United Technologies Corporation

Raytheon Technologies Corporation

Rtx Corporation

Steven Bruce Gautschi

San Quach

Tracy A Propheter-Hinckley

Dominic J Mongillo, Jr

Matthew A Devore

Brandon W Spangler

Mark F Zelesky

Ricardo Trindade

Thomas N Slavens

Shawn M McMahon

Christopher King

Justin M Aniello

Ryan Alan Waite

Edward F Pietraszkiewicz

Scott D Lewis

Nicholas M LoRicco

Atul Kohli

Benjamin F Hagan

Lane Mikal Thornton

Christopher Corcoran

Mohamed Hassan

Robin Prenter

Ryan Lundgreen

David Barger

Nicholas J Madonna

Christopher Cosher

Brett Alan Bartling

Christopher Perron

Benjamin Bellows

James B Hoke

Dennis M Moura

Alexander Staroselsky

Thomas J Martin

Benjamin T Fisk

Dustin Wayne Davis

Christopher Whitfield

Richard M Salzillo

David J Hyland

David Donald Chapdelaine

Trevor Rudy

Thomas Allwood

Vincent E Simms

Jeffery Scott Hembree, Sr

Stephen Karl Kramer

James Tilsley Auxier

Matthew S Gleiner

Dave J Hyland

Lea Kennard Castle

Anita L Tracy

David J Candelori

Shayan Ahmadian

Brandon S Donnell

Sean D Bradshaw

Jaime G Ghigliotty

Yafet Girma

Christina E Botnick

Stephanie Santoro

Tracy A Propheter-Hincley

Jr Dominic J Mongillo

Benjamin D Bellows

Tracy A Propheter-Hinkley

Christoper King

Sr Jeffery Scott Hembree

Edward F Pietrasziewicz

A component for a gas turbine engine includes, among other things, an airfoil that includes a pressure sidewall and a suction sidewall that meet together at both a leading edge and a trailing edge, the airfoil extending radially from a platform to a tip, a tip pocket formed in the tip and terminating prior to the trailing edge, and one or more heat transfer augmentation devices formed in the tip pocket.

Airfoil tip pocket with augmentation features

A component for a gas turbine engine includes a trailing edge tip corner that at least partially defines a trailing edge cavity and a multiple of corner features within the trailing edge cavity, the multiple of corner features splayed along the trailing edge tip corner.

Splayed tip features for gas turbine engine airfoil

A vane ring for a gas turbine engine includes a multiple of vanes that extend between the inner vane platform and the outer vane platform, each of the multiple of vanes contains an airfoil cooling circuit that receives cooling airflow through a respective one of a multiple of feed passages; a multiple of metering passages in the outer vane platform, each of the multiple of metering passages in communication with the respective one of the multiple of feed passages; and a multiple of secondary passages in the outer vane platform, each of the multiple of secondary passages in communication with the respective one of the multiple of metering passages.

Redundant entry cooling air feed hole blockage preventer for a gas turbine engine

A vane ring for a gas turbine engine component includes a multiple of vanes that extend between an inner vane platform and an outer vane platform, each of the multiple of vanes contains an airfoil cooling circuit that receives cooing airflow through a feed passage located in the outer vane platform and an extension from the outer vane platform, the extension comprises a metering passage in communication with the feed passage.

Shape recessed surface cooling air feed hole blockage preventer for a gas turbine engine

A vane ring for a gas turbine engine component includes a multiple of vanes that extend between an inner vane platform and an outer vane platform, each of the multiple of vanes contains an airfoil cooling circuit that receives cooling airflow through a feed passage located in the outer vane platform and an extension from the outer vane platform, the extension comprises a metering passage in communication with the feed passage.

Extension air feed hole blockage preventer for a gas turbine engine

An airfoil is provided. The airfoil may comprise a cross over, an impingement chamber in fluid communication with the cross over, and a first trip strip disposed on a first surface of the impingement chamber. A cooling system is also provided. The cooling system may comprise an impingement chamber, a first trip strip on a first surface of the impingement chamber, and a second trip strip on a second surface of the impingement chamber. An internally cooled engine part is further provided. The internally cooled part may comprise a cross over and an impingement chamber in fluid communication with the cross over. The cross over may be configured to direct air towards a first surface of the impingement chamber. A first trip strip may be disposed on the first surface of the impingement chamber.

Internal cooling cavity with trip strips

A pivoting turbine vane has an airfoil, an inner bearing race and an outer bearing race, with the inner and outer bearing races on a pivot axis of the pivoting turbine vane. There are cooling air passages through at least one of the inner and outer bearing races to provide cooling air from a remote facing face of at least one of the inner and outer bearing races to an airfoil facing face of at least one of the inner and outer bearing races. A turbine section is also disclosed.

Rotating turbine vane bearing cooling

In a featured embodiment, a lost core assembly includes a ceramic component having a tapered shape in a radial direction. A refractory metal component extends radially from the ceramic core component. A method of molding a gas turbine engine component is also disclosed.

Lost core molding cores for forming cooling passages

An airfoil includes pressure and suction side walls that extend in a chord-wise direction between leading and trailing edges. The pressure and suction side walls extend in a radial direction to provide an exterior airfoil surface. A main-body core cooling passage is arranged between the pressure and suction walls in a thickness direction and extends radially toward a platform. A skin core cooling passage is arranged in one of the pressure and suction side walls to form a hot side wall and a cold side wall. The hot side wall defines a portion of the exterior airfoil surface and the cold side wall defines a portion of the core passage. The skin core cooling passage is divided by a wall into two discrete first and second skin core cooling passages each supplied with cooling fluid from opposing sides.

Double wall turbine gas turbine engine vane with discrete opposing skin core cooling configuration

A gas turbine engine component comprises an airfoil with a suction side and pressure side extending from a leading edge to a trailing edge. There are a plurality of cooling holes adjacent the leading edge, with the cooling holes having a non-circular shape, with a longer dimension and a smaller dimension. The airfoil defines a radial direction from a radially outer end to a radially inner end, and radially outer of the cooling holes spaced toward the radially outer end, which have the longer dimension extending closer to parallel to the radial direction. Radially inner cooling holes closer to the radially inner end having the longer dimension extend to be closer to perpendicular relative to the radial direction compared to the radially outer cooling holes.

Fan cooling hole array

A component for a gas turbine engine includes a root with a neck that extends into a fir tree with at least one tooth, the root includes a feed passage in communication with a multiple of cooling passages that extend through the neck and fir tree.

Disk lug impingement for gas turbine engine airfoil

A combustor for a gas turbine engine includes a combustor shell having a shell opening therethrough, a combustor panel having a stud attached thereto, the stud extending through the shell opening. The stud includes a standoff to define an intermediate passage between the combustor shell and the combustor panel. A retainer is attached to the stud. A washer surrounds the stud and is positioned between the retainer and the combustor shell. The washer at least partially defines a cooling flow passage configured to direct a cooling airflow through the shell opening to impinge the cooling flow on at least one of the stud or the standoff.

Washer for combustor assembly

A combustor panel may include an attachment feature. Because conventional attachment features of conventional combustor panels may be insufficiently cooled, the present disclosure provides various combustor configurations for reducing hotspots in the vicinity of attachment features and/or for providing cooling airflow to and in the vicinity of attachment features.

Cooling configurations for combustor attachment features

Growing community of inventors

Naugatuck, CT, United States of America

Steven Bruce Gautschi