General Electric Company

Leslie Eugene Leeke, Jr

Sean Robert Keith

Michael Joseph Danowski

David Alan Frey

Ronald Eugene Mc Rae, Jr

Richard William Albrecht, Jr

Ronald Eugene McRae, Jr

Michael Beverley

Lawrence Paul Timko

Ching-Pang Lee

Brian Keith

Dean Thomas Lenahan

Chander Prakash

Kevin Samuel Klasing

Mark Edward Stegemiller

Robert Ingram Ackerman

Daniel John Harris

Andrew Edward Obermeyer

An apparatus enables a turbine nozzle, including at least two turbine nozzle singlets to be fabricated. Each singlet includes inner and outer bands and a vane extending therebetween. The vane also includes first and second sidewalls coupled together at a leading edge and a trailing edge. The apparatus comprises a fixture, and at least two support members extending from the fixture. Each support member has first and second abutment surfaces, wherein a first of the at least two support members contacts at least one of the leading and trailing edges of a first of the vanes. At least two locating features extend from the fixture, wherein a first of the locating features contacts one of the first and second sidewalls of the first vane. At least two biasing members are coupled to the fixture, wherein a first of the biasing members biases the first vane against the first locating feature.

Methods and apparatus for fabricating a turbine nozzle assembly

A rotor blade assembly includes a shank, an airfoil that is formed integrally with the shank, and a removable platform coupled between said shank and said airfoil via a friction fit. A method of assembling a gas turbine engine rotor blade assembly that includes a removable platform, and a gas turbine engine rotor assembly including the removable platform are also described herein.

Turbine rotor blade and method of assembling the same

A removable rotor blade platform includes a first platform leg, a second platform leg, and a platform portion coupled to the first and second platform legs. The first platform leg is configured to be retained by a first retainer coupled to a first rotor blade, and the second platform leg is configured to be retained by a second retainer coupled to a second adjacent rotor blade. A method of assembling a blade assembly that includes a removable platform, a rotor assembly including the removable platform, and a gas turbine engine assembly including the removable platform, are also described herein.

Turbine rotor blade assembly and method of assembling the same

An airfoil for a rotor blade including an uncoated profile substantially in accordance with Cartesian coordinate values of X, Y and Z to facilitate balancing performance and durability of the rotor blade and to facilitate improving an operating efficiency of a high-pressure turbine is provided. The profile is carried only to four decimal places, wherein Y represents a distance from a platform on which the airfoil is mounted, and X and Z are coordinates defining the profile at each distance Y from the platform.

Rotor blade profile optimization

An airfoil for a gas turbine engine includes a root, a tip, a leading edge, a trailing edge, and opposed pressure and suction sidewalls extending generally along a radial axis. The airfoil includes a tip cap extending between the pressure and suction sidewalls; and spaced-apart suction-side and pressure-side tip walls extending radially outward from the tip cap to define a tip cavity therebetween. The pressure-side tip wall includes a continuously concave curved arcuate portion, at least a section of which extends circumferentially outward from a radial axis of the airfoil.

Turbine airfoil with curved squealer tip

A method for fabricating a turbine rotor blade includes casting a turbine rotor blade including a dovetail, a platform having an outer surface, an inner surface, and a cast-in plenum defined between the outer surface and the inner surface, and an airfoil, and forming a plurality of openings between the platform inner surface and the platform outer surface to facilitate cooling an exterior surface of the platform.

Method and apparatus for cooling gas turbine engine rotor blades

A method for fabricating a rotor blade includes casting the turbine rotor blade to include a shank, and a platform having an upper surface and a lower surface, and coupling a first component to the rotor blade such that a first substantially hollow plenum is defined between the first component, the shank, and the platform lower surface.

A hollow cooled turbine airfoil is provided having pressure and suction side walls and a plurality of trailing edge cooling passages that feed cooling air bleed slots at the trailing edge. The suction side wall has a selectively reduced thickness, which allows shortened trailing edge slots, improving trailing edge cooling.

Turbine airfoil for gas turbine engine

A turbine includes a row of blades each having an integral airfoil, platform, and dovetail. Each platform has opposite first and second side edges corresponding with the opposite pressure and suction sides of the airfoil. The platform first side edge is radially higher than the platform second side edge continuously between opposite forward and aft edges of the platform. In this way, adjacent platforms define a down step therebetween for preventing obstruction of combustion gases flowable downstream thereover.

Step-down turbine platform

A hollow cooled turbine airfoil is provided having pressure and suction side walls and a plurality of trailing edge cooling passages that feed cooling air bleed slots at the trailing edge. The airfoil trailing edge is selectively thickened in the root portion so as to allow shortened trailing edge slots, thereby improving trailing edge cooling and reducing mechanical and thermal stresses.

A dovetail assembly including non-parallel relief faces that facilitates reduced pressure face brinelling in turbine engines. The assembly includes a plurality of rotor blades, each including a dovetail. Each dovetail includes at least a pair of blade tangs including blade relief faces. The dovetail assembly also includes a rotor disk including a plurality of dovetail slots, each sized to receive a dovetail. Each dovetail slot is defined by at least one pair of opposing disk tangs including disk relief faces. The disk relief faces are non-parallel to the blade relief faces when the dovetail is mounted in the dovetail slot.

Method and apparatus for non-parallel turbine dovetail-faces

A gas turbine engine blade includes pressure and suction sides extending between leading and trailing edges and root to tip. The pressure side includes a tip rib recessed therein to define a tip shelf terminating in an inclined ramp. The ramp may be aligned with streamlines of combustion gas flow for preventing interruption thereof and the increase of heat transfer therefrom.

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Burlington, KY, United States of America

Leslie Eugene Leeke, Jr