United Technologies Corporation

Raytheon Technologies Corporation

Rtx Corporation

Hamilton Sundstrand Corporation

Raymond C Surace

Tracy A Propheter-Hinckley

Robert A White, Iii

Eric A Hudson

Jonas S Banhos

Michael G McCaffrey

Dominic J Mongillo, Jr

Matthew A Devore

James T Roach

Russell Kim

San Quach

Bryan P Dube

Alan David Cetel

Timothy J Jennings

Jeffrey J DeGray

Emma J Place

Christopher J Peters

Andrew D Milliken

Jeremy B Fredette

Brian T Hazel

Dilip M Shah

Howard J Liles

Zhongfen Ding

Jonathan Ortiz

Tracy A Propheter

Mark F Zelesky

Kathryn S Read

William Abdel-Messeh

Shawn J Gregg

Jeremy Drake

Eleanor D Kaufman

Renee J Jurek

John D Teixeira

Ravon Venters

Brian D Merry

Brandon W Spangler

Thomas N Slavens

Jon E Sobanski

Francis P Marocchini

Gregory E Reinhardt

David Wasserman

Eric A Grover

Anthony P Cherolis

John R Farris

Edwin Otero

Jeffrey R Levine

Adam P Generale

Olivier H Sudre

Justin D Piggush

Venkatarama K Seetharaman

Tyler G Vincent

Brendan M Lenz

Mark Borja

Alex J Schneider

Amarnath Ramlogan

Gregory DiVincenzo

David A Krause

Alyson T Burdette

Gregory M Dolansky

Lucas Dvorozniak

Loi Cheng

John W Magowan

Eric P Letizia

Robert L Hazzard

Yongxiang D Xue

Mike R Lyon

Dilip N Shah

Tracy A Propheter-hinckley

Renee J Behlman

Daniel Carlson

Frank Thomas Cucinella

Dennis I Colon

Robert A White Iii

Dominic J Mongillo

Jeffrey J Degray

Jeffrey J Degraym

Tracy A Propheter- Hinckley

Jonathon Ortiz

Justin D Pigguish

Eleanor Kaufman

An assembly for use in releasing a preform from a tooling fixture comprises a solid body engageable with the fixture via a seal, a manifold formed in the solid body and in flow communication with the preform via a plurality of holes in the tooling fixture, and a fluid source in flow communication with the manifold and configured to supply fluid to the manifold. The fluid comprises at least one of a gas and a slurry.

Ceramic matrix composite part release from chemical vapor infiltration tooling

A gas turbine engine combustor panel assembly includes a shell defining an outer periphery of a combustion chamber. The shell has an outer skin with a first plurality of holes and a combustor panel, which includes a first end including a second plurality of holes on an inner face, which extend through the inner face to an outer face and are configured to allow fluid flow from the first plurality of holes to the combustion chamber. The combustor panel further includes a second end comprising a dovetail interface section extending axially from a central axis and from a panel wall of the combustor panel. The assembly further includes a plurality of dovetail tail connectors configured to extend along a central axis and interface with the second end of the combustor panel to connect the combustor panel to the shell.

Dovetail features on panel rails or impingement sheet to hold panels

A method of assembly includes locating a plurality of vanes about an inner circumference of a ring, the ring has a plurality of axially-extending projections circumferentially-spaced around the ring, each of the vanes has a radially outer platform, a flange that projects radially outwardly from the radially outer platform, and a notch that extends axially through the flange such that the notch opens at forward and aft flange faces of the flange, aligning the notch of each of the vanes with a corresponding one of the axially-extending projections on the ring, and inserting the projection into the notch to thereby prevent the vanes from rotating relative to the ring.

Airfoil anti-rotation ring and assembly

A method includes providing a continuous mandrel body disposed along an axis. The mandrel body defines, in order along the axis from a first end thereof, a pilot lead-in section, a first article section, a tapered transition section, and a second article section. Beginning at the pilot lead-in section, a continuous fabric sleeve is progressively formed on the mandrel body to the first article section, to the tapered transition section, and then to the second article section to form a fabric-covered mandrel. The portions of the continuous fabric sleeve that are on the first article section and the second article section form, respectively, a first article preform and a second article preform. The fabric-covered mandrel is then divided to separate the article preforms, followed by densifying the article preforms with a ceramic matrix material to thereby form CMC articles.

Method for forming multiple ceramic matrix composite articles

An airfoil for a gas turbine engine includes an airfoil section and a platform that has a non-gaspath side, a flange that extends from the non-gaspath side, and a gaspath side from which the airfoil section extends. The flange is comprised of a sandwich composite that includes first and second ceramic matrix composite (CMC) skins that each have at least one 2-D ceramic fiber ply, and a cellular core disposed between the first and second CMC skins.

Airfoil with sandwich composite flange

A tooling assembly suitable for use in infiltrating a fibrous preform comprises an outer tooling fixture and an inner tooling fixture. The outer tooling fixture comprises a first portion removably attached to a second portion, and a first pair of oppositely disposed windows defined by the first portion and the second portion. Each of the first portion and the second portion comprise a base with a recess. The inner tooling fixture comprises a body defined by at least one portion, the body configured to at least partially surround the fibrous preform, a plurality of holes extending through a thickness of the at least one portion, each hole of the plurality of holes having a length (L) and a diameter (D), and a first neck region engageable with a recess of the first portion or the second portion.

