Abstract: A turbine nozzle includes an airfoil shape. The airfoil shape may have a nominal profile substantially in accordance with Cartesian coordinate values of X, Y and Z set forth in Table I. The Cartesian coordinate values of X, Y and Z are non-dimensional values from 0% to 100% convertible to dimensional distances in inches by multiplying the Cartesian coordinate values of X, Y and Z by a height of the airfoil in inches. The X and Y values, when connected by smooth continuing arcs, define airfoil profile sections at each distance Z. The airfoil profile sections at Z distances may be joined smoothly with one another to form a complete airfoil shape.

Abstract: The present application provides a turbine nozzle including an airfoil shape. The airfoil shape may have a nominal profile substantially in accordance with Cartesian coordinate values of X, Y and Z set forth in Table I. The Cartesian coordinate values of X, Y and Z are non-dimensional values from 0% to 100% convertible to dimensional distances in inches by multiplying the Cartesian coordinate values of X, Y and Z by a height of the airfoil in inches. The X and Y values, when connected by smooth continuing arcs, define airfoil profile sections at each distance Z. The airfoil profile sections at Z distances may be joined smoothly with one another to form a complete airfoil shape.

Abstract: A turbine blade includes an airfoil having a tip shroud. The tip shroud has an edge, and the edge has a profile substantially in accordance with values of X and Y in a Cartesian coordinate system set forth in Table 1 at points 1-30. The X and Y values represent distances that may be proportionally scaled by a common multiplier which, once scaled and connected, define the profile of the edge of the tip shroud. The profile of the edge lies in an envelope within +/?20%, +/?10% or +/?0% in a direction normal to any location along the edge set forth by the points in Table 1.

Abstract: A turbine blade includes pressure and suction surfaces connected to define an interior through which coolant is passable. First and second pedestal arrays, each include pedestals respectively coupled to radially outboard portions of respective interior faces of one of the pressure and suction surfaces. The pedestals of the first pedestal array are separated from and directly opposed to pedestals of the second pedestal array by gaps respectively defined therebetween.

Abstract: A turbine blade includes an airfoil having a tip shroud. The tip shroud has an edge, and the edge has a profile substantially in accordance with values of X and Y in a Cartesian coordinate system set forth in Table 1 at points 1-30. The X and Y values represent distances that may be proportionally scaled by a common multiplier which, once scaled and connected, define the profile of the edge of the tip shroud. The profile of the edge lies in an envelope within +/?20%, +/?10% or +/?0% in a direction normal to any location along the edge set forth by the points in Table 1.

Abstract: A blade has an airfoil, and the blade is configured for use with a turbomachine. The airfoil has a throat distribution measured at a narrowest region in a pathway between adjacent blades, at which adjacent blades extend across the pathway between opposing walls to aerodynamically interact with fluid flow. The airfoil defines the throat distribution, and the throat distribution reduces aerodynamic loss and improves aerodynamic loading on the airfoil. The airfoil has a linear trailing edge profile.

Abstract: The present application thus provides a method for reducing stress on a turbine blade wherein each of the turbine blades includes a dovetail with lockwire tab. The method may include the steps of (a) determining a starting line for a backcut relative to a lockwire tab end, (b) determining a cut angle for the backcut, and (c) removing material from the lockwire tab according to the starting line and the cut angle to form the dovetail backcut. The starting line may be positioned about 0.6 inches (about 15.24 millimeters), plus or minus 0.065 inches (about 1.65 millimeters) from the lockwire tab end along the dovetail axis.

Abstract: A rotor blade includes a mounting portion comprising a mounting body that is formed to receive a coolant therein. An airfoil portion extends substantially radially outwardly from the mounting body and includes an airfoil body. The mounting body and the airfoil body define a plurality of primary cooling passages that extend substantially radially therein for routing the coolant through the airfoil body. Each of the primary cooling passages includes a cooling flow outlet that is formed along a tip portion of the airfoil body. The airfoil body further defines a plurality of trailing edge cooling passages, each having a coolant outlet that is formed along a trailing edge portion of the airfoil body. At least a portion of the trailing edge cooling passages are formed along a radially outer portion of the trailing edge proximate to the tip portion of the airfoil body.

