Abstract: An apparatus is provided for a turbine engine. This turbine engine apparatus includes a combustor wall, and the combustor wall includes a panel, a deflector and a quench aperture that extends along an aperture centerline through the combustor wall to an aperture outlet into a combustion chamber. The panel extends axially along and circumferentially about an axis. The panel extends radially between a first panel side and a second panel side. The first panel side forms a peripheral boundary of the combustion chamber. The panel forms a downstream side of the aperture outlet at the first panel side. The deflector projects out from the panel at the first panel side into the combustion chamber. The deflector forms an upstream side of the aperture outlet.

Abstract: An airfoil array may include an endwall, and a plurality of airfoils radially projecting from the endwall. Each airfoil may have a first side and an opposite second side extending axially in chord between a leading edge and a trailing edge. The airfoils may be circumferentially spaced apart on the endwall thereby defining a plurality of flow passages between adjacent airfoils. The airfoil array may further include a convex profiled region extending from the endwall adjacent to the first side of at least one of said plurality of airfoils near the leading edge of the at least one of said plurality of airfoils, and a concave profiled region in the endwall and extending across said at least one of said plurality of flow passages.

Abstract: An airfoil for a gas turbine engine comprises a radially extending body having a transverse cross-section. The transverse cross-section comprises a leading edge, a trailing edge, a pressure side and a suction side. The pressure side extends between the leading edge and the trailing edge with a predominantly concave curvature. The suction side extends between the leading edge and the trailing edge with a predominantly convex curvature. The suction side includes an approximately flat portion flanked by forward and aft convex portions. In another embodiment, the suction side includes a series of local curvature changes that produce inflection points in the convex curvature of the suction side spaced from the trailing edge.

Abstract: A turbine blade for a gas turbine engine includes an airfoil that includes leading and trailing edges joined by spaced apart pressure and suction sides to provide an exterior airfoil surface extending from a platform in a radial direction to a tip. The external airfoil surface is formed in substantial conformance with multiple cross-sectional profiles of the airfoil described by a set of Cartesian coordinates set forth in Table 1. The Cartesian coordinates are provided by an axial coordinate scaled by a local axial chord. A circumferential coordinate is scaled by a local axial chord, and a span location. The local axial chord corresponds to a width of the airfoil between the leading and trailing edges at the span location.

Abstract: An airfoil for a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, an airfoil body including a leading edge and a trailing edge joined by a pressure side and a suction side spaced apart from the pressure side to provide an external airfoil surface extending in a radial direction from at least one platform. The external airfoil surface is formed in substantial conformance with multiple cross-sectional profiles of the airfoil described by a set of Cartesian coordinates set forth in Table 1. The Cartesian coordinates are provided by an axial coordinate scaled by a local axial chord, a circumferential coordinate scaled by a local axial chord, and a span location. The local axial chord corresponds to a width of the airfoil between the leading edge and the trailing edge at the span location, and the Cartesian coordinates in Table 1 have a tolerance relative to the specified coordinates of up to ±0.050 inches (±1.27 mm).

Abstract: An airfoil for a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, an airfoil body including a leading edge and a trailing edge joined by a pressure side and a suction side spaced apart from the pressure side to provide an external airfoil surface extending in a radial direction from at least one platform. The external airfoil surface is formed in substantial conformance with multiple cross-sectional profiles of the airfoil described by a set of Cartesian coordinates set forth in Table 1. The Cartesian coordinates are provided by an axial coordinate scaled by a local axial chord, a circumferential coordinate scaled by a local axial chord, and a span location. The local axial chord corresponds to a width of the airfoil between the leading edge and the trailing edge at the span location, and the Cartesian coordinates in Table 1 have a tolerance relative to the specified coordinates of up to ±0.050 inches (±1.27 mm).

Abstract: An airfoil for a gas turbine engine comprises a radially extending body having a transverse cross-section. The transverse cross-section comprises a leading edge, a trailing edge, a pressure side and a suction side. The pressure side extends between the leading edge and the trailing edge with a predominantly concave curvature. The suction side extends between the leading edge and the trailing edge with a predominantly convex curvature. The suction side includes an approximately flat portion flanked by forward and aft convex portions. In another embodiment, the suction side includes a series of local curvature changes that produce inflection points in the convex curvature of the suction side spaced from the trailing edge.

Abstract: A turbine blade for a gas turbine engine includes an airfoil that includes leading and trailing edges joined by spaced apart pressure and suction sides to provide an exterior airfoil surface extending from a platform in a radial direction to a tip. The external airfoil surface is formed in substantial conformance with multiple cross-sectional profiles of the airfoil described by a set of Cartesian coordinates set forth in Table 1. The Cartesian coordinates are provided by an axial coordinate scaled by a local axial chord. A circumferential coordinate is scaled by a local axial chord, and a span location. The local axial chord corresponds to a width of the airfoil between the leading and trailing edges at the span location.

