United Technologies Corporation

Raytheon Technologies Corporation

San Quach

Steven Bruce Gautschi

Tracy A Propheter-Hinckley

Jeffrey T Morton

Matthew A Devore

Matthew D Parekh

Christopher King

Dominic J Mongillo, Jr

Thomas N Slavens

Adam P Generale

Alberto A Mateo

Lucas Dvorozniak

Jason Shenny

Bryan P Dube

Timothy J Jennings

Nicholas M LoRicco

John S Tu

Robert M Sonntag

Sasha M Moore

John J Rup, Jr

John McBrien

Clifford J Musto

Gregory Anselmi

Robert Selinsky, Jr

Raymond C Surace

Carey J Clum

Eric A Hudson

Donald William Peters

Christopher Whitfield

Lane Mikal Thornton

Tyler G Vincent

Brandon S Donnell

Anthony Van

Howard J Liles

Scott D Lewis

Atul Kohli

Christopher Corcoran

Winston Gregory Smiddy

Alex J Schneider

David J Candelori

Russell Deibel

Steven Taffet

Daniel C Nadeau

David Donald Chapdelaine

Fred NguyenLoc

Michael G McCaffrey

Brandon W Spangler

Brian T Hazel

James B Hoke

Robert A White, Iii

Matthew S Gleiner

Jeffrey S Beattie

Matthew Andrew Hough

Allan N Arisi

Monica Pacheco-Tougas

Ryan Alan Waite

Andreas Sadil

Stephen G Pixton

Jeffrey Leon

Anita L Tracy

Frank J Trzcinski

Joseph James Sedor

William F Werkheiser

Aaron S Butler

Jeremy Styborski

Jessica L Serra

John R Battye

Scott A Elliott

Caroline Karanian

Neil B Ridgeway

Yafet Girma

Andrew Cervoni

Jaime G Ghigliotty

Christina E Botnick

Gerard V Cadieux

Richard G Ullrich

Erik R Granstand

Cheng Gao

Kaitlin M Tomeo

Lisa A Jenkins

Lisa Tse

Kristy Rose Riley

Joanne V Rampersad

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

An airfoil vane assembly according to an exemplary embodiment of this disclosure, among other possible things includes a vane piece having a first vane platform, a second vane platform, and a hollow airfoil section joining the first vane platform and the second vane platform, a spar piece having a spar platform and a spar extending from the spar platform into the hollow airfoil section, and at least one seal arranged at a sealing surface of the spar platform and sealing between the spar platform and the first vane platform. The airfoil vane assembly also includes a thermal barrier coating disposed on the spar piece. The sealing surface is free from the thermal barrier coating. A gas turbine engine and a method of making a spar piece for an airfoil vane assembly are also disclosed.

Ceramic component with support structure

A gas turbine engine includes a ceramic wall for bounding an engine core gas path. The ceramic wall has a ceramic wall first side that faces the engine core gas path and a ceramic wall second side that faces away from the engine core gas path. There is a metallic wall adjacent the ceramic wall second side. The metallic wall has a metallic wall first side that faces the ceramic wall and a metallic wall second side that faces away from the ceramic wall. The metallic wall and the ceramic wall are spaced apart such that there is a channel there between. There is a seal on the ceramic wall second side, and the metallic wall has at least one cooling hole adjacent the seal for emitting cooling air to cool the seal.

Ceramic wall seal interface cooling

A gas turbine engine includes a blade outer air seal, a ceramic vane, and a cooling passage circuit that extends through a first internal passage in the blade outer air seal and a second internal passage in the ceramic vane.

Engine with cooling passage circuit for air prior to ceramic component

A vane arc segment includes an airfoil piece that defines first and second platforms and a hollow airfoil section that has an internal cavity and extends between the first and second platforms. The first platform defines a gaspath side, a non-gaspath side, and a flange that projects from the non-gaspath side. Support hardware supports the airfoil piece via the flange. There is a conformal thermal insulation blanket disposed on the flange.

Vane arc segment with conformal thermal insulation blanket

An airfoil includes a ceramic airfoil that defines a leading edge, a trailing edge, a pressure side, a suction side, a first radial end, and a second radial end. The ceramic airfoil section has an internal cavity and a rib that divides the internal cavity into a first radial passage and a second radial passage. The first radial passage is open at both the first radial end and the second radial end, and the second radial passage is open at least at the second radial end. A cooling passage circuit includes a first radial leg through the first radial passage, a second radial leg though the second radial passage, and a turn leg outside of the internal cavity at the second radial end. The turn leg connects the first radial leg and the second radial leg.

