United Technologies Corporation

Raytheon Technologies Corporation

Rtx Corporation

John S Tu

James B Hoke

Paul K Sanchez

Kalpendu J Parekh

William A Daniels

San Quach

Monica Pacheco-Tougas

Anthony Van

Robert Selinsky, Jr

Paul F Croteau

Matthew Andrew Hough

Russell B Hanson

Robert M Sonntag

John J Rup, Jr

Stephen C Harmon

Joshua C Rathgeb

Timothy S Snyder

Jonathan Jeffery Eastwood

Jaisukhlal V Chokshi

Dennis John Duhamel

Steven Wayne Burd

Brian T Hazel

Dennis M Moura

Robert Sze

John T Ols

Alexander W Williams

Kevin Joseph Low

Nurhak Erbas-Sen

James P Bangerter

Randal G McKinney

Jeffrey Jay Lienau

Jeffrey Leon

Frank J Trzcinski

Joseph James Sedor

William F Werkheiser

David Kwoka

Craig F Smith

Aaron S Butler

Jonathan M Jause

Sean D Bradshaw

William J Yeager

Jessica L Serra

Scott A Elliott

Neil B Ridgeway

Christopher R Brdar

Robert Newman

Andrew Cervoni

Gerard V Cadieux

David A Burns

Rajendra Prasad Uppara Allabanda

Kaitlin M Tomeo

Lisa A Jenkins

Lisa Tse

Shakira A Ramos

Venu Sambarapu

Prabaharan Manoharan

Jr John J Rup

Jr Robert Selinsky

Scott A Elliot

Dennis J Duhamel

Robert Selinsky Jr

John J Rup Jr

Jonathan Jeffrey Eastwood

Joseph J Sedor

Robert Selinsky

John J Rup

Flow diverters for installation in mid-turbine frame systems at a conduit outlet of gas turbine engines are described. The flow diverters include a diverter body having a connector portion defining a diverter inlet, a diverter extension at least partially defining a diverter outlet, and a curved portion arranged between the connector portion and the diverter extension, the curved portion configured to change a direction of flow from a first direction to a second direction that is about 90&#xb0; from the first direction as the flow passes from the diverter inlet to the diverter outlet.

Flow diverter for mid-turbine frame cooling air delivery

A diffuser case support structure for a gas turbine engine includes a fairing disposed circumferentially about a longitudinal axis. The fairing defines a plurality of passages circumferentially spaced apart and forming at least a portion of a fluid path between a compressor and a combustor of the gas turbine engine. A diffuser frame includes a plurality of struts. Each of the plurality of struts is disposed between a pair of adjacent passages of the plurality of passages. The diffuser frame is configured to couple an inner diffuser case to an outer diffuser case. A sliding joint is formed between the fairing and the diffuser frame, the sliding joint comprising a fastener extending generally axially through an aperture disposed in the fairing.

Diffuser case support structure

A combustor panel of a combustor may include a combustion facing surface, a cooling surface opposite the combustion facing surface, and heat transfer pins extending from the cooling surface. A grouping of the heat transfer pins may include a metallic coating.

Systems and methods for combustor panel

A diffuser case assembly for a gas turbine engine includes a fairing disposed circumferentially about a longitudinal axis. The fairing defines a plurality of passages circumferentially spaced apart and forming at least a portion of a fluid path between a compressor and a combustor of the gas turbine engine. A diffuser frame includes a plurality of struts. Each of the plurality of struts is disposed between a pair of adjacent passages of the plurality of passages. The diffuser frame is configured to couple an inner diffuser case to an outer diffuser case.

Diffuser case assembly

A diffuser case support structure for a gas turbine engine is disclosed. The diffuser case support structure includes a fairing disposed circumferentially about a longitudinal axis. The fairing forms at least a portion of a fluid path between a compressor and a combustor of the gas turbine engine. The fairing defines a plurality of apertures extending through the fairing. At least one spoke extends through at least one respective aperture of the plurality of apertures. The at least one spoke is configured to couple an inner diffuser case and an outer diffuser case of the gas turbine engine.

