Abstract: A method for forming a passage in a ceramic matrix composite component incudes forming a core for a ceramic matrix composite component; embedding a hollow member into the core at a desired location for a passage in the ceramic matrix composite component; wrapping the core with a ceramic material; and inserting a rod through the hollow member and into the core.

Abstract: A rotor blade damper is provided that includes a body having a base, a tip, a first contact surface, a second contact surface, a trailing edge surface, and a leading edge surface. The trailing edge and the leading edge surfaces extend between the contact surfaces. The first contact surface, second contact surface, trailing edge surface, and leading edge surface all extend lengthwise between the base and the tip. The body includes at least one cooling aperture disposed adjacent the base, that has a diameter that is approximately equal to or greater than the width of the trailing edge surface adjacent the tip. The body tapers between the base and the tip such that a first widthwise cross-sectional area adjacent the base is greater than a second widthwise cross-sectional area adjacent the tip.

Abstract: A rotor blade for a rotor assembly is provided that includes a root, an airfoil, and a damper. The airfoil has a length, a base, a tip, a first side wall, a second side wall, and at least one cavity. The length extends the base and the tip. The at least one cavity is disposed between the side walls, and the channel is defined by a first wall portion and a second wall portion. The damper, which is selectively received within the channel, includes a first bearing surface, a second bearing surface, a forward surface, and an aft surface, all of which extend lengthwise. At least one of the surfaces is shaped to form a lengthwise extending passage within the channel. The passage has a flow direction oriented along the length of the at least one surface to permit cooling air travel along the at least one surface in a lengthwise direction. According to one aspect of the present invention, the damper has an arcuate lengthwise extending centerline.

Abstract: A rotor blade for a rotor assembly is provided that includes a root, an airfoil, a platform, and a damper. The airfoil has at least one cavity. The platform is disposed between the root and the airfoil. The platform includes an inner surface, an outer surface, and a damper aperture disposed in the inner surface. The damper has a body and a base. The base and the damper aperture have mating geometries that enable the base to rotate within the damper aperture without substantial impediment from the mating geometries.

Abstract: A rotor blade is provided having a hollow airfoil with a cavity and an attached root. The root has a leading edge conduit, at least one mid-body conduit, and a trailing edge conduit. Each conduit has a centerline. The leading edge conduit includes an inlet having a forward side, a suction side, and a pressure side that diverge from the centerline of the leading edge conduit, and an aft side. Each of the mid-body conduits includes an inlet having a suction side and a pressure side that diverge from the centerline of the mid-body conduit, and an aft side and a forward side. The trailing edge conduit includes an inlet having a suction side and a pressure side that diverge from the centerline of the trailing edge conduit. The trailing edge conduit inlet further includes a forward side and an aft side.

Abstract: The present invention relates to a method for repairing a turbine blade having a crack in a trailing edge portion of the blade. The method comprises the steps of cutting back a first surface of the turbine blade adjacent the blade trailing edge portion where the crack is located, and cutting back a second surface of the turbine blade adjacent the blade trailing edge portion where the crack is located. Each cut back step comprises cutting back the respective surface by a depth greater than the length of the crack and less than the trailing edge radius to remove the crack and form a cut back trailing edge portion. A compound radius is used to prevent a blunt transition into the trailing edge that would result in aerodynamic losses and to reduce the airfoil root stresses. The method also includes applying a thermal barrier coating to the turbine blade to increase service life.

Abstract: A method for repairing a turbine blade having a crack in a trailing edge portion of the blade includes the steps of cutting back a first surface of the turbine blade adjacent the blade trailing edge portion where the crack is located, and cutting back a second surface of the turbine blade adjacent the blade trailing edge portion where the crack is located. Each cut back step comprises cutting back the respective surface by a depth greater than the length of the crack and less than the trailing edge radius to remove the crack and form a cut back trailing edge portion. A compound radius is used to prevent a blunt transition into the trailing edge that would result in aerodynamic losses and to reduce the airfoil root stresses. The method also includes applying a thermal barrier coating to the turbine blade to increase service life. Prior to applying the coating, the tip portion of the turbine blade is modified to account for the change in the thermal characteristics of the turbine blade.