Abstract: A method of increasing a heat transfer of a cooling fluid passing through a component of a gas turbine engine. The method including the steps of: directing a cooling fluid between an interior surface of an internal cooling cavity of the component and an exterior surface of a baffle insert located in the internal cooling cavity; and creating a plurality of vortices in the cooling fluid as it passes between the exterior surface of the baffle insert and the interior surface of the internal cooling cavity, wherein the internal cooling cavity is elliptical in shape.

Abstract: A combustor panel may include an attachment feature. Because conventional attachment features of conventional combustor panels may be insufficiently cooled, the present disclosure provides various combustor configurations for reducing hotspots in the vicinity of attachment features and/or for providing cooling airflow to and in the vicinity of attachment features.

Abstract: A combustor panel may include an attachment feature. Because conventional attachment features of conventional combustor panels may be insufficiently cooled, the present disclosure provides various combustor configurations for reducing hotspots in the vicinity of attachment features and/or for providing cooling airflow to and in the vicinity of attachment features.

Abstract: A combustor panel may include an attachment feature. Because conventional attachment features of conventional combustor panels may be insufficiently cooled, the present disclosure provides various combustor configurations for reducing hotspots in the vicinity of attachment features and/or for providing cooling airflow to and in the vicinity of attachment features.

Abstract: Airfoils and gas turbine engines having the same, the airfoils including an airfoil body having a leading edge and a trailing edge extending in a radial direction, a cooling cavity defined within the airfoil body, wherein at least one wall of the cooling cavity is a curved end wall, and an end wall contoured pedestal positioned adjacent the curved end wall. The end wall contoured pedestal has a first portion with a contoured side wall and a parallel side wall, and a second portion with tapering side walls, wherein the contoured side wall faces the curved end wall, the contoured side wall paralleling a contour of the curved end wall and defining a meter section therebetween.

Abstract: Disclosed is a module for a gas turbine engine, the module including: a disk rotatable about an engine central axis and a vane fixedly disposed upstream of the disk, the vane being hollow and including a vane cavity, the vane cavity having a vane longitudinal span between a vane upstream opening and vane downstream opening, the vane including: a baffle fixedly supported within the vane cavity, the baffle having a hollow interior and having a baffle longitudinal span between a baffle upstream surface and a baffle downstream surface, the baffle longitudinal span being less than the vane longitudinal span to define a partial space eater baffle; wherein a baffle opening area fluidly connects the vane cavity to the baffle interior, and the baffle opening area is between 5.0×10??5 square inches and 1.5×10??3 square inches.

Abstract: In one embodiment, a component for a gas turbine engine is provided. The component including: an airfoil having a tip portion; a tip shelf located in the tip portion; a first plurality of cooling openings located in an edge of the tip shelf that extends along at least a portion of a pressure side of the airfoil; and a second plurality of cooling openings located in an edge of the tip portion proximate to the tip shelf that extends along at least a portion of a pressure side of the tip portion.

Abstract: A combustor for a gas turbine engine includes a combustor shell having a shell opening therethrough, a combustor panel having a stud attached thereto, the stud extending through the shell opening. The stud includes a standoff to define an intermediate passage between the combustor shell and the combustor panel. A retainer is attached to the stud. A washer surrounds the stud and is positioned between the retainer and the combustor shell. The washer at least partially defines a cooling flow passage configured to direct a cooling airflow through the shell opening to impinge the cooling flow on at least one of the stud or the standoff.

Abstract: An airfoil for a gas turbine engine includes pressure and suction side walls joined to one another at leading and trailing edges to provide an exterior airfoil surface. The pressure and suction side walls are spaced apart from one another in a thickness direction. A stagnation line is located near the leading edge. A cooling passage is provided between the pressure and suction side walls. The showerhead cooling holes are arranged at least one of adjacent to or on the stagnation line. At least one of the showerhead cooling holes has a metering hole fluidly connecting the cooling passage to a diffuser arranged at the exterior airfoil surface. At least one showerhead cooling hole is arranged on each of opposing sides of the stagnation line. Each showerhead cooling hole has the diffuser with a first diffuser angle that expands downstream in the thickness direction in opposing directions from one another when separated by the stagnation line.

Abstract: An engine component assembly includes at least one cavity that is in communication with a source of cooling air. An insert disposed within the cavity includes a plurality of scoops protruding into a flow of cooling air for directing cooling air through the insert and against an inner surface of the cavity.

