Rolls-royce Corporation

Rolls-royce North American Technologies Inc.

Rolls-royce High Temperature Composites Inc.

Bruce Edward Varney

Ted Joseph Freeman

Daniel Kent Vetters

Aaron D Sippel

Jeffrey Allen Walston

Brett J Barker

David John Thomas

Jack Dwane Petty, Sr

Kenneth William Froemming

Eugene Clemens

Donald E Pinnick

Simon James Jones

Adam Lee Chamberlain

Sean E Landwehr

Andrew J Eifert

Steven Hillier

Peter Broadhead

Todd S Taylor

Brian J Shoemaker

Charles Alan Bulgrin

Paul Antony Davis

Jun Shi

Andrew Joseph Lazur

Joseph P Lamusga

Li Li

Ann Bolcavage

Eric Koenig

James Carl Loebig

Ngunjoh Lawrence Ndamka

Christopher Barrett

Kevin P Holley

Vickie Rae Smiley

Kyle Uhlenhake

Paul H Davis

David Thomas

A combustion assembly for a gas turbine engine may be provided. The combustion assembly may include a ceramic matrix composite combustor shell, which may include a chamber defined by a wall of the ceramic matrix composite combustor shell, and the ceramic matrix composite combustor shell may include a ceramic matrix composite chute integral with the ceramic matrix composite combustor shell. The ceramic matrix composite chute may extend towards a midline of the chamber. A method for fabricating a ceramic matrix composite chute may be provided. At least one chute may be woven in three dimensions into a ceramic preform. A layup tool may be inserted into the chute. The chute may be enlarged with the layup tool. The ceramic preform may be formed into a ceramic matrix composite body, which includes a combustor shell and the chute.

CMC combustor shell with integral chutes

A vane assembly for a gas turbine engine is disclosed in this paper. The vane assembly includes an inner platform, an outer platform, and a ceramic-containing airfoil. The ceramic-containing airfoil extends between the inner platform and the outer platform. A reinforcement spar extends between the inner platform and the outer platform through a hollow core of the ceramic-containing airfoil.

Vane assembly for a gas turbine engine

A method includes predicting a composition of calcium-magnesium-aluminum-silicate (CMAS) to be encountered by a high temperature mechanical system during use of the high temperature mechanical system. The method further includes selecting a composition of a CMAS-resistant barrier coating layer based at least in part on the predicted composition of CMAS. The CMAS-resistant barrier coating layer includes a base composition and at least one secondary oxide selected based on the predicted composition of CMAS. The at least one secondary oxide includes at least one of an oxide of a divalent element, an oxide of a trivalent element, or an oxide of a tetravalent element. The CMAS-resistant barrier coating layer comprises greater than 0 mol. % and less than about 7 mol. % of the at least one secondary oxide.

CMAS-resistant barrier coatings

A liner for a combustor includes a support member, an intermediate member, and a liner member. The intermediate member is positioned intermediate the support member and the liner member and has a plurality of protrusions and a plurality of recesses. The support member is coupled to the intermediate member at a tangent of each protrusion. Additionally, the liner member is comprised of a ceramic matrix composite material. The liner member is coupled to the intermediate member at a tangent of each recess.

Liner for a combustor of a gas turbine engine with metallic corrugated member

An airfoil with dual-wall cooling for a gas turbine engine comprises a spar having a pressure side wall and a suction side wall meeting at a leading edge and a trailing edge of the airfoil. An interior of the spar comprises a coolant cavity. The suction side wall includes an arrangement of rails on an outer surface thereof, where each rail extends along a non-chordal direction and terminates at or near the trailing edge. The airfoil also comprises a suction side coversheet overlying the suction side wall, where an inner surface of the suction side coversheet is in contact with the arrangement of rails so as to define a plurality of angled channels between the suction side wall and the suction side coversheet. The suction side wall also comprises inlet holes in fluid communication with the coolant cavity for feeding coolant to the angled channels, and the arrangement of rails is configured to direct the coolant through the angled channels and toward the trailing edge of the airfoil.

