Abstract: A method is provided for machining an airfoil section (12) of a turbine blade or vane produced by a casting process. The airfoil section (12) has an outer wall (18) delimiting an airfoil interior having one or more internal cooling passages (28). The method involves: receiving design data pertaining to the airfoil section (12), including a nominal outer airfoil form (40N) and nominal wall thickness (TN) data; generating a machining path by determining a target outer airfoil form (40T), the target outer airfoil form (40T) being generated by adapting the nominal outer airfoil form (40N) such that a nominal wall thickness (TN) is maintained at all points on the outer wall around the one or more internal cooling passages (28) in a subsequently machined airfoil section; and machining an outer surface (18a) of the airfoil section (12) produced by the casting process according to the generated machining path, to remove excess material to conform to the generated target outer airfoil form (40T).

Abstract: A gas turbine blade having a casted metal airfoil, the airfoil has a main wall defining at least one interior cavity, having a first side wall and a second side wall, which are coupled to each other at a leading edge and a trailing edge, extending in a radial direction from a blade root to a blade tip and defining a radial span from 0% at the blade root to 100% at the blade tip. The main airfoil has a radial span dependent chord length defined by a straight line connecting the leading edge and the trailing edge as well as a radial span dependent solidity ratio of metal area to total cross-sectional area. Solidity ratios in a machined zone of the airfoil from 80% to 85% of span are below 35%, in particular all solidity ratios in the zone.

Abstract: A bladed rotor system includes first and second sets of blades with respective airfoils each having at least one internal cavity. The airfoils of both the first and second sets of blades have identical outer shapes defined by an outer surface of an outer wall of the respective airfoils. The airfoils of the first set of blades are distinguished from the airfoils of the second set of blades by a geometry and/or position of at the least one internal cavity, which is unique to blades of a given set. The natural frequency of a blade of the first set differs from the natural frequency of a blade of the second set by a predetermined amount. The blades of the first set and the second set are alternately arranged in a periodic fashion in said circumferential row, to provide a frequency mistuning to stabilize flutter of the blades.

Abstract: A bladed rotor system includes first and second sets of blades with respective airfoils each having at least one internal cavity. The airfoils of both the first and second sets of blades have identical outer shapes defined by an outer surface of an outer wall of the respective airfoils. The airfoils of the first set of blades are distinguished from the airfoils of the second set of blades by a geometry and/or position of at the least one internal cavity, which is unique to blades of a given set. The natural frequency of a blade of the first set differs from the natural frequency of a blade of the second set by a predetermined amount. The blades of the first set and the second set are alternately arranged in a periodic fashion in said circumferential row, to provide a frequency mistuning to stabilize flutter of the blades.

Abstract: A gas turbine blade having a casted metal airfoil, the airfoil has a main wall defining at least one interior cavity, having a first side wall and a second side wall, which are coupled to each other at a leading edge and a trailing edge, extending in a radial direction from a blade root to a blade tip and defining a radial span from 0% at the blade root to 100% at the blade tip. The main airfoil has a radial span dependent chord length defined by a straight line connecting the leading edge and the trailing edge as well as a radial span dependent solidity ratio of metal area to total cross-sectional area. Solidity ratios in a machined zone of the airfoil from 80% to 85% of span are below 35%, in particular all solidity ratios in the zone.

Abstract: A method is provided for machining an airfoil section (12) of a turbine blade or vane produced by a casting process. The airfoil section (12) has an outer wall (18) delimiting an airfoil interior having one or more internal cooling passages (28). The method involves: receiving design data pertaining to the airfoil section (12), including a nominal outer airfoil form (40N) and nominal wall thickness (TN) data; generating a machining path by determining a target outer airfoil form (40T), the target outer airfoil form (40T) being generated by adapting the nominal outer airfoil form (40N) such that a nominal wall thickness (TN) is maintained at all points on the outer wall around the one or more internal cooling passages (28) in a subsequently machined airfoil section; and machining an outer surface (18a) of the airfoil section (12) produced by the casting process according to the generated machining path, to remove excess material to conform to the generated target outer airfoil form (40T).

Abstract: A gas turbine engine blade (10), including a base portion (12) having a cast wall, and a tip portion (14) attached to the base portion and having a wall (60) formed by an additive manufacturing process. The tip portion wall may be formed to be solid and less than 2 mm in thickness, or it may be corrugated and be greater than 2 mm in thickness. Openings (80) defining the wall corrugations may be semi-circular, rectangular, trapezoidal, or elliptical in cross-sectional shape. The resulting blade has lower tip mass while retaining adequate mechanical properties. The tip portion may be formed to have a directionally solidified grain structure on a base portion having an equiaxed grain structure.

Abstract: A system and method to minimize flow induced vibration in a gas turbine exhaust is provided. The system includes a turbine exhaust manifold connected to a turbine exhaust cylinder establishing a fluid flow path, the fluid flow path bounded radially outward by an outer cylindrical surface and bounded radially inward by an inner cylindrical surface. At least one tangential strut is arranged between the outer cylindrical surface and the inner cylindrical surface. A first flap is arranged diagonally between the tangential strut and the outer cylindrical surface or the inner cylindrical surface where the first flap minimizes vortex shedding of the fluid flow from the tangential strut.

Abstract: Disclosed are a casing arrangement and a method to reduce vibrations in a gas turbine casing. The casing arrangement includes a turbine exhaust cylinder connected to a turbine exhaust manifold establishing a fluid flow path, the fluid flow path including an inner and an outer flow path. A damping blanket damps the vibrations and is coupled to a surface of the inner flow path via a constraining layer.