Abstract: A blade has an airfoil, and the blade is configured for use with a turbomachine. The airfoil has a throat distribution measured at a narrowest region in a pathway between adjacent blades, at which adjacent blades extend across the pathway between opposing walls to aerodynamically interact with fluid flow. The airfoil defines the throat distribution, and the throat distribution reduces aerodynamic loss and improves aerodynamic loading on the airfoil. The airfoil has a linear trailing edge profile.

Abstract: A turbine blade includes pressure and suction surfaces connected to define an interior through which coolant is passable. First and second pedestal arrays, each include pedestals respectively coupled to radially outboard portions of respective interior faces of one of the pressure and suction surfaces. The pedestals of the first pedestal array are separated from and directly opposed to pedestals of the second pedestal array by gaps respectively defined therebetween.

Abstract: The present application thus provides a method for reducing stress on at least one of a turbine disk and a turbine blade. The method may include the steps of (a) determining a starting line for a dovetail backcut relative to a datum line, (b) determining a cut angle for the dovetail backcut, and (c) removing material from at least one of the blade dovetail or the disk dovetail slot according to the starting line and the cut angle to form the dovetail backcut. The datum line may be positioned about 2.866 inches (about 72.796 millimeters) from a forward face of the blade dovetail and wherein step (a) is practiced such that for the pressure side of the dovetail, the starting line of the dovetail backcut is at least about 2.566 inches (about 65.176 millimeters) in a forward direction from the datum line.

Abstract: A turbine rotor blade includes a tip portion having a pressure tip wall and a suction tip wall, a tip leading edge and a tip trailing edge, wherein the pressure tip wall and the suction tip wall define a trailing edge tip thickness. Also included is a squealer cavity at least partially defined by the pressure tip wall and the suction tip wall, the squealer cavity including a trench extending fully to the tip trailing edge to form an open flow path out of the tip trailing edge. Further included is a suction side wall and a pressure side wall extending from a root portion of the turbine rotor blade to the tip portion, wherein the suction side wall and the pressure side wall define a trailing edge blade thickness, the trailing edge tip thickness being greater than the trailing edge blade thickness.

Abstract: A rotor blade includes a mounting portion comprising a mounting body that is formed to receive a coolant therein. An airfoil portion extends substantially radially outwardly from the mounting body and includes an airfoil body. The mounting body and the airfoil body define a plurality of primary cooling passages that extend substantially radially therein for routing the coolant through the airfoil body. Each of the primary cooling passages includes a cooling flow outlet that is formed along a tip portion of the airfoil body. The airfoil body further defines a plurality of trailing edge cooling passages, each having a coolant outlet that is formed along a trailing edge portion of the airfoil body. At least a portion of the trailing edge cooling passages are formed along a radially outer portion of the trailing edge proximate to the tip portion of the airfoil body.

Abstract: A bucket tip shroud measurement fixture includes a frame extending from a first end to a second end through an intermediate portion having a first surface and an opposing second surface, a first tip shroud fixing member extending from the second surface of the frame at the first end, and a second tip shroud fixing member extending from the second surface of the frame at the second end. The first and second tip shroud fixing members are configured and disposed to retain and establish an orientation of a bucket tip shroud relative the frame. A plurality of reference points are provided on the frame. The plurality of reference points are configured and disposed to receive a coordinate measuring machine (CMM) probe.

Abstract: A bucket tip shroud measurement fixture includes a frame extending from a first end to a second end through an intermediate portion having a first surface and an opposing second surface, a first tip shroud fixing member extending from the second surface of the frame at the first end, and a second tip shroud fixing member extending from the second surface of the frame at the second end. The first and second tip shroud fixing members are configured and disposed to retain and establish an orientation of a bucket tip shroud relative the frame. A plurality of reference points are provided on the frame. The plurality of reference points are configured and disposed to receive a coordinate measuring machine (CMM) probe.

Abstract: A turbine engine, bucket and method for modifying an airfoil shroud of turbine bucket is disclosed. A reference location is located in a second end edge of the airfoil shroud proximate a seal rail of the airfoil shroud. A relief cut is formed in the airfoil shroud to remove the reference location. Additionally, another reference location may be located in a first end edge of the airfoil shroud proximate the seal rail. Another relief cut may be formed in the airfoil shroud to remove the other reference location.