United Technologies Corporation

General Electric Company

William S Kvasnak

Ronald S LaFleur

Friedrich O Soechting

Douglas A Hayes

Christopher R Joe

Joe Moroso

James Preston Downs

Thomas A Auxier

Scott Edmond Ellis

James Anthony West

Hans R Przirembel

Brendan F Sexton

Kenneth K Landis

Samuel D Draper

William H Calhoun

The present application provides a hot gas path component for use in a hot gas path of a gas turbine engine. The hot gas path component may include an internal wall, an external wall facing the hot gas path, an impingement wall, a number of internal wall pedestals positioned between the internal wall and the impingement wall, and a number of external wall pedestals positioned between the external wall and the impingement wall.

Hot gas path component with impingement and pedestal cooling

A turbine nozzle cooling sub-system is provided. The sub-system includes at least one turbine nozzle segment. The segment includes an arcuate, radially outermost endwall, an arcuate, radially innermost endwall, and at least one airfoil vane. The endwalls each include at least one open passage. The airfoil vane extends between and is coupled to the endwalls. The vane further includes a cavity, a leading edge, a trailing edge, and an airfoil vane external surface. The cavity includes an airfoil vane internal surface and a plurality of turbulators. The cavity and the open passages are in flow communication such that an airfoil cooling air stream flow is facilitated. The sub-system also includes at least one diffuser in flow communication with a compressor assembly and the segment. The diffuser includes at least one diffuser wall and cavity. The diffuser wall extends from the compressor assembly to the segment such that a channeling of the airfoil cooling air stream to the segment is facilitated. The airfoil cooling air stream includes at least a portion of a compressor assembly discharge air stream flow.

Methods and apparatus for cooling combustion turbine engine components

A turbine engine component, such as a blade or a vane, is provided by the present invention. The turbine engine component has a pressure side and a suction side. Each of the pressure and suction sides has an external wall and an internal wall. A first set of fluid passageways is located on the pressure side between the external wall and the internal wall. A second set of fluid passageways is located on the suction side between the external wall and the internal wall. Each of the fluid passageways in the first set and in the second set has a wavy configuration. The turbine engine component may also have one or more wavy trailing edge cooling passageways for cooling a trailing edge portion of the component.

Drillable super blades

An internally-cooled turbomachine element has an airfoil extending between inboard and outboard ends. A cooling passageway is at least partially within the airfoil and has at least a first turn. Means are in the passageway for limiting a turning a loss of the first turn. The turbomachine element may result from a reengineering of an existing element configuration lacking such means.

Cooling passageway turn

A cooling circuit is provided disposed between a first wall portion and a second wall portion of a wall for use in a gas turbine engine, one or more inlet apertures, and one or more exit apertures. The inlet aperture(s) provides a cooling airflow path into the cooling circuit and the exit aperture(s) provides a cooling airflow path out of the cooling circuit. The cooling circuit includes a plurality of first pedestals extending between the first wall portion and the second wall portion. The first pedestals are arranged in one or more rows.

Method and apparatus for cooling a wall within a gas turbine engine

A method and apparatus for cooling a wall within a gas turbine engine is provided which comprises the steps of: (1) providing a wall having an internal surface and an external surface; (2) providing a cooling microcircuit within the wall that has a passage for cooling air that extends between the internal surface and the external surface; and (3) increasing heat transfer from the wall to a fluid flow within the passage by increasing the average heat transfer coefficient per unit flow within the microcircuit. According to one aspect, the present invention method and apparatus can be tuned to substantially match the thermal profile of the wall at hand.

Method and apparatus for cooling an airfoil

A coolable airfoil is disclosed which includes an internal cavity, an external wall, a plurality of first apertures, and a plurality of second apertures. The external wall includes a suction-side portion and a pressure-side portion. The external wall portions extend chordwise between a leading edge and a trailing edge, and spanwise between an inner radial surface and an outer radial surface. The mean camber line of the airfoil passes through the leading edge and the trailing edge along a path equidistant between the outer surfaces of the pressure-side and suction-side walls. The first apertures, which are disposed in the external wall adjacent the trailing edge, extend a distance within the suction-side wall portion and exit the external wall through the pressure-side wall portion. The second apertures extend through the pressure-side wall portion and exit the pressure-side wall portion upstream of and in close proximity to the first apertures.

Trailing edge cooling apparatus for a gas turbine airfoil

A coolable airfoil is provided which includes an internal cavity, an external wall, a plurality of first cooling apertures, and a plurality of second cooling apertures. The external wall includes a suction side portion and a pressure side portion. The wall portions extend chordwise between a leading edge and a trailing edge and spanwise between an inner radial surface and an outer radial surface. The first cooling apertures are disposed in the external wall adjacent the trailing edge, exiting the airfoil through the pressure side portion. The second cooling apertures are disposed in the external wall adjacent the trailing edge, exiting the airfoil through the suction side portion.

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Guilford, CT, United States of America

William S Kvasnak