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 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 gas turbine engine according to an example of the present disclosure includes, among other things, a rotor and a vane spaced axially from the rotor, and a blade outer air seal spaced radially from the rotor. At least one of the rotor and the vane includes an airfoil section extending from a platform. At least one of the airfoil section, the platform and the blade outer air seal includes a first cavity extending in a first direction, the first cavity defining a reference plane along a parting line formed by a casting die, and a plurality of trip strips including a first set of trip strips distributed in the first direction along a surface of the first cavity and on a first side of the reference plane, each of the plurality of trip strips defining a respective groove axis extending longitudinally between a first end and an opposed, second end of a respective one the plurality of trip strips, and the groove axes being oriented with respect to a pull direction of the casting die.

Abstract: A gas turbine engine component includes an airfoil that has a leading edge and extends from a first side of a platform that has a leading edge overhang. A root portion extends from a second side of the platform opposite the first side. A cooling passage extends through the platform and beneath a trailing edge of the airfoil. The cooling passage includes an inlet located on a second opposite side of the platform. The inlet is located axially upstream of the leading edge of the airfoil and the inlet is located axially upstream of the root portion in the leading edge overhang.

Abstract: A vane according to an exemplary aspect of the present disclosure includes, among other things, a platform extending from an edge face and between spaced apart lateral faces and an airfoil extending outwardly from the platform. The platform includes at least one ejection port in the edge face and at least one passage connected to the at least one ejection port. A method of cooling a component is also disclosed.

Abstract: A component for a gas turbine engine includes a body portion that extends between a leading edge and a trailing edge of the component. The trailing edge includes a flared region and a non-flared region. At least one discharge slot is disposed at least partially within the flared region of the component.

Abstract: A component for a gas turbine engine includes a body portion that extends between a leading edge and a trailing edge of the component. The trailing edge includes a flared region and a non-flared region. At least one discharge slot is disposed at least partially within the flared region of the component.

Abstract: A gas turbine engine according to an example of the present disclosure includes, among other things, a plurality of blade outer air seals and a plurality of airfoils, at least one of seals and airfoils including at least one cooling passage. The cooling passage includes a first wall and an opposed second wall bounding the cooling passage, a surface contour of the first wall having a plurality of first surface features and a surface contour of the second wall having a plurality of second surface features. The first surface features and the second surface features are arranged such that a width of the cooling passage varies along a length of the cooling passage defined by the first surface features and the second surface features. The first surface features have a first profile, and the second surface features have a second, different profile. A casting core for forming cooling passages in an aircraft component is also disclosed.

Abstract: A gas turbine engine component includes an airfoil that has a leading edge and extends from a first side of a platform that has a leading edge overhang. A root portion extends from a second side of the platform opposite the first side. A cooling passage extends through the platform and beneath a trailing edge of the airfoil. The cooling passage includes an inlet located on a second opposite side of the platform. The inlet is located axially upstream of the leading edge of the airfoil and the inlet is located axially upstream of the root portion in the leading edge overhang.

Abstract: A gas turbine engine according to an example of the present disclosure includes, among other things, a plurality of blade outer air seals and a plurality of airfoils, at least one of seals and airfoils including at least one cooling passage. The cooling passage includes a first wall and an opposed second wall bounding the cooling passage, a surface contour of the first wall having a plurality of first surface features and a surface contour of the second wall having a plurality of second surface features. The first surface features and the second surface features are arranged such that a width of the cooling passage varies along a length of the cooling passage defined by the first surface features and the second surface features. The first surface features have a first profile, and the second surface features have a second, different profile. A casting core for forming cooling passages in an aircraft component is also disclosed.

Abstract: A vane according to an exemplary aspect of the present disclosure includes, among other things, a platform extending from an edge face and between spaced apart lateral faces and an airfoil extending outwardly from the platform. The platform includes at least one ejection port in the edge face and at least one passage connected to the at least one ejection port. A method of cooling a component is also disclosed.

