United Technologies Corporation

Raytheon Technologies Corporation

United Technology Corporation

Georgia Tech Research Corporation

Rtx Corporation

Edward F Pietraszkiewicz

Dominic J Mongillo, Jr

Lisa P O'Neill

James Preston Downs

Mark F Zelesky

Alexander Staroselsky

Sergey Mironets

Frank J Cunha

Jeffrey R Levine

Om Parkash Sharma

Rafael A Perez

Tracy A Propheter-Hinckley

James M Koonankeil

Seth J Thomen

Todd R Coons

Brandon W Spangler

Steven Bruce Gautschi

Glenn Levasseur

Shawn J Gregg

Kirk David Hlavaty

Thomas N Slavens

Matthew A Devore

Atul Kohli

Thomas J Martin

Bret M Teller

Stanley J Funk

Christopher Corcoran

Richard M Salzillo

Christopher King

Craig R McGarrah

Young H Chon

Thomas R Davis

Michael J Bruskotter

Bryan P Dube

David A Litton

Kevin W Schlichting

Ricardo Trindade

Wieslaw A Chlus

Robert A Charbonneau

Scott W Gayman

Francisco J Cunha

Justin D Piggush

Jason Edward Albert

Anton G Banks

Mohamed Hassan

Andrew D Milliken

David James Hiskes

David J Candelori

Irwin D Singer

John C Calderbank

Barry M Beckwith

Richard Silverman

Kevin J Klinefelter

Sandra S Pinero

Kenneth A Lonczak

Heather Ann Terry

Laengheng Khoun

Ryan Shepard Levy

Norman F Roeloffs

Christina E Botnick

Markus W Fritch

Erik R Granstrand

Joseph A Peto

Kenneth P Botticello

David Michael Kontrovitz

Joshua E Kinnaman

Joseph F Englehart

Richard William Varsell

Takao Fukuda

Paul Thomas Rembish

Christopher B Jelks

Kenneth M Boucher

Jaime G Ghigliotty

Stephen D Doll

Domenico Valerio

William A Agli, Iii

Moon-Kyoo Brian Kang

Kar D Chin

Scott D Hartmann

Benjamin Adair

W Steven Johnson

Jesse Gannelli

Stephen J Yee

Wesley Harris

Steve Palmer

Norm Roeloffs

Richard Stockton

Wanda Widmer

David M Konopka

Luke A Hmiel

Dominic Mongillo

Francis M Tamiso

Pedro L Castaneda

Nestor A Alvarez

Jamie G Ghigliotty

Karl A Schachtner

Jason Daley

Travis Goeller

Vincent Barker

Khoun Laengheng

A combustor liner for a gas turbine engine includes at least one liner segment that has an external wall dimensioned to bound a combustion chamber. The external wall extends between leading and trailing edges in an axial direction and extends between opposed mate faces in a circumferential direction. A cooling circuit is defined by the external wall. A plurality of heat transfer features are distributed in the cooling circuit to define first and second prioritized flow regions on opposed sides of a first restricted flow region.

Preferential flow distribution for gas turbine engine component

A component for a gas turbine engine according to an example of the present disclosure includes, among other things, an airfoil section extending in a radial direction from a platform, the airfoil section extending in an axial direction between an airfoil leading edge and an airfoil trailing edge, and the airfoil section extending in the circumferential direction between pressure and suction sides. The platform extends in the axial direction between a platform leading edge and a platform trailing edge, and extends in the circumferential direction between a first mate face and a second mate face. The platform has a radially facing surface joined with the airfoil section and has a cold side surface opposed to the radially facing surface. A first thickness is defined between the radially facing surface and the cold side surface adjacent the first mate face, and a second thickness is defined between the radially facing surface and the cold side surface adjacent the second mate face. The first thickness is greater than the second thickness from at least the airfoil leading edge to the airfoil trailing edge with respect to the axial direction. A method of assembly is also disclosed.

Mate face arrangement for gas turbine engine components

A cooling circuit for a gas turbine engine includes a gas turbine engine component having at least one internal cooling cavity defined by an internal wall surface and a plurality of turbulent flow features extending outwardly from the internal wall surface. Each turbulent flow feature is spaced apart from an adjacent turbulent flow feature in a first direction. At least one trench extends through the turbulent flow features in the first direction, and a plurality of cooling holes are formed within the at least one trench. A gas turbine engine and a method of forming a cooling circuit for a gas turbine engine component are also disclosed.

Method of reducing manufacturing variation related to blocked cooling holes

A combustor liner for a gas turbine engine according to an example of the present disclosure includes, among other things, at least one liner segment that has an external wall dimensioned to bound a combustion chamber. The external wall extends between leading and trailing edges in an axial direction and extends between opposed mate faces in a circumferential direction. A cooling circuit is defined by the external wall. A plurality of heat transfer features are distributed in the cooling circuit to define a first restricted flow region that tapers from the leading edge to the trailing edge and to define at least one prioritized flow region that extends substantially from the leading edge to the trailing edge such that the at least one prioritized flow region is bounded by a perimeter of the first restricted flow region, and the at least one prioritized flow region has a lesser concentration of the plurality of heat transfer features than the first restricted flow region.

