Abstract: An airfoil having one-sided pedestals is disclosed. The airfoil may define various cavities, such as an inflow feed cavity, an impingement cavity, and an outflow cavity. The various cavities may be connected by crossover sections such as an inflow crossover section and an outflow crossover section. Cooling air may be conducted into the inflow feed cavity, out of the inflow feed cavity through an inflow crossover section into an impingement cavity, and through an impingement cavity. The cooling air may be conducted out of the impingement cavity and into an outflow cavity through an outflow crossover section. Various cavities may include one-wall pedestals. One-wall pedestals may be structures extending from a wall of a cavity into the void of the cavity, whereupon cooling air may impinge, effectuating convective cooling.

Abstract: An airfoil for a gas turbine engine includes pressure and suction surfaces provided by pressure and suction walls that extend in a radial direction and are joined at a leading edge and a trailing edge. A cooling passage is arranged between the pressure and suction walls and extends to the trailing edge. Elongated pedestals are arranged in the cooling passage and interconnect the pressure and suction walls. The elongated pedestals are spaced apart from one another in the radial direction and extend from a plane to the trailing edge. A metering pedestal includes at least a portion that is arranged between the plane and the trailing edge. The portion is provided between adjacent elongated pedestals in the radial direction.

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: A turbomachinery component includes an airfoil section with a cooling passage extending within the airfoil section. A wall is defined between the cooling passage and an exterior surface of the airfoil. At least one row of cooling holes is positioned along a cooling portion of the external wall proximate a leading edge of the airfoil for fluid communication between the cooling passage and the exterior surface of the airfoil. The cooling portion of the airfoil wall is thicker than an average adjacent airfoil wall thickness.

Abstract: A gas turbine engine airfoil includes a body that provides an exterior airfoil surface that extends in a radial direction to a tip. The exterior surface has a leading edge in a forward direction and a trailing edge in an aft direction. The tip includes a squealer pocket that has a recess surface. A cooling passage is arranged in the body. Each of the cooling holes extends from an inlet at the cooling passage to an outlet at the recessed surface. The inlet and outlet are arranged at an angle in an angular direction relative to the recessed surface. The angular direction is toward at least one of the forward and aft directions.

Abstract: A component for a gas turbine engine including a body having at least one internal cooling cavity and a plurality of film cooling holes disposed along a first edge of the body. At least one of the film cooling holes includes a metering section defining an axis, and a diffuser section having a centerline. The centerline of the diffuser section is offset from the axis of the metering section.

Abstract: The gas turbine component showerhead cooling hole layouts described herein include minimal lateral cooling hole exit diffusion on the middle showerhead cooling hole rows and interior facing sides of outer rows. In this way, rows of cooling holes may be placed close together. Stated another way, the outer showerhead cooling hole rows substantially only include lateral cooling hole exit diffusion in the direction away from the other rows to again allow the rows to be placed close together.

Abstract: A component for a gas turbine engine, the component having: a cooling slot located on a surface of the component, the cooling slot being defined by a plurality of diffuser portions each extending from a respective one of a plurality of cooling openings providing cooling fluid to the cooling slot.

Abstract: A seal arrangement for a gas turbine engine according to an example of the present disclosure includes, among other things, a component including a body having a cold side surface adjacent to a mate face, and a seal member including a leading edge region and a trailing edge region spaced by sidewalls. The seal member defines one or more grooves. A length of the one or more grooves abuts the cold side surface to define one or more cooling passages, with at least one of the one or more cooling passages having a flared inlet defined by a corresponding one of the one or more grooves.

Abstract: An airfoil structure for a gas turbine engine includes an airfoil which includes a leading edge and a trailing edge. A platform is located adjacent a first end of the airfoil and includes a core passage that extends through the platform, a mate-face for engaging an adjacent airfoil structure and a set of impingement cooling holes in communication with the core passage that extend through the mate-face adjacent the trialing edge of the airfoil.