Abstract: Disclosed are a system for damping vibrations of a gas turbine exhaust and a method to damp vibrations of a gas turbine exhaust. The system includes a turbine exhaust cylinder connected to a turbine exhaust manifold establishing a fluid flow path, the fluid flow path including an inner and an outer flow path. A damping blanket damps the vibrations and is coupled to a surface of the inner flow path via a constraining layer clamped by a plurality of studs.

Abstract: Disclosed are a casing arrangement and a method to reduce critical panel mode response in a gas turbine casing. The casing arrangement includes a turbine exhaust cylinder connected to a turbine exhaust manifold establishing a fluid flow path, the fluid flow path including an inner and an outer flow path. A plurality of stiffening ribs are coupled to a surface of the inner flow path which effectively increases the stiffness reducing the critical panel mode response.

Abstract: A gas turbine engine blade (10), including a base portion (12) having a cast wall, and a tip portion (14) attached to the base portion and having a wall (60) formed by an additive manufacturing process. The tip portion wall may be formed to be solid and less than 2 mm in thickness, or it may be corrugated and be greater than 2 mm in thickness. Openings (80) defining the wall corrugations may be semi-circular, rectangular, trapezoidal, or elliptical in cross-sectional shape. The resulting blade has lower tip mass while retaining adequate mechanical properties. The tip portion may be formed to have a directionally solidified grain structure on a base portion having an equiaxed grain structure.

Abstract: Disclosed are a casing arrangement and a method to reduce vibrations in a gas turbine casing. The casing arrangement includes a turbine exhaust cylinder connected to a turbine exhaust manifold establishing a fluid flow path, the fluid flow path including an inner and an outer flow path. A damping blanket damps the vibrations and is coupled to a surface of the inner flow path via a constraining layer.

Abstract: Disclosed are a system for damping vibrations of a gas turbine exhaust and a method to damp vibrations of a gas turbine exhaust. The system includes a turbine exhaust cylinder connected to a turbine exhaust manifold establishing a fluid flow path, the fluid flow path including an inner and an outer flow path. A damping blanket damps the vibrations and is coupled to a surface of the inner flow path via a constraining layer clamped by a plurality of studs.

Abstract: A system and method to minimize flow induced vibration in a gas turbine exhaust is provided. The system includes a turbine exhaust manifold connected to a turbine exhaust cylinder establishing a fluid flow path, the fluid flow path bounded radially outward by an outer cylindrical surface and bounded radially inward by an inner cylindrical surface. At least one tangential strut is arranged between the outer cylindrical surface and the inner cylindrical surface. A first flap is arranged diagonally between the tangential strut and the outer cylindrical surface or the inner cylindrical surface where the first flap minimizes vortex shedding of the fluid flow from the tangential strut.

Abstract: Disclosed are a casing arrangement and a method to reduce critical panel mode response in a gas turbine casing. The casing arrangement includes a turbine exhaust cylinder connected to a turbine exhaust manifold establishing a fluid flow path, the fluid flow path including an inner and an outer flow path. A plurality of stiffening ribs are coupled to a surface of the inner flow path which effectively increases the stiffness reducing the critical panel mode response.

Abstract: A vane structure for a gas turbine engine. The vane structure includes a radially outer platform and a radially inner platform, and an airfoil having an outer wall extending radially between the outer platform and the inner platform. A cooling passage is defined within the outer wall and has a plurality of radially extending channels. An outer end turn structure is located at the outer platform to conduct cooling fluid in a chordal direction between at least two of the channels. The outer end turn structure includes an enlarged portion wherein the enlarged portion is defined by an enlarged dimension, in a direction transverse to the chordal direction, between the at least two channels.

Abstract: A turbine airfoil (31) with an end portion (42) that tapers (44) toward the end (43) of the airfoil. A ridge (46) extends around the end portion. It has proximal (66) and distal (67) sides. A shroud platform (50) is bi-cast onto the end portion around the ridge without bonding. Cooling shrinks the platform into compression (62) on the end portion (42) of the airfoil. Gaps between the airfoil and platform are formed using a fugitive material (56) in the bi-casting stage. These gaps are designed in combination with the taper angle (44) to accommodate differential thermal expansion while maintaining a gas seal along the contact surfaces. The taper angle (44) may vary from lesser on the pressure side (36) to greater on the suction side (38) of the airfoil. A collar portion (52) of the platform provides sufficient contact area for connection stability.

Abstract: A turbine blade assembly includes a turbine blade having a pressure sidewall and an opposed suction sidewall and a first snubber assembly associated with one of the pressure sidewall and the suction sidewall. The first snubber assembly includes a first base portion extending outwardly from the one of the pressure sidewall and the suction sidewall, and a first snubber portion. The first base portion is integrally cast with the turbine blade and includes first connection structure. The first snubber portion is bicast onto the first base portion and includes second connection structure that interacts with the first connection structure to substantially prevent separational movement between the first base portion and the first snubber portion.

Abstract: A turbine vane for a turbine machine comprising an intermediate section having a nominal airfoil profile substantially in accordance with Cartesian coordinate values of X, Y and Z set forth in Table I wherein Z is a radial distance along a stacking axis that is normal to a centerline of the turbine machine and contain the X and Y values with Z value beginning at innermost aerodynamic point and the Z values represent a radial height of the vane and the X and Y values define the nominal airfoil profile at each radial height Z.