to the pressure side; and second skin passageways each lacking film cooling outlets to the pressure side. Linking passageways extend between the first skin passageways and the second skin passageways. The first skin passageways and second skin passageways are directly fed from one or more inlets of the airfoil element.

Abstract: A turbine vane assembly for a gas turbine engine is disclosed herein. The turbine vane assembly includes a turbine vane including a leading edge, a pressure edge, a suction edge, and a trailing edge, a core defined by the turbine vane, an outer platform end wall connected to the turbine vane, the outer platform end wall defining an interior space, the interior space being in fluid communication with the core, and a plurality of cooling holes formed in the turbine vane, the plurality of cooling holes being in fluid communication with the core.

Abstract: An airfoil for a gas turbine engine according to an example of the present disclosure includes an airfoil section extending in a radial direction from a root section to a tip portion. The airfoil section has an external wall and an internal wall. The airfoil section establishes an internal cooling arrangement including a first cooling passage having a first section and a tip flag section. The tip flag section extends in the chordwise direction along the tip portion from the first section. The first section includes a plurality of branched paths established by at least one turning vane.

Abstract: A method of preparing a casting article for use in manufacturing a gas turbine engine part according to an exemplary aspect of the present disclosure includes, among other things, communicating a powdered material to an additive manufacturing system and preparing a casting article that includes at least one trunk and a skin core that extends from the at least one trunk out of the powdered material. A casting article is also disclosed.

Abstract: An airfoil includes an airfoil section that has an airfoil wall that defines a leading end, an arced trailing end, and first and second sides that join the leading end and the arced trailing end. The first and second sides span in a longitudinal direction between first and second ends. The airfoil wall circumscribes an internal core cavity that includes an exit region that spans between the first and second ends and that opens through the arced trailing end. The exit region includes pedestals arranged in a plurality of longitudinal pedestal rows. At least one of the longitudinal pedestal rows is straight and at least one of the longitudinal pedestal rows is arced.

Abstract: 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.

Abstract: A manufacturing method is provided during which a preform component for a turbine engine is provided. The preform component includes a substrate and a locating feature at an exterior surface of the substrate. An outer coating is applied over the substrate. The outer coating covers the locating feature. At least a portion of the preform component and the outer coating are scanned with an imaging system to provide scan data indicative of a location of the locating feature. A cooling aperture is formed in the substrate and the outer coating based on the scan data.