United Technologies Corporation

Raytheon Technologies Corporation

Yafet Girma

Brandon S Donnell

Thomas N Slavens

James Tilsley Auxier

Dominic J Mongillo, Jr

San Quach

Steven Bruce Gautschi

Scott D Lewis

Christopher Corcoran

Anita L Tracy

Jason L O'Hearn

Gina C O'Dell

In a featured embodiment, a gas turbine engine component comprises an airfoil having a leading edge, a trailing edge, and pressure and suction side walls extending from the leading edge to the trailing edge. The airfoil extends from a base to a tip. A shelf is formed in the tip, and extends from the pressure side wall, around the leading edge, to the suction side wall.

Enhanced cooling for blade tip

An airfoil according to an exemplary aspect of the present disclosure includes, among other things, an airfoil section having an external wall and an internal wall. The internal wall defines a first reference plane extending in a spanwise direction and through a thickness of the internal wall. A first cavity and a second cavity are separated by the internal wall. A plurality of crossover passages within the internal wall connects the first cavity to the second cavity. Each of the plurality of crossover passages defines a passage axis. The plurality of crossover passages are distributed in the spanwise direction and arranged such that the passage axis of each of the plurality of cooling passages intersects a surface of the second cavity. The plurality of crossover passages include a first set of crossover passages and a second set of crossover passages positioned on opposite sides of the first reference plane. The passage axis of each of the first set of crossover passages is arranged at a first vertical angle relative to a spanwise axis, and the passage axis of each of the second set of crossover passages is arranged at a second, different vertical angle relative to the spanwise axis. A casting core for an airfoil is also disclosed.

Staggered crossovers for airfoils

A component for a gas turbine engine according to an exemplary aspect of this disclosure includes a surface having a plurality of cooling holes. The surface includes a first region and a second region each having a plurality of cooling holes. The cooling holes within the first region are arranged differently than the cooling holes in the second region.

Cooling hole arrangement for engine component

An airfoil for a gas turbine engine includes pressure and suction walls spaced apart from one another and joined at leading and trailing edges to provide an airfoil having an exterior surface that extends in a radial direction to a tip. A camber line at the tip extends from the leading edge to the trailing edge. Pressure and suction side shelves are arranged in the exterior surface on opposing sides of the camber line respectively in the pressure and suction side walls. A plateau is proud of and separates the pressure and suction side shelves. The plateau is arranged along the camber line and extends to the leading edge.

Gas turbine engine turbine blade tip cooling

A damper body extends between an axial first end, an opposing axial second end, a first lateral side, and an opposing second lateral side. A first pair of lugs are disposed at the first axial end and a second pair of lugs are disposed at the second axial end. In the first lug pair, a lug has an axially outermost foot which extends laterally beyond a lateral wall. The second lug pair has a lug with an axial outermost point with a foot extending laterally beyond a lateral wall. The foot on the first axial end and the foot on the second axial end are on opposed lateral sides.

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Hartford, CT, United States of America

Yafet Girma