Floating tooling assembly for chemical vapor infiltration

A casting core assembly for a gas turbine engine component according to an implementation includes a skin core corresponding to a first cooling passage of an airfoil. The first cooling passage includes a first section and a tip flag section joined at a first bend. The skin core includes a first portion corresponding to the first section and a tip flag portion corresponding to the tip flag section. The skin core includes at least one arcuate slot corresponding to at least one turning vane of the airfoil. A method of forming an airfoil for a gas turbine engine is also disclosed.

Cooling arrangement for gas turbine engine component

A method of coating a seal slot includes applying a coating to a slot in a component and machining the coating to provide an open height configured to receive a seal. After the machining a surface roughness of the coating is less than about 100 ra. A scaling assembly is also disclosed.

Seal slot with coating

A combustor includes a combustion chamber and a liner that bounds at least a portion of the combustion chamber. The liner includes a first side facing the combustion chamber and a second side facing away from the combustion chamber. The liner is formed of a lay-up of ceramic matrix composite (CMC) plies that have a first CMC ply on the first side, a second CMC ply on the second side, and intermediate CMC plies between the first and second CMC plies. At least one of the intermediate CMC plies has a pattern of voids that define cooling channels in the combustor panel. The cooling channels are bound on lateral channel sides by the at least one of the intermediate CMC plies.

Combustor liner with pattern of voids in CMC fiber ply for cooling channels

A gas turbine engine combustor panel assembly includes a shell defining an outer periphery of a combustion chamber with an outer skin with a first plurality of holes, a combustor panel including a first end, a second plurality of holes extending through an inner face to an inner void, to an outer face with a third plurality of holes to allow fluid flow from the first plurality of holes to the combustion chamber, a plurality of bolt holes, a second end with an internal combustion chamber section extending axially from a central axis and from a panel wall of the combustor panel. The assembly includes a sheet metal seal plate abutting the plurality of bolt holes and the interface, and a plurality of bolts extending along a central axis and interfacing with the plurality of bolt holes to connect the combustor panel and shell.

T-head bolts or hollow flanged bushing with one thread or grooved end in a multi-liner arrangement

Disclosed is a method of reducing play in a vane arc segment. The vane arc segment includes an airfoil piece that defines first and second platforms and a hollow airfoil section that has an internal cavity and that extends between the first and second platforms. The first platform defines a gaspath side, a non-gaspath side, and a radial flange that projects from the non-gaspath side. Support hardware supports the airfoil piece via the radial flange, and a thermal insulation element is located adjacent the radial flange. The method includes performing a light scan of the radial flange to produce a digital three-dimensional model of the radial flange, and then machining the thermal insulation element in accordance with the digital three-dimensional model to provide a low-tolerance fit between the radial flange and the thermal insulation element that limits play between the airfoil piece and the thermal insulation element.

Vane arc segment with thermal insulation element

A ceramic matrix composite airfoil includes a high-pressure surface and a low-pressure surface connected at a leading edge and a trailing edge. The high-pressure surface and the low-pressure surface extend from a first end to a second end. A leading edge cooling passage includes an inlet portion, a midspan portion and an outlet portion. A cross-sectional flow area of the midspan portion is less than a cross-sectional flow area of either the inlet portion or the outlet portion.

Airfoil leading edge venturi cooling passage

An airfoil for a gas turbine engine includes an airfoil section and a platform that has a non-gaspath side and a gaspath side from which the airfoil section extends. The platform is comprised of a sandwich composite that has first and second ceramic matrix composite (CMC) skins that each include at least one 2-D ceramic fiber ply. The first and second CMC skins are disposed, respectively, on the gaspath side and the non-gaspath side, and there is a cellular core disposed radially between the first and second CMC skins.

Airfoil with sandwich composite

An airfoil for a gas turbine engine according to an example of the present disclosure includes an airfoil section extending in a radial direction from a root section to a tip portion. The airfoil section has an external wall and an internal wall. The airfoil section establishes an internal cooling arrangement including a first cooling passage having a first section and a tip flag section. The tip flag section extends in the chordwise direction along the tip portion from the first section. The first section includes a plurality of branched paths established by at least one turning vane.

A plurality of static stator vane are circumferentially spaced in a row axially intermediate the rows of turbine blades. Each stator vane has at least an outer platform with mount structure. Each of the stator vanes are formed of composite materials. The mount structure is provided with sacrificial material. At least a first of the stator vanes is circumferentially adjacent to a second of the stator vanes. An orientation of the first of the stator vanes is re-staggered relative to the second of the stator vanes. The sacrificial material of the first of the stator vanes is machined away such that a final trailing edge of the first of the stator vanes mount structure is now better axially aligned with the trailing edge of the second of the stator vanes mount structure trailing edge.

Removable layer to adjust mount structure of a turbine vane for re-stagger

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Middletown, CT, United States of America

Raymond C Surace