Abstract: A method of repairing a tip for an airfoil is disclosed herein. In various embodiments, the method comprises: removing a portion of the airfoil that reduces a radial height of the airfoil and at least partially removes a defect of the airfoil; removing a remainder of the defect; joining, via gas tungsten arc welding, a filler material to the airfoil to repair the defect; and joining a first mating surface of the airfoil to a second mating surface of a blade tip body via a solid-state joining process.

Abstract: An electrode includes a metallic electrode body that has a surface region that is enriched in at least one of aluminum or chromium. The aluminum and/or chromium that is vaporized from the surface region suppresses vaporization loss of aluminum or chromium from the machined surface of the metallic workpiece during electric discharge machining.

Abstract: A turbine blade has an attachment root and an airfoil. A cooling passageway system has a plurality of trunks extending from respective inlets along the root inner diameter end from a leading trunk near a first axial end to a trailing trunk near a second axial end; and a plurality of outlets along the airfoil including trailing edge outlets fed by the trailing trunk. Viewed normal to a root end-to-end centerplane: the trailing trunk has a turn passing forward and then rearward; an outside of the turn protrudes forward; and the outside of the turn has a tighter curvature than an inside of the turn.

Abstract: A component for a gas turbine engine includes a mateface with a purge flow interface, the mateface comprises a pocket located in communication with a feather seal slot in a mateface. A vane for a gas turbine engine includes a platform that extends from the airfoil, the platform comprising a mateface with a feather seal slot and a pocket in communication with the feather seal slot, wherein the pocket is of a cross-sectional shape larger than the feather seal slot.

Abstract: A turbine blade for a gas turbine engine. The turbine blade having a plurality of cooling holes defined therein, wherein the plurality of cooling holes are located in the turbine blade according to the coordinates of Table 1 and at least some of the plurality of cooling holes are located in an airfoil of the turbine blade.

Abstract: A turbine blade has an attachment root and an airfoil. A cooling passageway system has a plurality of trunks extending from respective inlets along the root inner diameter end from a leading trunk near a first axial end to a trailing trunk near a second axial end; and a plurality of outlets along the airfoil including trailing edge outlets fed by the trailing trunk. Viewed normal to a root end-to-end centerplane: the trailing trunk has a turn passing forward and then rearward; an outside of the turn protrudes forward; and the outside of the turn has a tighter curvature than an inside of the turn.

Abstract: An airfoil for a gas turbine engine according to an example of the present disclosure includes a first cavity, a second cavity, and a rib extending from a suction sidewall to a pressure sidewall and separating the first cavity from the second cavity. The rib includes a central portion, a first edge portion extending from the pressure sidewall to the central portion, and a second edge portion extending from the suction sidewall to the central portion. The rib defines one or more communication openings from the first cavity to the second cavity in the first or second edge portion.

Abstract: Vane assemblies for turbine engines are described. The vane assemblies include a vane having an internal cavity and a vane platform and a vane rail defining, in part, an outer diameter supply cavity. A blade outer air seal support (BOAS support) is arranged adjacent the vane and engages with a portion of the vane, the blade outer air seal support having BOAS support rail, and a BOAS supported on the BOAS support and engaging with a portion of the vane. The BOAS support includes a first cooling flow aperture configured to enable a cooling flow to cool at least the BOAS and a second cooling flow aperture formed in the BOAS support rail. The vane rail includes a third cooling flow aperture to form a cooling flow path through the second cooling flow aperture and the third cooling flow aperture to fluidly connect to the outer diameter supply cavity.

Abstract: A method is provided that involves rotational equipment that includes a case and a first rotor blade within the case. During the method, a probe is arranged within the case adjacent the first rotor blade. A magnetic characteristic of the first rotor blade is measured using the probe. Presence of internal corrosion within the first rotor blade is determined based on the measured magnetic characteristic.

Abstract: A turbine blade for a gas turbine engine. The turbine blade having a plurality of cooling holes defined therein, wherein the plurality of cooling holes are located in the turbine blade according to the coordinates of Table 1 and at least some of the plurality of cooling holes are located in an airfoil of the turbine blade.

