Rtx Corporation

United Technologies Corporation

Raytheon Technologies Corporation

Andrew J Murphy

Lynn Macdonald

Sean Stevens

Cagan Gurer

John Mcwhirter

Naxin Tu

George D Yancopoulos

Vera Voronina

Karolina A Hosiawa

Nicholas J Papadopoulos

Marc J Muldoon

Jesper Gromada

Matthew E Bintz

Elena Burova

Richard Flavell

Michael E McCune

Michael G McCaffrey

William K Ackermann

Joel H Martin

David Frendewey

Ka-Man Venus Lai

Aris N Economides

Richard A Flavell

Jeffrey T Morton

Lisa I Brilliant

Joseph Parkos, Jr

James Olaf Hansen

Alexander O Mujica

Markus G Manz

Steven B Johnson

John C DiTomasso

Christopher J Hertel

Panayiotis Stevis

Matthew P Forcier

Yuan Dong

O Gavin Thurston

Steven D Roberts

Bindu Varghese

Elizabeth Eynon

Susan D Croll

Yingzi Xue

Anthony Rongvaux

Faith Harris

Paul J Hiester

Charles P Gendrich

Madhav Dhodapkar

Meghan Drummond Samuelson

Suzanne Hartford

Yonaton Ray

Jordan T Wall

Taryn Narrow

David R Buckler

Chozhavendan Rathinam

Daniel J Monahan

Jay Thomas Abraham

Jeffery Haines

Lucas D Sorensen

Michael E Mccune

Kihwa Kang

Ashley P Phillips

William J Beeson

Jamie Orengo

Li-Hsien Wang

Karolina A Meagher

Markus Manz

Tim Willinger

Till Strowig

George Yancopoulos

Brian Zambrowicz

Ana Kostic

Jr Joseph Parkos

Jay T Abraham

Jee H Kim

Chiara Borsotti

Christopher A Hertel

Ivan B Lobov

Katherine Diana Cygnar

Yajun Tang

An assembly for an aircraft propulsion system includes an open propulsor rotor having variable geometry propulsor blades and an open guide vane/stator assembly having a plurality of guide vanes (stator blades) in a fixed position relative to the rotor and assembly. The geometry (e.g., pitch, camber, etc.) of the guide vanes can be manually adjusted by ground personnel but are incapable of adjustment during a flight mission (a flight). The open propulsor rotor is configured to rotate about an axis, and the guide vane assembly includes the plurality of guide vanes arranged circumferentially about the axis and is disposed axially next to and downstream of the propulsor rotor.

Open rotor with fixed guide vane/stator

An assembly for a gas turbine engine includes at least one rotational assembly, an engine static structure, a compressor, one or more compressed air loads, and a particulate separator assembly. The at least one rotational assembly includes a shaft, a bladed compressor rotor, and a bladed turbine rotor. The engine static structure includes an engine case assembly. The engine case assembly surrounds the at least one rotational assembly. The compressor includes the bladed compressor rotor. The compressor is configured to form a compressed air flow. The one or more compressed air loads are disposed within the engine case assembly. The particulate separator assembly includes a plurality of particulate separators. The plurality of particulate separators are disposed outside of the engine case assembly. The plurality of particulate separators are configured to separate particulate from the compressed air flow and direct the compressed air flow to the one or more compressed air loads.

Particulate separator assembly for a gas turbine engine

A gas turbine engine includes a fan shaft rotatable about an axis, a fan connected to the fan shaft, and an outer housing surrounding the fan to define a bypass passage. A compressor section has both a low pressure compressor that is fixed to rotate with the fan shaft and a high pressure compressor. A turbine section has a low pressure turbine driving a low speed spool and a high pressure turbine driving the high pressure compressor. The low pressure turbine includes rotating blades and a static structure aft of the rotating blades. A fan drivetrain includes a geared architecture connecting the low speed spool to the fan shaft such that the fan and low pressure compressor rotate at a lower speed than the low pressure turbine. A thrust bearing is axially aft of the geared architecture and axially forward of the high pressure compressor. A tower shaft is rotatably driven by the low speed spool.

Geared turbofan with over-speed protection

Turbine engine systems and aircraft having such systems are described. The turbine engine systems include a combustor arranged along a core flow path of the turbine engine, a drive shaft having at least a compressor section and a turbine section coupled thereto, a cryogenic fuel tank configured to supply a fuel to the combustor, and an expansion turbine mechanically coupled to the drive shaft, the expansion turbine configured to receive fuel from the cryogenic fuel tank and expand said fuel, wherein expansion of said fuel by the expansion turbine drives rotation of the expansion turbine to provide power input to the drive shaft.

Gas turbine engines having cryogenic fuel systems

A rotor blade for a gas turbine engine includes a blade extending from a root and a contoured tip portion at a first end of the blade. The first end is opposite the root. The contoured tip portion includes a first sloped region and a second sloped region. The second sloped region is steeper than the first sloped region, relative to a platform.

Rotor blade for a gas turbine engine

An electroformed sheath for protecting an airfoil includes a sheath body and a mandrel insert is provided. The sheath body includes a leading edge. The sheath body includes a pressure side wall and an opposed suction side wall, which side walls meet at the leading edge and extend away from the leading edge to define a cavity between the side walls. The sheath body includes a head section between the leading edge and the cavity. The mandrel insert is positioned between the pressure side and suction side walls, and includes a generally wedge-shaped geometry. A method for protecting an airfoil includes: 1) securing a mandrel insert to a mandrel; 2) electroplating a sheath body onto the mandrel and the mandrel insert; 3) removing the mandrel from the sheath body so that a sheath cavity is defined within the sheath body; and 4) securing the airfoil within the sheath cavity.

Electroformed sheath

In one exemplary embodiment, a gas turbine engine system for cooling engine components includes an engine core, a core housing containing the engine core, an engine core driven fan forward of the core housing, a nacelle surrounding the fan and the core housing, and a bypass duct defined between an outer diameter of the core housing and an inner diameter of the nacelle. Also included is a thermal management system having a coolant circuit including at least one of a first heat exchanger disposed on the inner diameter of the nacelle and a second heat exchanger disposed on a leading edge of a BiFi spanning the bypass duct. The first heat exchanger is in thermal communication with the second heat exchanger.

Thermal management system for a gas turbine engine with an integral oil tank and heat exchanger in the nacelle

A compressor for use in a gas turbine engine comprises a compressor rotor including blades and a disc, with a bore defined radially inwardly of the disc. A high pressure air tap includes a lower temperature tapped path and a higher temperature tapped path and a valve for selectively delivering one of the lower temperature tapped path and the higher temperature tapped path into the bore of the disc. The valve is operable to selectively block flow of either of the lower pressure and higher pressure tapped paths to the bore of the disc, with the disc including holes to allow air from compressor chambers to communicate with the bore of the disc. A gas turbine engine and a method of operating a gas turbine engine are also disclosed.

High pressure compressor rotor thermal conditioning using discharge pressure air

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Old Saybrook, CT, United States of America

Andrew J Murphy