Abstract: An apparatus for a nozzle assembly for a turbine engine and a method of forming the nozzle assembly are described herein. The nozzle assembly can include a set of nozzles including an inner band and an outer band, with one or more airfoils extending between the inner and outer bands. The nozzles can be coupled to adjacent nozzles at the inner and outer bands to form the nozzle assembly. A throat can be defined at the minimum cross-sectional area between adjacent airfoils in the nozzle assembly. A formed portion can be formed on either of or both of the inner or outer band at the throat.

Abstract: A method for hole formation using an electrical discharge machining tool configured to penetrate metal. The method includes the steps of: penetrating the metal using an electrode to form a hole by moving the electrode away from a starting point; verifying that the hole is complete using the electrical discharge machine; and wherein the verifying step includes the step of probing the hole with the electrode.

Abstract: A method for hole formation using an electrical discharge machining tool configured to penetrate metal. The method includes the steps of: penetrating the metal using an electrode to form a hole by moving the electrode away from a starting point; verifying that the hole is complete using the electrical discharge machine; and wherein the verifying step includes the step of probing the hole with the electrode.

Abstract: A turbine nozzle includes a row of vanes extending radially between annular outer and inner bands. The outer band includes a pair of radial flanges defining an annular seal groove therebetween. One of the flanges is crenelated to improve nozzle life.

Abstract: A turbine nozzle includes a row of vanes extending radially between annular outer and inner bands. The outer band includes a pair of radial flanges defining an annular seal groove therebetween. One of the flanges is crenelated to improve nozzle life.

Abstract: A method for measuring a throat area of a turbine nozzle assembly that includes at least one airfoil extending between an inner band and an outer band is provided. The method includes locating a plurality of datum points on the at least one airfoil using a measuring system, measuring a plurality of measurement points within a flow path at least partly defined within the turbine nozzle assembly using the measuring system, comparing the plurality of datum points to a plurality of corresponding datum points measured on a baseline turbine nozzle assembly, and calculating a throat area variance between the measured turbine nozzle assembly and the baseline turbine nozzle assembly model based on the comparison to the baseline turbine nozzle assembly.

Abstract: The fourth-stage buckets have airfoil profiles substantially in accordance with Cartesian coordinate values of X, Y and Z set forth in inches in Table I wherein Z is a perpendicular distance from a plane normal to a radius of the turbine centerline and containing the X and Y values with the Z value commencing at zero in the X, Y plane at the radially innermost aerodynamic section of the airfoil and X and Y are coordinate values defining the airfoil profile at each distance Z. The X, Y and Z values may be scaled as a function of the same constant or number to provide a scaled-up or scaled-down airfoil section for the bucket.

Abstract: The fourth-stage buckets have airfoil profiles substantially in accordance with Cartesian coordinate values of X, Y and Z set forth in inches in Table I wherein Z is a perpendicular distance from a plane normal to a radius of the turbine centerline and containing the X and Y values with the Z value commencing at zero in the X, Y plane at the radially innermost aerodynamic section of the airfoil and X and Y are coordinate values defining the airfoil profile at each distance Z. The X, Y and Z values may be scaled as a function of the same constant or number to provide a scaled-up or scaled-down airfoil section for the bucket.

Abstract: An interface seal for a turbine engine including a rotor shroud hanger including at least one opening, an outer band including an aft rail, and a seal spring mounted in the opening and extending over an interface between the outer band and the shroud hanger, is described. The seal provides a barrier that prevents the leakage of high pressure cooling air through the interface into the combustion gas flowpath bypassing turbine engine components, and thus forces the high pressure cooling air through the necessary turbine engine components.

Abstract: A turbine seal includes a first arcuate segment defining a flowpath boundary between combustion gases and air, and includes a radially outwardly extending rail at one end thereof. A second arcuate segment is disposed coaxially with the first segment for defining a continuation of the flowpath boundary, and has a radially extending face adjoining the rail. A leaf seal bridges the rail and the face for sealing leakage therebetween, and a plurality of pins extend through the leaf seal for providing mounting to the rail. Each pin includes a head and an opposite tip, with the pins being fixedly joined to the rail solely about the pin tips for freely supporting the pin heads in a cantilever without obstruction therearound for eliminating supporting tabs. A leaf spring is mounted on the pins between the heads and the leaf seal to pre-load the leaf seal against the rail and face to effect the sealing therebetween.

Abstract: A method and apparatus are disclosed for treating ice and debris in a gas turbine engine. More particularly, a specially configured and slidable scoop is positioned between the booster exit and the compressor inlet in a gas turbine engine. An actuator mechanism can be coupled through a linkage to the scoop to position the scoop within the core path at low power conditions and retract the scoop from the core flow path at high power conditions.