Abstract: A component of a gas turbine engine includes a cooling hole extending from a first side to a second side that includes an inlet portion disposed about an axis that includes an area defining an inlet area through a first surface. The cooling hole further includes a diffuser portion in communication with the inlet portion. The diffuser portion defines an exit area and an area ratio of the exit area to the inlet area is provided that provides improved cooling efficiencies.

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: A seal arrangement for a gas turbine engine according to an example of the present disclosure includes, among other things, a component including a body having a cold side surface adjacent to a mate face, and a seal member including a leading edge region and a trailing edge region spaced by sidewalls. The seal member defines one or more grooves. A length of the one or more grooves abuts the cold side surface to define one or more cooling passages, with at least one of the one or more cooling passages having a flared inlet defined by a corresponding one of the one or more grooves.

Abstract: A gas turbine engine blade includes a blade portion having a leading edge and a trailing edge. A first surface connects the leading edge to the trailing edge and a second surface connects the leading edge to the trailing edge. A tip section is located at a first end of the blade portion and includes a pocket protruding into the tip section from an outermost end of the tip section. The pocket has a first side wall adjacent the first surface and a second side wall adjacent the second surface. At least one of the first side wall and the second side wall have a curve distinct from a curve of the corresponding adjacent surface.

Abstract: A seal arrangement for a gas turbine engine according to an example of the present disclosure includes, among other things, a component including a body having a cold side surface adjacent to a mate face, and a seal member including a leading edge region and a trailing edge region spaced by sidewalls. The seal member defines one or more grooves. The one or more grooves abut the cold side surface to define one or more cooling passages, with at least one of the one or more cooling passages having a flared inlet defined by a corresponding one of the one or more grooves.

Abstract: A seal for a gas turbine engine includes a plurality of seal arc segments. Each of the seal arc segments includes radially inner and outer sides and sloped first and second circumferential sides. The seal arc segments are circumferentially arranged about an axis such that the sloped first and second circumferential sides define gaps circumferentially between adjacent ones of the seal arc segments. Each of the gaps extends from the radially inner sides along a respective central gap axis that slopes with respect to a radial direction from the axis.

Abstract: An airfoil is provided. The airfoil comprises a crossover and an impingement cavity in fluid communication with the crossover and having an internal surface. At least a portion of the internal surface comprises an undulating internal surface. A plurality of trip strips may be disposed on the at least a portion of the internal surface to define the undulating internal surface. A gas turbine engine and an internally-cooled engine part are also provided.

Abstract: A rotor blade for a gas turbine engine is provided. The rotor blade having: an attachment; an airfoil extending from the attachment to a tip; and a tip shelf located in a surface of the tip proximate to a pressure side of the airfoil, wherein the tip shelf has a ledge portion extending from the pressure side to a wall portion extending upwardly from the ledge portion to the tip and wherein the wall portion is configured to have a convex portion with respect to the pressure side of the airfoil as it extends from a leading edge to a trailing edge of the airfoil.

Abstract: A gas turbine engine component is provided. The gas turbine engine component comprises a platform, a platform cooling core circuit inside of the platform, an airfoil extending from the platform, and a plurality of one-wall pedestals inside the platform cooling core circuit. The plurality of one-wall pedestals partially extend from only a first endwall of the platform cooling core circuit toward a second endwall of the platform cooling core circuit. The gas turbine engine and a system for cooling an airfoil platform of a gas turbine engine component are also provided.

Abstract: A rotor blade for a gas turbine engine is provided. The rotor blade having: an attachment; an airfoil extending from the attachment to a tip; and a squealer pocket located in a surface of the tip, wherein the squealer pocket is at least partially surrounded by a first surface of a wall located between the squealer pocket and a pressure side of the airfoil, wherein the first surface of the wall has a convex configuration with respect to the pressure side of the airfoil as it extends from a leading edge to a trailing edge of the airfoil.