Abstract: A component for a gas turbine engine includes a mateface with a purge flow interface, the mateface comprises a pocket located in communication with a feather seal slot in a mateface. A vane for a gas turbine engine includes a platform that extends from the airfoil, the platform comprising a mateface with a feather seal slot and a pocket in communication with the feather seal slot, wherein the pocket is of a cross-sectional shape larger than the feather seal slot.

Abstract: Vane assemblies for turbine engines are described. The vane assemblies include a vane having an internal cavity and a vane platform and a vane rail defining, in part, an outer diameter supply cavity. A blade outer air seal support (BOAS support) is arranged adjacent the vane and engages with a portion of the vane, the blade outer air seal support having BOAS support rail, and a BOAS supported on the BOAS support and engaging with a portion of the vane. The BOAS support includes a first cooling flow aperture configured to enable a cooling flow to cool at least the BOAS and a second cooling flow aperture formed in the BOAS support rail. The vane rail includes a third cooling flow aperture to form a cooling flow path through the second cooling flow aperture and the third cooling flow aperture to fluidly connect to the outer diameter supply cavity.

Abstract: A turbine blade for a gas turbine engine having a plurality of cooling holes defined therein, the plurality of cooling holes are located in an airfoil of the turbine blade according to the coordinates of Table 1.

Abstract: A turbine blade for a gas turbine engine. The turbine blade having a plurality of cooling holes defined therein, wherein the plurality of cooling holes are located in the turbine blade according to the coordinates of Table 1 and at least some of the plurality of cooling holes are located in an airfoil of the turbine blade.

Abstract: A method is provided that involves rotational equipment that includes a case and a first rotor blade within the case. During the method, a probe is arranged within the case adjacent the first rotor blade. A magnetic characteristic of the first rotor blade is measured using the probe. Presence of internal corrosion within the first rotor blade is determined based on the measured magnetic characteristic.

Abstract: A gas turbine engine component includes a wall that provides an exterior surface and an interior flow path surface. The wall has a wall thickness. A protrusion is arranged on the wall that extends a height beyond the wall thickness and provides a portion of the interior flow path surface. A film cooling hole that has an inlet is provided on the protrusion and extends to an exit on the exterior surface.

Abstract: Divided baffles for a gas turbine engines are provided. The divided baffles include a baffle body having a first end and a second end, at least one divider located within the baffle body and extending from the first end to the second end arranged to divide an interior of the baffle body into two or more feed cavities, at least one cap arranged at the first end to bound a first feed cavity of the two or more feed cavities, and at least one cap arranged at the second end to bound a second feed cavity of the two or more feed cavities.

Abstract: One exemplary embodiment of this disclosure relates to a method of forming an engine component. The method includes forming an engine component having an internal passageway, the internal passageway formed with an initial dimension. The method further includes establishing a flow of machining fluid within the internal passageway, the machining fluid changing the initial dimension.

Abstract: Core assemblies for manufacturing airfoils and airfoils for gas turbine engines are described. The core assemblies include a tip flag cavity core having an upstream portion, a tapering portion, and a downstream portion, with the tapering portion located between the upstream portion and the downstream portion and the downstream portion defines an exit in a formed airfoil. The upstream portion has a first radial height H1, the downstream portion has a second radial height H2 that is less than the first radial height H1, the tapering portion transitions from the first radial height H1 at an upstream end to the second radial height H2 at a downstream end, and at least one metering pedestal aperture is located within the tapering portion.

Abstract: An airfoil includes pressure and suction side walls that extend in a chord-wise direction between leading and trailing edges. The pressure and suction side walls extend in a radial direction to provide an exterior airfoil surface. A main-body core cooling passage is arranged between the pressure and suction walls in a thickness direction and extends radially toward a platform. A skin core cooling passage is arranged in one of the pressure and suction side walls to form a hot side wall and a cold side wall. The hot side wall defines a portion of the exterior airfoil surface and the cold side wall defines a portion of the core passage. The skin core cooling passage is divided by a wall into two discrete first and second skin core cooling passages each supplied with cooling fluid from opposing sides.