United Technologies Corporation

Siemens Energy, Inc.

Siemens Westinghouse Power Corporation

Robert J McClelland

Robert J Kraft

Thomas A Auxier

Kenneth B Hall

Edwin Lee Kite

Charles M Evans

Eric Munoz

A gas turbine engine blade (), including an airfoil substrate () having an exterior surface, wherein: a base () of the airfoil substrate is located at a 0% radial on an inner platform surface () and a tip () 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.

Gas turbine blade configuration

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.

Cooling system for a turbine vane

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.

Airfoil vibration damping device

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. Trip strips may be employed to augment the pressure drop distribution

Cooled blades for a gas turbine engine

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.

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Palm City, FL, United States of America

Robert J McClelland