A blade assembly includes a blade and a blade platform secured to the blade. The blade extends radially from the blade platform. The blade platform includes at least one platform airflow passage located therein. A gusset extends from the blade to the blade platform. The gusset includes a gusset airflow passage fluidly connected to the platform airflow passage to convey an airflow to the platform airflow passage. a gas turbine engine includes a combustor and a plurality of gas turbine engine components located in fluid communication with the combustor. The gas turbine engine component includes an airfoil portion and a platform secured to the airfoil portion. The platform includes at least one platform airflow passage positioned therein. A gusset extends from the airfoil portion to the platform. The gusset includes a gusset airflow passage fluidly connected to the platform airflow passage to convey an airflow to the platform airflow passage.

Blade platform gusset with internal cooling

An airfoil includes a cooling air passage for receiving a cooling air flow. A chevron including a first rib and a second rib extends from a common tip is disposed within the cooling passage for generating a turbulent flow to improve heat transfer. The chevron includes an angle between the first rib and the second rib that is greater than 90 degrees.

Obtuse angle chevron trip strip

An airfoil for a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, an airfoil body and a cooling circuit disposed inside the airfoil body and including a leading edge cavity with a first portion extending from a radially inner wall to a radially outer wall of the airfoil body and a second portion that extends from a leading edge inner wall to a trailing edge inner wall of the airfoil body. The cooling circuit is configured to communicate cooling airflow through the first portion and the second portion prior to exiting the leading edge cavity into a second cavity of the cooling circuit.

Gas turbine engine airfoil cooling circuit

A method of designing a turbine blade includes the steps of forming at least two notches on a tip of a turbine blade, each of the at least two notches having a known dimension. The turbine blade has a pressure side and a suction side. The method further includes the step of operating a gas turbine engine including the turbine blade to expand a length of the turbine blade such that the tip of the turbine engages a casing. The method further includes the steps of viewing the tip of the turbine blade after the step of operating of the gas turbine engine, determining an appearance of the notches on the tip and determining a manufacturing length of the turbine blade based on the step of determining the appearance the notches.

Method of measuring turbine blade tip erosion

A vane structure includes a baffle movably mounted within an aperture, the baffle movable to control a cooling flow between a first cooling cavity and a second cooling cavity.

Modulated turbine vane cooling

A Blade Outer Air Seal (BOAS) includes a body manufactured of a metal alloy, the body includes a face opposite a forward interface and an aft interface, the face includes a cavity. A non-metallic insert within the cavity such that the insert is flush with the face.

Hybrid blade outer air seal for gas turbine engine

An airfoil of a gas turbine engine is provided. The airfoil includes an airfoil body having at least one internal flow passage, the body having a first surface and a second surface, the first surface defining a wall of the at least one internal flow passage and a bleed port fluidly connecting the at least one internal flow passage to the second surface. The bleed port includes a bleed orifice extending from the second surface toward the internal flow passage and a bleed port cavity extending from the first surface toward the second surface, the bleed port cavity and the bleed orifice fluidly connected. The bleed port cavity is defined by a bleed port cavity wall and a base wall surrounding the bleed orifice. The bleed port cavity wall extends from the first surface to the base wall.

Low loss airflow port

A gas turbine engine component includes first and second walls spaced apart from one another to provide a cooling passage. First and second trip strips are respectively provided on the first and second walls and arranged to face one another. The first and second trip strips are arranged in an interleaved fashion with respect to one another in a direction.

Gas turbine engine component cooling with interleaved facing trip strips

A component for a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a platform having a non-gas path surface and a gas path surface, a cover plate positioned relative to the non-gas path surface and a cooling passage that extends between the cover plate and the non-gas path surface. At least one cooling entrance is formed through the cover plate and configured to bias the flow of a cooling fluid toward a prioritized location of the non-gas path surface.

Gas turbine engine component providing prioritized cooling

A component for a gas turbine engine is provided. The component includes an internal cooling passage disposed within the component, and an s-shaped trip strip formed on a surface of the internal cooling passage.

S-shaped trip strips in internally cooled components

A platform is disclosed. The platform may include an airfoil section with a cooling passage and a platform. The platform may have various cooling features, such as a platform cooling apparatus. The platform cooling apparatus may have a cooling passage forming a channel disposed at least partially through the platform and the platform cooling apparatus may have an inflow channel in fluidic communication with the channel and the cooling passage so that cooling air may travel from the cooling cavity of the blade airfoil section and into the platform cooling apparatus. Moreover, the platform cooling apparatus may have a cooling cover apparatus at least partially fluidically sealing the platform cooling apparatus.

Serpentine platform cooling structures

Growing community of inventors

Southington, CT, United States of America

Edward F Pietraszkiewicz