United Technologies Corporation

Richard H Page

Bryan P Dube

William Abdel-Messeh

Ryan Shepard Levy

Daniel Herrera

Paul M Lutjen

Anthony P Cherolis

Ricardo Trindade

Susan M Tholen

Justin D Piggush

Wieslaw A Chlus

Michael F Blair

Richard W Hoff

Jesse R Christophel

Roger Gates

Karl A Mentz

The present disclosure is directed to a refractory metal core for use in forming varying thickness microcircuits in turbine engine components, a process for forming the refractory metal core, and a process for forming the turbine engine components. The refractory metal core is used in the casting of a turbine engine component. The core is formed by a sheet of refractory metal material having a curved trailing edge portion integrally formed with a leading edge portion.

Process and refractory metal core for creating varying thickness microcircuits for turbine engine components

An investment casting core combination includes a metallic casting core and a ceramic feedcore. A first region of the metallic casting core is embedded in the ceramic feedcore. A mating edge portion of the metallic casting core includes a number of projections. The first region is along at least some of the projections. A number of recesses span gaps between adjacent projections. The ceramic feedcore includes a number of compartments respectively receiving the metallic casting core projections. The ceramic feedcore further includes a number of portions between the compartments and respectively received in the metallic casting core recesses.

Investment casting cores and methods

A gas turbine engine has a turbine blade having a tip cap to close off internal cooling passages. The tip cap is formed with a plurality of purge holes, with there being at least two rows of purge holes. The increased number of purge holes spreads the cooling air outwardly across more of the surface area of the tip cap, and provides the tip cap with a better ability to withstand the extreme temperatures it faces in use.

Turbine blade with improved durability tip cap

A gas turbine engine is provided with an airfoil component that is disclosed as a turbine blade. Internal cooling passages circulate air within the blade. An impingement rib separates a cooling air channel from a pedestal array, and includes cross-over holes to meter the flow of air into the pedestal array. A lower end of the impingement rib is split to reduce stress concentrations between the impingement rib and a platform. Further, the pedestals adjacent to the lower end of the impingement rib are formed to be stubs, which do not extend entirely across a width of the blade. The disclosed structure reduces the stress concentration at an area where the impingement rib meets the platform.

Turbine blade with split impingement rib

A blade outer air seal (BOAS) casting core has first and second end portions and a plurality of legs. Of these legs, first legs each have: a proximal end joining the first end portion; a main body portion; and a free distal portion. Second legs each have: a proximal end joining the second end portion; a main body portion; and a free distal portion.

Blade outer air seal cores and manufacture methods

A turbine blade for use in a gas turbine engine is provided. The turbine blade includes an airfoil portion having a tip end, a shroud attached to the tip end, which shroud has an outer surface, and a knife edge attached to an outer surface of the shroud. The knife edge has a pair of cutter blades disposed substantially over an axis of the airfoil portion.

Shroud honeycomb cutter

A turbine blade having improved cooling characteristics is provided. The turbine blade has an airfoil portion having a span, and at least one cooling passageway in the airfoil portion extending from a root portion of the airfoil portion to a tip portion of the airfoil portion. A plurality of turbulation promotion devices are placed in the at least one cooling passageway. The turbulation promotion devices have a P/e ratio which varies along the span of the airfoil portion, where P is the pitch between adjacent turbulation promotion devices and e is the height of the turbulation promotion devices.

Tailored turbulation for turbine blades

A turbine blade for use in a gas turbine engine is provided. The turbine blade includes an airfoil portion having a tip end, a shroud attached to the tip end, which shroud has an outer surface, and a knife edge attached to an outer surface of the shroud. The knife edge has a pair of cutter blades protruding outwardly from the knife edge.

A turbine blade having improved cooling has an airfoil with a root end and a tip end and at least one cooling passageway in the airfoil. Each cooling passageway extends from the root end to the tip end and has a circular cross-section. A plurality of turbulation promotion devices are arranged in each cooling passageway. Each of the turbulation promotion devices is arcuate in shape and circumscribes an arc less than 180 degrees.

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Guilford, CT, United States of America

Richard H Page