Airfoil with dual-wall cooling and angled cooling channels

A blade track for a gas turbine engine includes a plurality of blade track segments. The blade track segments are arranged circumferentially around a central axis to form the blade track.

Keystoned blade track

A turbine blade is secured with a turbine disk by a retention assembly. The turbine blade includes a blade stem and the disk includes at least one mount post for mounting of the turbine blade. The retention assembly is secured with the disk by forming a diffusion joint with the disk.

Blisk bonded CMC airfoil having attachment

An example gas turbine engine component includes a component configured to separate a cooling air plenum from a heated gas environment. The component includes a substrate defining a surface, and a unitary structure. The unitary structure includes a cooling region and a cover layer. The cover layer defines a hot wall surface configured to face the heated gas environment. The cooling region is disposed between the cover surface and the substrate and includes a plurality of support structures extending between the cover layer and the surface of the substrate. At least some of the support structures define a respective bond surface bonded to the substrate at the surface of the substrate. An example technique for fabricating the gas turbine engine component includes additively depositing the unitary structure on the surface of the substrate.

Additively deposited gas turbine engine cooling component

A component for a gas turbine engine that separates a cooling air plenum from a heated gas environment. The component defines a hot section surface adjacent to the heated gas environment having a plurality of cooling apertures fluidically connecting the cooling air plenum to the heated gas environment to allow a cooling air to flow from the cooling air plenum to the heated gas environment through the plurality of cooling apertures. The plurality of cooling apertures each have an aperture diameter of less than about 3 millimeters (mm) and an average surface roughness of less than about 1 micrometer (1 &#x3bc;m).

Air-film cooled component for a gas turbine engine

An assembly for a gas turbine engine includes ceramic material containing (i.e. ceramic matrix composite) segments and joints that couple the segments together.

Ceramic matrix composite joints

A gas turbine engine may comprise a blade track and a method of making the same. The blade track may be constructed of ceramic matrix composite components including main body members and joints.

Ceramic matrix full hoop blade track

Techniques for forming a dual-walled component for a gas turbine engine that include chemically etching at least one of a hot section part or a cold section part to form an etched part having plurality of support structures and bonding the etched part to a corresponding cold section part or a corresponding hot section part to form a dual-walled component, with the plurality of support structures defining at least one cooling channel between the at least one of the hot section part or the cold section part and the corresponding cold section part or the corresponding hot section part.

Dual-walled components for a gas turbine engine

A system for a hot section dual wall component in a gas turbine engine is used to avoid blockage by minimizing particulate deposits. The system includes impingement apertures formed in first wall of a cooling passageway of the dual wall component, and posts included on a second wall of the cooling passageway. The impingement apertures and the posts are respectively aligned opposite each other in the cooling passageway in operative cooperation so that working fluid exhausting into the cooling passageway from the impingement apertures in a first direction is directed to flow in a second direction along the cooling passageway to provide a laminar flow of the working fluid through the cooling passageway in order to minimize deposition of particles.

Hot section dual wall component anti-blockage system

An airfoil for use in a gas turbine engine is disclosed herein. The airfoil includes a pressure side wall, a suction side wall, and a cooling air distribution system. The suction side wall is arranged opposite the pressure side wall. The cooling air distribution system is formed internal to at least one of the pressure side wall and the suction side wall. The cooling air distribution system is configured to distribute cooling air through the airfoil to cool the airfoil.

Turbine engine components with cooling features

An annular assembly for a gas turbine engine and a method of making the annular assembly. The annular assembly comprises a supporting member comprising metal and including a support hole therethrough, and a liner tile made of ceramic matrix composite material. A plurality of the liner tile form an annular liner to shield hot combustion gases produced by the gas turbine engine. The liner tile is disposed adjacent the supporting member. A connecting member comprising ceramic material has a base element and a stem. The base element is embedded in the liner tile during construction of the liner tile to form an integrated structure therewith. A distal end of the stem passes through the support hole and is mated with a retaining member to retain the liner tile adjacent the supporting member.

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Greenwood, IN, United States of America

Bruce Edward Varney