A wall assembly that may be for a combustor of a gas turbine engine includes a liner having a hot face that defines a combustion chamber, an opposite cold face, and a plurality of effusion holes. A shell of the assembly is spaced outward from the cold face and includes a plurality of impingement holes each having a centerline orientated substantially normal to the cold face. A plurality of cooling member arrays of the liner each include a first plurality of members that may be pins projecting outward from the cold face to conduct heat out of the liner. Each array is spaced between adjacent effusion holes and is symmetrically orientated about the respective centerline.

Cooled wall assembly for a combustor and method of design

A heat shield panel for a gas turbine engine includes a substrate layer having a first substrate surface and a second substrate surface opposite the first substrate surface. The first substrate surface and the second substrate surface define a substrate layer thickness therebetween. One or more thermally protective coating layers are applied to the first substrate surface of the substrate layer. The one or more coating layers have a constant coating layer thickness and the substrate layer thickness tapers along an axial length of the heat shield panel.

Combustor panel for gas turbine engine

A liner associated with an engine of an aircraft is described. The liner includes a panel and an array of projections configured to enhance a cooling of the panel and distributed on at least part of a first side of the panel that corresponds to a cold side of the panel.

Impingement film-cooled floatwall with backside feature

A turbine engine includes a compressor section, a combustor section in fluid communication with the compressor section, a high pressure turbine in fluid communication with the combustor, a low pressure turbine in fluid communication with the high pressure turbine, and a mid-turbine frame located axially between the high pressure turbine and the low pressure turbine. The mid-turbine frame includes an outer frame case, an inner frame case, and a plurality of hollow spokes that distribute loads from the inner frame case to the outer frame case. The spokes are hollow to allow cooling airflow to be supplied through the spokes to the inner frame case.

Internally cooled spoke

A wall panel assembly includes a first liner panel and a coating. The first liner panel has an inner first liner panel surface and a first liner panel outer surface each axially extending between a first liner panel first end and a first liner panel second end. The coating is disposed on at least one of the first liner panel inner surface and the first liner panel outer surface. The coating has an overall thickness that varies axially between the first liner panel first end and the first liner panel second end.

Wall panel assembly for a gas turbine engine

A liner panel for use in a combustor for a gas turbine engine includes a forward region including a multiple of effusion passages each directed at an orientation consistent with a local swirl direction of combustion gases. Another liner panel for use in a combustor for a gas turbine engine includes a forward region forward of a dilution passage and an aft region aft of the forward region. The forward region includes a multiple of effusion passages each directed generally circumferentially. The aft region includes a multiple of effusion passages each directed at an orientation generally axially. A method of cooling a wall assembly within a combustor of a gas turbine engine includes orienting a multiple of effusion passages within a forward region consistent with a local swirl direction of combustion gases.

Co-swirl orientation of combustor effusion passages for gas turbine engine combustor

A combustor of a gas turbine engine including a combustor shell having an interior surface defining a combustion chamber, a first panel mounted to the interior surface at a first position, the first panel having a first surface and a first rail extending from the first surface toward the combustor shell, the first rail configured at a first angle relative to the first surface, and a second panel mounted to the interior surface at a second position axially adjacent to the first panel, the second panel having a second surface and a second rail extending from the second surface toward the combustor shell, the second rail configured at a second angle relative to the second surface. The first and second rails are proximal to each other and define a circumferential gap there between and at least one of the first or second angles is an acute angle.

Combustor panels having angled rail

A combustor of a gas turbine engine including a combustor shell having an interior surface, a first panel mounted to the interior surface at a first position and a second panel mounted to the interior surface at a second position. The first panel has a first end, a first combustion chamber surface parallel with the interior surface, a first rail extending from the first combustion chamber surface toward the interior surface of the combustor shell, and a first extension extending axially from the first rail to the end of the first panel. The second panel has a second end, a second combustion chamber surface, and a second rail extending from the second combustion chamber surface toward the interior surface of the combustor shell. The first end and the second end are proximal to each other and define a circumferentially extending gap there between.

Combustor panels having recessed rail

A mid-turbine frame (MTF) for a gas turbine engine includes an inner manifold directing air to a turbine rotor of the gas turbine engine. The MTF includes an outer MTF case and an inner MTF case. The inner manifold of the MTF is located in the inner case of the MTF.

Internal manifold for turning mid-turbine frame flow distribution

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West Hartford, CT, United States of America

John S Tu