Ceramic airfoil with cooling air turn

In one exemplary embodiment, a flow path component assembly includes a support structure that is a unitary component having an axial portion and a radial portion and an outer portion arranged between the support structure and a flow path. The outer portion defines a gap between the outer portion and the support structure axial portion. A flange extends into the gap.

CMC component flow discourager flanges

An airfoil includes an vane airfoil, which has a ceramic matrix composite airfoil section including an outer wall that defines an internal cavity. The vane airfoil has an associated temperature profile which defines at least one high temperature area and at least one low temperature area. The vane airfoil includes a first zone which is defined in a first radial extent and corresponds to the high temperature area of the temperature profile and a second zone defined in a second radial extent and corresponds to the low temperature area of the temperature profile. An insert is situated in the internal cavity. The insert includes a first zone defined in a first radial extent that is aligned with the first zone of the vane airfoil. The first zone includes a first plurality of cooling holes which are configured to provide a first cooling density. A second zone is defined in a second radial extent that is aligned with the second zone of the vane airfoil. The second zone includes a second plurality of cooling holes that provide a second cooling density that is lower than the first cooling density such that the high temperature area and the low temperature area are balanced to lower a thermal gradient across the first and second zones of the vane airfoil. An vane airfoil and a method of assembling a ceramic matrix composite vane airfoil are also disclosed.

Baffle with impingement holes

A gas turbine engine includes one of a turbine section and a compressor section having multiple stages. At least one of the stages defines an outer diameter comprised of a plurality of circumferentially arranged blade outer air seals. Each blade outer air seal is spaced from each adjacent blade outer air seal in the plurality of circumferentially arranged blade outer air seals via a mateface gap. The mateface gap is oblique to a radius of the gas turbine engine, such that air entering the mateface gap is directed to an inner diameter surface of at least one of the blade outer air seals in the plurality of blade outer air seals.

Mateface for blade outer air seals in a gas turbine engine

A gas turbine engine component according to an example of the present disclosure includes a wall extending in a thickness direction between first and second wall surfaces. The first wall surface bounds an internal cavity. The wall includes a plurality of cooling passages. Each of the cooling passages extend in a first direction between an inlet port and an outlet port coupled to a respective diffuser, and the inlet port coupled to the internal cavity along the first wall surface. Sidewalls of adjacent diffusers are conjoined to establish a common diffuser region interconnecting the diffusers and a common outlet along the second wall surface. A method of cooling a gas turbine engine component is also disclosed.

Cooling arrangement including overlapping diffusers

A gas turbine engine component includes a wall that has an inner surface and an outer surface. An inlet is defined by the inner surface. At least one non-rectangular slot is defined by the outer surface and includes at least one protrusion extending into the slot. A slot passage fluidly connects the inlet to the at least one non-rectangular slot. The slot passage comprises an inlet portion that extends through the wall from the inlet to an intermediate portion. An outlet portion extends through the wall from the intermediate portion to the at least one non-rectangular slot.

Cooling passages for gas turbine engine component

An airfoil for a gas turbine engine is disclosed. In various embodiments, the airfoil includes a cooling passage; an outer wall separating a core flow path from the cooling passage; a diffuser in fluid communication with the cooling passage and opening into the core flow path, the diffuser being characterized by a linear ridge on a downstream end of the diffuser; and a thermal barrier coating covering the outer wall and the linear ridge.

Film cooling diffuser hole

An airfoil vane includes an airfoil section including an outer wall that defines an internal cavity; and a baffle situated in the internal cavity, the baffle including a baffle wall that defines a central cavity having a leading end and a trailing end corresponding to a leading end and a trailing end of the airfoil section, and a tail extending from the baffle wall, the tail including at least one feature configured to disturb an airflow surrounding the tail. A baffle for the airfoil vane assembly and a method of assembling a ceramic matrix composite airfoil vane are also disclosed.

Baffle with tail

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

An airfoil for a gas turbine engine according to an example of the present disclosure includes, among other things, an airfoil section that has an internal wall and an external wall. The external wall defines pressure and suction sides that extends in a chordwise direction between a leading edge and a trailing edge, a first impingement cavity and a second impingement cavity bounded by the external wall at a leading edge region that defines the leading edge. A first crossover passage within the internal wall is connected to the first impingement. The first crossover passage defines a first passage axis that intersects a surface of the first impingement cavity. A second crossover passage within the internal wall is connected to the second impingement cavity. The second crossover passage defines a second passage axis that intersects a surface of the second impingement cavity.

Impingement cooling arrangement for airfoils

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East Hartford, CT, United States of America

San Quach