Abstract: A gas turbine engine according to an example of the present disclosure includes, among other things, a rotor and a vane spaced axially from the rotor, and a blade outer air seal spaced radially from the rotor. At least one of the rotor and the vane includes an airfoil section extending from a platform. At least one of the airfoil section, the platform and the blade outer air seal includes a first cavity extending in a first direction, the first cavity defining a reference plane along a parting line formed by a casting die, and a plurality of trip strips including a first set of trip strips distributed in the first direction along a surface of the first cavity and on a first side of the reference plane, each of the plurality of trip strips defining a respective groove axis extending longitudinally between a first end and an opposed, second end of a respective one the plurality of trip strips, and the groove axes being oriented with respect to a pull direction of the casting die.

Abstract: A gas turbine engine component includes an airfoil that extends from a first side of a platform and a cooling passage that extends through the platform and includes an inlet located on a second opposite side of the platform. The inlet is located axially upstream of the airfoil.

Abstract: A casting core for an airfoil according to an example of the present disclosure includes, among other things, a first portion extending in a first direction and corresponding to a first cavity of an airfoil. The first portion defines a reference plane along a parting line formed by a casting die. The first portion defines a plurality of grooves corresponding to a plurality of trip strips of the airfoil. Each of the plurality of grooves defines a respective groove axis, and the plurality of grooves are distributed in the first direction along a first side of the reference plane such that one or more of the groove axes are oriented with respect to a pull direction of the casting die. A method for fabricating a gas turbine engine component is also disclosed.

Abstract: A vane according to an exemplary aspect of the present disclosure includes, among other things, a platform extending from an edge face and between spaced apart lateral faces and an airfoil extending outwardly from the platform. The platform includes at least one ejection port in the edge face and at least one passage connected to the at least one ejection port.

Abstract: A gas turbine engine component according to an exemplary aspect of the present disclosure includes, among other things, at least one cooling passage. The at least one cooling passage includes a first wall and a second wall bounding the at least one cooling passage, the first wall having a plurality of first surface features and the second wall having a plurality of second surface features. The plurality of first surface features and the plurality of second surface features are arranged such that a width of the cooling passage varies along a length of the cooling passage defined by the plurality of first surface features and the plurality of second surface features. The plurality of first surface features has a first profile, and the plurality of second surface features has a second, different profile. A casting core for forming cooling passages in an aircraft component is also disclosed.

Abstract: An airfoil for use in a gas turbine engine is provided. The airfoil having: a pressure surface and a suction surface each extending axially from a leading edge to a trailing edge of the airfoil, at least one of the pressure surface, the suction surface, the leading edge and the trailing edge terminating at an edge of a tip section of the airfoil; a plurality of internal cooling channels located within the airfoil; and at least one cooling hole in fluid communication with at least one of the plurality of internal cooling channels, wherein the at least one cooling hole is aligned with an opening or diffuser that extends directly from the at least one cooling hole and wherein the opening or diffuser is formed in and extends through the edge of the tip section of the airfoil.

Abstract: A component for a gas turbine engine includes a platform having an outer surface and an inner surface. A cover plate can be positioned adjacent to the outer surface of the platform. The outer surface of the platform can include a pocket and the cover plate is positioned relative to the pocket to establish a platform cooling channel therebetween.

Abstract: A gas turbine engine component includes a cooling hole. The component includes a first wall having an inlet, a second wall having an outlet and a metering section extending downstream from the inlet and having a substantially convex first surface and a substantially concave second surface. The component also includes a diffusing section extending from the metering section to the outlet. A gas turbine engine wall includes first and second surfaces and a cooling hole extending between an inlet at the first surface and an outlet at the second surface. The cooling hole includes a metering section commencing at the inlet and a diffusing section in communication with the metering section and terminating at the outlet. The metering section includes a top portion having a first arcuate surface and a bottom portion having a second arcuate surface. The first and second arcuate surfaces have arcs extending in substantially similar directions.

Abstract: An airfoil component is described herein. The airfoil component may include an OD platform comprising a gaspath face and a non-gaspath face coupled together via a plurality of cooling holes. The airfoil component may include an airfoil extending in from the outer diameter platform. The plurality of cooling holes comprises a plurality of groups of cooling holes disposed in the outer diameter platform proximate a suction side of the airfoil and a pressure side of the airfoil.