Abstract: Inefficiencies associated with the formation of horseshoe vortices in a gas turbine engine having a row of radially extending airfoils, are mitigated with an improved flow passage defined by the suction and pressure surface of two adjacent airfoils and endwalls (disposed proximal to the radially inner and outer ends of the airfoils). The flow passage includes a ridge disposed in one of the endwalls adjacent a suction surface of one of the airfoils, the ridge extending longitudinally from a location at or near the maximum circumferential extent of the airfoil suction side to a downstream location proximal to the trailing edge of the suction surface.

Abstract: An airfoil array includes a laterally extending endwall 56 with a series of airfoils such as 28 or 38 projecting from the endwall. The airfoils cooperate with the endwall to define a series of fluid flow passages 74. The endwall has a trough 100 toward a pressure side of the passage and a more elevated profile toward a suction side of the passage for reducing secondary flow losses.

Abstract: A blade or vane includes a platform 130 and an airfoil 132 extending from the platform. The airfoil has a pressure surface offset in a first direction D1 from a part span mean camber line 148, a suction surface offset in a second direction D2 from the part span mean camber line, and a base 146 that is laterally enlarged in the first direction for reducing secondary flow losses.

Abstract: A vane for a turbine engine comprises an airfoil section, an inner platform and an outer platform. The airfoil section comprises pressure and suction surfaces extending from a leading edge to a trailing edge. The inner platform is attached to the airfoil section along an inner flow boundary, where the inner flow boundary extends from an upstream inlet region of the vane to a downstream outlet region of the vane. The outer platform is attached to the airfoil section along an outer flow boundary, where the outer flow boundary extends from the inlet region to the outlet region. An area ratio of the outlet region to the inlet region is greater than 2.4.

Abstract: An airfoil with turning flow comprises a first surface, a second surface, a leading edge and a trailing edge. The first and second surfaces extend from the leading edge to a trailing edge. A chord extends from a midpoint of the leading edge to a midpoint of the trailing edge, defining a duct angle of at least fifteen degrees with respect to a turbine axis. An axial length of the chord defines an aspect ratio of at most one with respect to an average radial span of the leading and trailing edges. In a region of the trailing edge, an axial angle is at most four degrees and a circumferential angle is at most ten degrees, where the circumferential angle is defined in a direction opposite the turning flow.

Abstract: A transition duct assembly for a gas turbine engine includes an outer diameter surface and an inner diameter surface. A substantially annular flowpath having a radial offset is defined between the inner diameter surface and the outer diameter surface. A curvature distribution of the inner diameter surface includes a first region of increasing slope in a radial direction and a second region of decreasing slope in the radial direction adjacent to and downstream of the first region. The first and second regions are configured to reduce flow separation along the outer diameter surface of fluid traveling along the substantially annular flowpath.

Abstract: Inefficiencies associated with the formation of horseshoe vortices in a gas turbine engine having a row of radially extending airfoils, are mitigated with an improved flow passage defined by the suction and pressure surface of two adjacent airfoils and endwalls (disposed proximal to the radially inner and outer ends of the airfoils). The flow passage includes a ridge disposed in one of the endwalls adjacent a suction surface of one of the airfoils, the ridge extending longitudinally from a location at or near the maximum circumferential extent of the airfoil suction side to a downstream location proximal to the trailing edge of the suction surface.

Abstract: An airfoil array includes a laterally extending endwall 56 with a series of airfoils such as 28, or 38 projecting from the endwall. The airfoils cooperate with the endwall to define a series of fluid flow passages 74. The endwall has a hump 84 toward a pressure side of the passage and a less elevated profile toward a suction side of the passage for reducing secondary flow losses.

Abstract: An airfoil with turning flow comprises a first surface, a second surface, a leading edge and a trailing edge. The first and second surfaces extend from the leading edge to a trailing edge. A chord extends from a midpoint of the leading edge to a midpoint of the trailing edge, defining a duct angle of at least fifteen degrees with respect to a turbine axis. An axial length of the chord defines an aspect ratio of at most one with respect to an average radial span of the leading and trailing edges. In a region of the trailing edge, an axial angle is at most four degrees and a circumferential angle is at most ten degrees, where the circumferential angle is defined in a direction opposite the turning flow.

Abstract: A transition duct assembly for a gas turbine engine includes an outer diameter surface and an inner diameter surface. A substantially annular flowpath having a radial offset is defined between the inner diameter surface and the outer diameter surface. A curvature distribution of the inner diameter surface includes a first region of increasing slope in a radial direction and a second region of decreasing slope in the radial direction adjacent to and downstream of the first region. The first and second regions are configured to reduce flow separation along the outer diameter surface of fluid traveling along the substantially annular flowpath.

Abstract: A blade or vane includes a platform 130 and an airfoil 132 extending from the platform. The airfoil has a pressure surface offset in a first direction D1 from a part span mean camber line 148, a suction surface offset in a second direction D2 from the part span mean camber line, and a base 146 that is laterally enlarged in the first direction for reducing secondary flow losses.

Abstract: An airfoil array includes a laterally extending endwall 56 with a series of airfoils such as 28, or 38 projecting from the endwall. The airfoils cooperate with the endwall to define a series of fluid flow passages 74. The endwall has a hump 84 toward a pressure side of the passage and a less elevated profile toward a suction side of the passage for reducing secondary flow losses.