Abstract: An air-cooled turbine blade and methods of manufacturing the blade are provided. The blade includes a suction side flow circuit formed within its interior and defined at least by an interior surface of a convex suction side wall, a pressure side flow circuit formed within the blade interior and defined at least by an interior surface of a concave pressure side wall, and a center flow circuit including a first section and a second section, the first section disposed between the suction side flow circuit and the pressure side flow circuit, and the second section in flow communication with the first section and a plurality of openings of a leading edge wall and defined at least partially by an interior surface of the leading edge wall.

Abstract: A turbine blade includes a convex suction side wall, a concave pressure side wall, a tip wall, an internal cooling circuit, and a plurality of tip edge channels. The tip wall is recessed from a first tip edge of the suction side wall and a second tip edge of the concave pressure side wall to define a suction side wall tip section and a pressure side wall tip section, and the suction side wall tip section is shorter than the pressure side wall tip section. The internal cooling circuit is formed at least partially between the convex suction side wall, the concave pressure side wall, and the tip wall. The plurality of tip edge channels formed through the first tip edge of the convex suction side wall extend to the internal cooling circuit. Methods of manufacturing turbine blades are also provided.

Abstract: An air-cooled turbine blade and methods of manufacturing the blade are provided. The blade includes a suction side flow circuit formed within its interior and defined at least by an interior surface of a convex suction side wall, a pressure side flow circuit formed within the blade interior and defined at least by an interior surface of a concave pressure side wall, and a center flow circuit including a first section and a second section, the first section disposed between the suction side flow circuit and the pressure side flow circuit, and the second section in flow communication with the first section and a plurality of openings of a leading edge wall and defined at least partially by an interior surface of the leading edge wall.

Abstract: A turbine blade includes a convex suction side wall, a concave pressure side wall, a tip wall, an internal cooling circuit, and a plurality of tip edge channels. The tip wall is recessed from a first tip edge of the suction side wall and a second tip edge of the concave pressure side wall to define a suction side wall tip section and a pressure side wall tip section, and the suction side wall tip section is shorter than the pressure side wall tip section. The internal cooling circuit is formed at least partially between the convex suction side wall, the concave pressure side wall, and the tip wall. The plurality of tip edge channels formed through the first tip edge of the convex suction side wall extend to the internal cooling circuit. Methods of manufacturing turbine blades are also provided.

Abstract: A turbine blade tip and shroud clearance control coating system comprising an abrasive blade tip coating and an abradable shroud coating are provided. The abrasive layer may comprise abrasive particles of cubic zirconia, cubic hafnia or mixtures thereof, and the abradable layer may be a nanolaminate thermal barrier coating that is softer than the abrasive layer. The invention further provides an alternate coating system comprising an abradable blade tip coating and an abrasive shroud coating.

Abstract: A turbine blade tip and shroud clearance control coating system comprising a dense abrasive blade tip layer and an abradable shroud layer are provided. The dense abrasive coating may comprise cubic zirconia, hafnia or mixtures thereof and the abradable layer may be a nanolaminate thermal barrier coating that is softer than the dense abrasive layer.

Abstract: Methods for rapid prototype casting metal components, wherein the metal components are cast in a secondary ceramic mold, the secondary ceramic mold is cast in a primary mold, and the primary mold is formed by rapid prototyping or rapid manufacturing. The secondary ceramic mold may comprise a one-piece integral shell and core(s), and the metal components may have at least one hollow portion or void therein, such as a hollow airfoil for a gas turbine engine.

Abstract: A turbine blade tip and shroud clearance control coating system comprising an abrasive blade tip coating and an abradable shroud coating are provided. The abrasive layer may comprise abrasive particles of cubic zirconia, cubic hafnia or mixtures thereof, and the abradable layer may be a nanolaminate thermal barrier coating that is softer than the abrasive layer. The invention further provides an alternate coating system comprising an abradable blade tip coating and an abrasive shroud coating.