Abstract: There is provided a component formed from a plurality of laminates stacked on one another, thereby defining a stacked laminate structure having a leading edge and a trailing edge. Each of the plurality of laminates is formed from a ceramic matrix composite material. In addition, a plurality of interior cooling channels are defined within an interior of the stacked laminate structure and extend longitudinally between the leading edge and the trailing edge. A metal support structure is arranged so as to extend through first openings in the laminates and through the stacked laminate structure.

Abstract: A gas turbine engine blade (22), including an airfoil substrate (10) having an exterior surface, wherein: a base (14) of the airfoil substrate is located at a 0% radial on an inner platform surface (20) and a tip (16) of the airfoil substrate is located at a 100% radial; wherein at the 0% radial a cross-sectional profile of the exterior surface is substantially characterized by nominal X and Y coordinates present in Table 1; and wherein at a 50% radial location a cross-sectional profile of the exterior surface is characterized by nominal X and Y coordinates present in Table 6.

Abstract: A gas turbine engine blade (10), having: an airfoil (12) having a pressure side (22) and a suction side (24); an inner platform (30), having: a gas path surface (40); a coolant surface (42); and at least one row (114) of film cooling holes (112) spanning there between and disposed on the pressure side of the airfoil. The film cooling holes are angled to include a directional component substantially parallel with the pressure side at a trailing end of the airfoil, and the at least one row is oriented substantially parallel to a mateface wall (122) of the inner platform. An array (130) of turbulators (132) is disposed on the coolant surface.

Abstract: A gas turbine engine blade (22), including an airfoil substrate (10) having an exterior surface, wherein: a base (14) of the airfoil substrate is located at a 0% radial on an inner platform surface (20) and a tip (16) of the airfoil substrate is located at a 100% radial; wherein at the 0% radial a cross-sectional profile of the exterior surface is substantially characterized by nominal X and Y coordinates present in Table 1; and wherein at a 50% radial location a cross-sectional profile of the exterior surface is characterized by nominal X and Y coordinates present in Table 6.

Abstract: A method and device for facilitating measurement of thermo-optically induced material phase change response in a thin planar or a grating film stack is disclosed. The method may include using small-spot visible and ultraviolet spectra (ellipsometric or reflectance) for measuring a material phase change response. The device may include a measurement system platform, at least one electrical resistor, at least one external electric probe, and ohmic contact circuitry.

Abstract: A method and device for facilitating measurement of thermo-optically induced material phase change response in a thin planar or a grating film stack is disclosed. The method may include using small-spot visible and ultraviolet spectra (ellipsometric or reflectance) for measuring a material phase change response. The device may include a measurement system platform, at least one electrical resistor, at least one external electric probe, and ohmic contact circuitry.

Abstract: A turbine vane usable in a turbine engine and having at least one cooling system. The cooling system including an aft cooling circuit formed from at least one serpentine cooling path. The serpentine cooling path having at least one rib may include bypass orifices for allowing air to pass through the rib to shorten the distance of the serpentine cooling path through which at least some of the air passes. The bypass orifices allow a greater quantity of air to pass through the vane and be expelled into a disc to which the vane is movably coupled as compared to a similar shaped and sized serpentine cooling path not having the bypass orifices.

Abstract: An apparatus 100 for removing surface non-uniformities is provided. This apparatus has a stage 103 for holding a substrate 127 to be processed. This substrate often includes a film thereon, where the film has the non-uniformities. The apparatus 200 includes a movable head 111, which can provide rotatable movement about a fixed axis 123. A drive motor 115 is operably attached to the movable head 111 to provide this rotatable movement. A pad 113 (e.g., polishing or planarizing pad) is attached to the movable head. This pad 113 comprises an abrasive material and also has a smaller length (e.g, diameter, etc.) relative to a length (e.g, diameter, etc.) of the substrate. The smaller pad is capable of selectively removing a portion of the non-uniformities on the film.

Abstract: A rotor blade for a rotor assembly is provided which includes a root, an airfoil, a platform, and a damper. The airfoil includes at least one cavity. The platform extends laterally outward from the blade between the root and the airfoil, and includes an airfoil side, a root side, and an aperture extending between the root side of the platform and the cavity within the airfoil. The damper, which includes at least one bearing surface, is received within the aperture and the cavity. The bearing surface is in contact with a surface within the cavity and friction between the bearing surface and the surface within the cavity damps vibration of the blade.

Abstract: An internally cooled turbine blade for a gas turbine engine is modified at the leading and trailing edges to include a dynamic cool air flowing radial passageway with an inlet at the root and a discharge at the tip feeding a plurality of radially spaced film cooling holes in the airfoil surface. Replenishment holes communicating with the serpentine passages radially spaced in the inner wall of the radial passage replenish the cooling air lost to the film cooling holes. The discharge orifice is sized to match the backflow margin to achieve a constant film hole coverage throughout the radial length.

Abstract: A rotor blade for a rotor assembly is provided comprising a root, an airfoil, and a damper. The airfoil includes a base, a tip, a first cavity, a second cavity, and a passage. The passage includes a pair of walls converging from the first cavity to the second cavity, thereby connecting the first and second cavities. The damper is received within the passage. According to one aspect of the present invention, a difference in gas pressure across the damper biases the damper against the converging walls of the passage. According to another aspect of the present invention, centrifugal force biases the damper against the converging walls of the passage.