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 includes one of a turbine section and a compressor section having multiple stages. At least one of the stages defines an outer diameter comprised of a plurality of circumferentially arranged blade outer air seals. Each blade outer air seal is spaced from each adjacent blade outer air seal in the plurality of circumferentially arranged blade outer air seals via a mateface gap. The mateface gap is oblique to a radius of the gas turbine engine, such that air entering the mateface gap is directed to an inner diameter surface of at least one of the blade outer air seals in the plurality of blade outer air seals.

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, among other things, an airfoil section that has an internal wall and an external wall. The external wall defines pressure and suction sides that extends in a chordwise direction between a leading edge and a trailing edge, a first impingement cavity and a second impingement cavity bounded by the external wall at a leading edge region that defines the leading edge. A first crossover passage within the internal wall is connected to the first impingement. The first crossover passage defines a first passage axis that intersects a surface of the first impingement cavity. A second crossover passage within the internal wall is connected to the second impingement cavity. The second crossover passage defines a second passage axis that intersects a surface of the second impingement cavity.

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 liner panel for use in a combustor of a gas turbine engine, the liner panel including a cold side; and a rail that extends from the cold side, the rail includes a first diffusion interface passage surface and a second diffusion interface passage surface, the first diffusion interface passage surface angled with respect to the second diffusion interface passage surface.

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 having an airfoil to platform fillet, wherein the fillet is contoured in accordance with the coordinates of Table 1.

Abstract: A gas turbine engine includes a compressor, a combustor fluidly connected to the compressor via a core flowpath, and a turbine fluidly connected to the combustor via the core flowpath. The turbine includes at least one stage having a plurality of rotors and a plurality of vanes. An outer diameter of the core flowpath at at least one stage is at least partially defined by a set of circumferentially arranged blade outer air seals. Each blade outer air seal includes a platform. An internal cooling cavity is defined within the platform. At least one mateface of the platform includes a cooling trench, and a first set of cooling holes connecting the internal cavity to the cooling trench.

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 gas turbine engine includes a compressor, a combustor fluidly connected to the compressor via a core flowpath, and a turbine fluidly connected to the combustor via the core flowpath. The turbine includes at least one stage having a plurality of rotors and a plurality of vanes. An outer diameter of the core flowpath at at least one stage is at least partially defined by a set of circumferentially arranged blade outer air seals. Each blade outer air seal includes a platform. An internal cooling cavity is defined within the platform. At least one mateface of the platform includes a cooling trench, and a first set of cooling holes connecting the internal cavity to the cooling trench.

Abstract: A gas turbine engine includes one of a turbine section and a compressor section having multiple stages. At least one of the stages defines an outer diameter comprised of a plurality of circumferentially arranged blade outer air seals. Each blade outer air seal is spaced from each adjacent blade outer air seal in the plurality of circumferentially arranged blade outer air seals via a mateface gap. The mateface gap is oblique to a radius of the gas turbine engine, such that air entering the mateface gap is directed to an inner diameter surface of at least one of the blade outer air seals in the plurality of blade outer air seals.

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: A combustor for a gas turbine engine includes a support shell; a first liner panel mounted to the support shell via a multiple of studs, the first liner panel including a first rail that extends from a cold side of the first liner panel such that the rail is non-perpendicular to the cold side and includes a concave surface to at least partially form a curved interface passage; and a second liner panel mounted to the support shell via a multiple of studs, the first liner panel including a second rail that extends from a cold side of the second liner panel and includes a convex surface to at least partially form the curved interface passage.

Abstract: A combustor for a gas turbine engine includes a support shell; a first liner panel mounted to the support shell via a multiple of studs, the first liner panel including a first rail that extends from a cold side of the first liner panel; a second liner panel mounted to the support shell via a multiple of studs, the second liner panel including a second rail that extends from a cold side of the second liner panel adjacent to the first rail to form an interface passage; and at least one heat transfer feature within the interface passage.

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 blade for a gas turbine engine includes an airfoil which includes pressure and suction side surfaces joined at leading and trailing edges to provide an exterior airfoil surface. The airfoil extends from a 0% span position to a 100% span position at a tip. A mean chamber line is provided at the tip and extends from the leading edge to the trailing edge. The mean chamber line has a leading edge tangent line at the leading edge at the tip and a trailing edge tangent line at the trailing edge at the tip. The leading edge tangent line and the trailing edge tangent line intersect to provide a total camber angle in a range of ?35° to 35°. The pressure side includes a tip shelf.

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 an airfoil body having a pressure side and a suction side that each radially extend between a first radial boundary end and a second radial boundary end and each axially extend between a leading edge and a trailing edge. The airfoil body defines a plurality of first skin core passages disposed proximate the suction side and radially extend from the first radial boundary end towards the second radial boundary end, and a plurality of second skin core passages that radially extend toward the second radial boundary end and circumferentially extend towards the suction side proximate the second radial boundary end.