Abstract: A method for orienting cooling holes of a nozzle singlet for a turbine engine is provided. The method includes providing a nozzle singlet having an inner band, an outer band, and at least one airfoil extending therebetween. The method also includes orienting at least one first row of cooling holes an angle with respect to at least one second row of cooling holes. The orientation of the at least one first row and the at least one second row provides a cooling hole pattern that accommodates a change in the airfoil angle without reorienting the cooling hole pattern.

Abstract: A method for orienting cooling holes of a nozzle singlet for a turbine engine is provided. The method includes providing a nozzle singlet having an inner band, an outer band, and at least one airfoil extending therebetween. The method also includes orienting at least one first row of cooling holes an angle with respect to at least one second row of cooling holes. The orientation of the at least one first row and the at least one second row provides a cooling hole pattern that accommodates a change in the airfoil angle without reorienting the cooling hole pattern.

Abstract: An airfoil for a turbine nozzle assembly of a gas turbine engine includes an outer side wall, an inner side wall, a leading edge extending from the outer side wall to the inner side wall, a trailing edge extending from the outer side wall to the inner side wall, a concave surface extending from the leading edge to the trailing edge on a pressure side of the airfoil, a convex surface extending from the leading edge to the trailing edge on a suction side of the airfoil, an outer cooling slot, an inner cooling slot, and at least one middle cooling slot formed in the concave side of the airfoil adjacent the trailing edge.

Abstract: An airfoil for a turbine nozzle assembly of a gas turbine engine includes an outer side wall, an inner side wall, a leading edge extending from the outer side wall to said inner side wall, a trailing edge extending from the outer side wall to the inner side wall, a concave surface extending from the leading edge to the trailing edge on a pressure side of the airfoil, a convex surface extending from the leading edge to the trailing edge on a suction side of said airfoil, an outer cooling slot, an inner cooling slot, and at least one middle cooling slot formed in the concave side of the airfoil adjacent the trailing edge.

Abstract: A method is provided for making, from a plurality of members brazed together, an article including an environmental resistant surface coating and a wear resistant surface portion. Prior to brazing, the members are assembled with at least one preform including the wear resistant material in a matrix including a first brazing alloy having a brazing temperature in a brazing temperature range. The assembly of members includes a second brazing alloy having a brazing temperature in the brazing temperature range. The assembly of members and wear resistant preform is heated in the brazing temperature range to provide a brazed article preform. Then the article preform is machined to a selected geometry and can be coated with the environmental coating.

Abstract: A gas turbine engine arcuate segment has an arcuate band segment, at least one tab and an arcuate rail segment spaced axially apart from the tab and located along an axial end of the band segment. The tab and the arcuate rail segment extend radially away from the band segment and a space is located between the tab and the rail segment. A tab aperture extends entirely through the tab. The rail segment has a rail aperture entirely therethrough. A leaf seal and a spring are disposed within the space to bias the leaf seal against the aft rail. In an exemplary embodiment, a spring aperture extends entirely through the spring, a seal aperture extends entirely through the leaf seal, and a headless pin is disposed in the apertures such that the leaf seal and the spring are radially restrained by the pin.

Abstract: A turbine nozzle includes outer and inner bands between which extend a plurality of vanes for channeling combustion gases. Each of the vanes includes leading and trailing edges, and pressure and suction sides extending therebetween, and also a bow along the trailing edge to increase pressure in the gases adjacent the inner band. The vanes also include a thermal barrier coating (TBC) selectively disposed solely along the suction side between the leading and trailing edges.

Abstract: A turbine nozzle includes outer and inner bands between which extend a plurality of vanes for channeling combustion gases. Each of the vanes includes leading and trailing edges, and pressure and suction sides extending therebetween, and also a bow along the trailing edge to increase pressure in the gases adjacent the inner band. The vanes also include a thermal barrier coating (TBC) selectively disposed solely along the suction side between the leading and trailing edges.