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: A method of forming a component for use in a gas turbine engine comprises the steps of determining a desired shape for a cooling hole on a gas turbine engine component, and determining the likely deposition of a coating to be provided on the component into the cooling hole. An intermediate cooling hole is formed that has an enlarged area from the desired shape to account for deposition of the coating. The component is then coated. A component and an intermediate component for use in a gas turbine engine are also disclosed.

Abstract: One exemplary embodiment of this disclosure relates to a gas turbine engine including a component. The component includes a platform having a mateface on a circumferential side thereof. The platform including a core passageway configured to communicate fluid to the mateface.

Abstract: A cooling circuit for a gas turbine engine comprises a gas turbine engine component having a body with at least one internal cavity defined by a cavity wall. A plurality of cooling holes formed within the cavity wall, wherein each cooling hole is defined by a length extending from a cooling hole inlet to a cooling hole outlet, and wherein the cooling holes are positioned relative to each other such that a minimum allowable ligament distance is maintained between adjacent cooling holes along the entire length of each cooling hole. A gas turbine engine and a method of forming a cooling circuit for a gas turbine engine are also disclosed.

Abstract: A gas turbine engine component includes a wall having an inner surface and an outer surface. At least one non-rectangular slot extends through the outer wall surface and is in communication with a passage. The passages extend through the wall and include an inlet portion located at the inner surface and an outlet portion located at the outer surface. An intermediate portion fluidly connects the inlet portion and the outlet portion.

Abstract: A turbomachine component includes a body defining an interior cooling channel in fluid communication with the exterior of the body for fluid communication with a cooling flow source. At least one flow modifying pedestal is disposed within the interior cooling channel extending in a first direction. The flow modifying pedestal includes at least one flow feature that extends from the pedestal in a second direction.

Abstract: A component according to an exemplary aspect of the present disclosure includes, among other things, an airfoil that includes a first sidewall and a second sidewall joined together at a leading edge and a trailing edge and extending from a base to a tip. A plenum is defined inside the airfoil. A first cooling cavity merges into the plenum and a second cooling cavity merges into the plenum. A rib extends from at least one of the first sidewall and the second sidewall at least partially into the plenum to separate the first cooling cavity from the second cooling cavity.

Abstract: An airfoil includes a leading edge, a trailing edge region, a suction surface, a pressure surface, a cooling passageway, and a column of flow separators. The suction surface and the pressure surface both extend axially between the leading edge and the trailing edge region, as well as radially from a root section of the airfoil to a tip section of the airfoil to define a central cavity of the airfoil. The cooling passageway is located within the central cavity at the trailing edge region. The column of flow separators is located in the cooling passageway adjacent the trailing edge and includes a first flow separator having a first longitudinal axis and a second flow separator having a second longitudinal axis. The first longitudinal axis is offset at an angle with respect to the first longitudinal axis.

Abstract: A wall of a gas turbine engine is provided. The wall may comprise an external surface adjacent a gas path and an internal surface adjacent an internal flow path. A film hole may have an inlet at the internal surface and an outlet at the external surface. A flow accumulator adjacent the inlet may protrude from the internal surface. A method of making an engine component is also provided and comprises the step of forming a component wall comprising an accumulator on an internal surface and a film hole defined by the component wall. The film hole may include an opening adjacent the accumulator and defined by the internal surface.

Abstract: A component for a gas turbine engine includes a trailing edge tip corner that at least partially defines a trailing edge cavity and a multiple of corner features within the trailing edge cavity, the multiple of corner features splayed along the trailing edge tip corner.

Abstract: A component for a gas turbine engine includes a root with a neck that extends into a fir tree with at least one tooth, the root includes a feed passage in communication with a multiple of cooling passages that extend through the neck and fir tree.

Abstract: A blade for a gas turbine engine includes an airfoil having a tip with a terminal end surface and multiple squealer pockets recessed into the terminal end surface.

Abstract: An airfoil for a gas turbine engine includes pressure and suction side walls joined to one another at leading and trailing edges. A stagnation line is located on the pressure side wall aft of the leading edge. A cooling passage is provided between the pressure and suction side walls. Forward-facing cooling holes are provided adjacent to the stagnation line on the pressure side wall and oriented toward the leading edge.

Abstract: A component for a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a body portion that includes a first wall spaced apart from a second wall and disposed about a centerline axis. At least one rib extends between the first wall and the second wall. The at least one rib extends along a rib axis that is transversely angled relative to the centerline axis. At least one impingement hole extends through the at least one rib.

Abstract: An airfoil comprises leading and trailing edges with pressure and suction surfaces defining a chord length therebetween. The pressure and suction surfaces extend from a root section of the airfoil to a tip section. A tip shelf is formed along the tip section between the pressure surface and a tip shelf wall spaced between the pressure surface and the suction surface. A squealer pocket is formed along the tip section between the tip shelf wall and a squealer tip wall extending from the suction surface. The tip shelf extends from within 10% of the cord length measured from the leading edge to within 10% of the chord length measured from the trailing edge. The squealer pocket extends from within 10% of the chord length measured from the leading edge to terminate between 10% and 25% of the chord length measured from the trailing edge.

Abstract: A component for a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a body portion that includes a first wall spaced apart from a second wall. At least one rib extends between the first wall and the second wall and at least one aperture extends through the at least one rib. The at least one aperture is angled relative to a radial axis of the at least one rib.

Abstract: A component according to an exemplary aspect of the present disclosure includes, among other things, an airfoil that includes a pressure sidewall and a suction sidewall that meet together at both a leading edge and a trailing edge. The airfoil extends to a tip. A tip pocket is formed in the tip and terminates prior to the trailing edge. A heat transfer augmentation device is formed in the tip pocket.

Abstract: An airfoil for a gas turbine engine includes an airfoil body that extends in a radial direction from a support. The airfoil body has pressure and suction side walls joined at leading and trailing edges to provide an exterior airfoil surface. A chord-wise direction extends between the leading and trailing edges and a thickness direction transverse to chord-wise direction and extending between the pressure and suction side walls. Cooling passages extend from the support into the airfoil body. The cooling passages include adjacent passageways in the thickness direction and are separated by a chord-wise wall. One of the adjacent passageways is adjacent to another passageway in the chord-wise direction and is separated by a rib in the thickness direction. The rib is discontinued at a location along the radial direction to provide an opening that fluidly connects one of the adjacent passageways to the other passageway.

Abstract: A component for a gas turbine engine according to an exemplary aspect of this disclosure includes a surface having a plurality of cooling holes. The surface includes a first region and a second region each having a plurality of cooling holes. The cooling holes within the first region are arranged differently than the cooling holes in the second region.

Abstract: A blade for a gas turbine engine includes an airfoil that has a tip with a terminal end surface. The terminal end surface includes a recess that has a depth of less than 40 mils (1.016 mm). The recess is filled with a thermal barrier coating. The recess is without any cooling holes.