Abstract: A component of a gas turbine engine having: an internal cooling cavity; a baffle insert configured to be inserted into the internal cooling cavity; a plurality of trip strips extending upwardly from an exterior surface of the baffle insert; at least one separating feature located between a pair of ends of a pair of the plurality of trip strips, wherein the plurality of trip strips and the at least one separating feature are spaced from an interior surface of the internal cooling cavity; a plurality of trip strips extending upwardly from the interior surface; and at least one separating feature located between a pair of ends of the plurality of trip strips located on the interior surface of the internal cooling cavity, wherein the plurality of trip strips and the at least one separating feature of the interior surface are spaced from the exterior surface of the baffle insert.

Abstract: A component of a gas turbine engine, the component having: an internal cooling cavity extending through an interior of the component; a baffle insert configured to be inserted into the internal cooling cavity; a plurality of trip strips extending upwardly from an exterior surface of the baffle insert; and at least one rib extending upwardly from the exterior surface of the baffle insert, wherein the plurality of trip strips and the at least one rib are spaced from an interior surface of the internal cooling cavity.

Abstract: A baffle insert for a component of a gas turbine engine is provided. The baffle insert having: a plurality of trip strips extending upwardly from an exterior surface of the baffle insert; and at least one rib extending upwardly from the exterior surface of the baffle insert.

Abstract: A method of increasing a heat transfer of a cooling fluid passing through a component of a gas turbine engine. The method including the steps of: directing a cooling fluid between an interior surface of an internal cooling cavity of the component and an exterior surface of a baffle insert located in the internal cooling cavity; and creating a plurality of vortices in the cooling fluid as it passes between the exterior surface of the baffle insert and the interior surface of the internal cooling cavity, wherein the internal cooling cavity is elliptical in shape.

Abstract: A gas path component for a gas turbine engine includes an element configured to be exposed to a gas path. The element includes a plurality of internal cooling passages. At least one of the internal cooling passages includes a cross section having a plurality of asymmetrical filleted corners.

Abstract: An array of components in a gas turbine engine includes first and second structures respectively including first and second surfaces that are arranged adjacent to one another to provide a gap. The first and second surfaces respectively have first and second rounded edges at the gap that are arranged in staggered relationship relative to one another.

Abstract: A gas turbine engine component subjected to a flow of high temperature gas includes a wall having first and second wall surfaces and a cooling hole extending through the wall. The cooling hole includes an inlet located at the first wall surface, an outlet located at the second wall surface, a metering section extending downstream from the inlet and a diffusing section extending from the metering section to the outlet. The diffusing section includes a first lobe, a second lobe and a transition region. The first lobe diverges longitudinally and laterally from the metering section and has a first shape. The second lobe is generally opposite the first lobe and diverges longitudinally and laterally from the metering section and has a second shape different from the first shape. The transition region is positioned between the first and second lobes and includes a downstream end adjacent the outlet.

Abstract: This disclosure relates to a gas turbine engine including a component having a leading edge, a pressure side and a suction side opposite the pressure side. The component includes a first group of showerhead holes in the leading edge and a second group of showerhead holes in one of the pressure side and the suction side. The component further includes a first core passageway and a second core passageway separate from the first core passageway. The first core passageway and the second core passageway are in communication with a respective one of the first group of showerhead holes and the second group of showerhead holes.

Abstract: An airfoil is provided. The airfoil may comprise a cross over, an impingement chamber in fluid communication with the cross over, and a first trip strip disposed on a first surface of the impingement chamber. A cooling system is also provided. The cooling system may comprise an impingement chamber, a first trip strip on a first surface of the impingement chamber, and a second trip strip on a second surface of the impingement chamber. An internally cooled engine part is further provided. The internally cooled part may comprise a cross over and an impingement chamber in fluid communication with the cross over. The cross over may be configured to direct air towards a first surface of the impingement chamber. A first trip strip may be disposed on the first surface of the impingement chamber.

Abstract: A component for a gas turbine engine includes a wall and a cooling hole extending through the wall. The wall has a first surface and a second surface. The cooling hole includes a metering section extending downstream from an inlet in the first surface of the wall and a diffusion section extending from the metering section to an outlet in the second surface of the wall. The diffusion section includes a first plurality of lobes diverging longitudinally and laterally from the metering section on a first side of a centerline axis of the cooling hole and a second plurality of lobes diverging longitudinally and laterally from the metering section on a second side of the centerline axis.