General Electric Company

Ge Avio S.r.l.

General Electric Company Polska Sp. Z O.o.

Ge Aviation Systems Limited

General Electric Deutschland Holding Gmbh

General Electric Company Polska Sp. Z 0.0.

Arthur William Sibbach

Brandon Wayne Miller

Nicholas Joseph Kray

Daniel Alan Niergarth

Jeffrey Douglas Rambo

Jeffrey Donald Clements

Bugra Han Ertas

Andrea Piazza

Scott Alan Schimmels

Tsuguji Nakano

Gary Willard Bryant, Jr

Elzbieta Kryj-Kos

Brian Lewis Devendorf

Frank Worthoff

Xiaohua Zhang

Miriam Manzoni

Ravindra Shankar Ganiger

Michael Anthony Benjamin

Ethan Patrick O'Connor

Erich Alois Krammer

James Ryan Reepmeyer

Michael John Franks

Timothy Richard DePuy

Paul Burchell Glaser

Bhaskar Nanda Mondal

Randy M Vondrell

Ian Francis Prentice

Wendy Wenling Lin

Stefan Joseph Cafaro

Scott Hunter

Aaron Michael Dziech

Christopher James Kroger

John Carl Glessner

Gregory Carl Gemeinhardt

Kevin Richard Graziano

Hao Huang

Mohamed Osama

Kirk Douglas Gallier

Joseph Alfred Iannotti

Ming Xie

Daniel Edward Mollmann

Constantinos Minas

Sibtosh Pal

Andrew Hudecki

John C Schilling

Egbert Geertsema

Sean Christopher Binion

Robert R Rachedi

Adam Tomasz Pazinski

Flavia Turi

Benjamin Thomas Van Oflen

Ryan T Roehm

Vincenzo Martina

Allan George Van De Wall

Timothy Richard Depuy

Adam Tomasz Paziński

Michael Macrorie

Vaibhav M Deshmukh

Sushilkumar Shevakari

Lawrence William Nurre

Douglas Downey Turner

Keith W Wilkinson

Jorge De Luis

Rakshit Tirumala

Brent Michael Push

Patrick Michael Marrinan

Rajesh Kumar

Richard Alan Wesling

Brian Gene Brzek

Eric Richard Westervelt

David Marion Ostdiek

Andrew J Martin

William Joseph Solomon

Donald Scott Yeager

Vishnu Vardhan Venkata Tatiparthi

Kevin Edward Hinderliter

Tomasz Jan Bulsiewicz

Steven B Morris

Tomasz Iglewski

Aaron Todd Sellinger

Shawn Patrick Riley

Douglas Lorrimer Armstrong

Inenhe Mohammed Khalid

Jacques Paul

Nagashiresha G

Keith Blodgett

Douglas Scott Jacobs

Koji Asari

Kedar S Vaidya

Ahmet Dindar

Christopher Michael Chapman

Avinash Natarajan

Prateek Mathur

Robert Eugene Dyson

Michał Tomasz Kuropatwa

James Laverty

Randy W Vondrell

Marcin Jacek Łobocki

Marcin Krzysztof Wachulec

A lubrication system for a turbine engine includes a reservoir that stores a lubricant, a primary lubricant supply circuit including a primary supply pump fluidly coupled to the reservoir, and an auxiliary lubricant supply circuit including an auxiliary supply pump fluidly coupled to the reservoir. A clutch is mechanically coupled to the auxiliary supply pump, and the clutch configured to engage the auxiliary supply pump and a shaft of the turbine engine when activated. The lubrication system further includes a pressure sensor that monitors a lubricant pressure within the primary lubricant supply circuit. When the lubricant pressure within the primary lubricant supply circuit falls below a predetermined lubricant threshold, the clutch is activated to engage the auxiliary supply pump and the shaft of the turbine engine.

Lubrication system and methods of lubricating a gearbox assembly

An airfoil assembly for a turbine engine, the airfoil assembly including a platform defining an inner surface and an outer surface, a variable pitch airfoil extending radially from the outer surface of the platform from a root to a tip to define a span length and a mounting structure connected to the platform.

Variable pitch airfoil

An aircraft engine assembly includes a gas turbine engine having an intake channel configured to receive an incoming flow of air and form an intake flow of air, the intake channel configured to turn the received incoming flow of air from an incoming flow direction to a first axial direction of the gas turbine engine, the incoming flow direction reverse of the first axial direction, and an electric machine coupled with the low pressure shaft and located at the aft end of the gas turbine engine proximate the intake channel, the electric machine in heat exchange communication with the intake flow of air such that the electric machine transfers heat to the incoming flow of air within the intake channel when the electric machine is operated.

Reverse flow gas turbine engine having electric machine

A turbomachine for a vehicle is provided. The turbomachine includes a first rotatable component; a first power source operatively coupled with the first rotatable component; a second power source selectively coupled with the first rotatable component; and a controller having one or more processors and one or more memory devices, the one or more memory devices storing instructions that when executed by the one or more processors cause the one or more processors to perform operations, in performing the operations, the one or more processors are configured to: receive an input indicating an engine shutdown of the turbomachine; and in response to the engine shutdown, cause the second power source to provide power to and rotate the first rotatable component.

System and method for mitigating bowed rotor in a gas turbine engine

A gas turbine engine includes: a turbomachine comprising a drive turbine and defining a working gas flowpath and an inlet to the working gas flowpath; a fan having a fan blade formed of a composite material, the fan blade defining a leading edge fan radius Rand a trailing edge fan radius R, and the fan defining a leading edge hub radius Rand a trailing edge hub radius R, the gas turbine engine defining a bypass ratio during operation of the gas turbine engine in a cruise operating mode; and a reduction gearbox mechanically coupling the drive turbine of the turbomachine to the fan; wherein the gas turbine engine defines a Fan Leading Edge to Trailing Edge Compression Factor (FLTCF) greater than or equal to 1.05 and less than or equal to 1.8.

Gas turbine engine having composite fan blades

A turbofan engine is provided. The turbofan engine includes a fan comprising a plurality of fan blades; a turbomachine operably coupled to the fan for driving the fan, the turbomachine comprising a compressor section, a combustion section, and a turbine section in serial flow order and together defining a core air flowpath; a nacelle surrounding and at least partially enclosing the fan; an inlet pre-swirl feature located upstream of the plurality of fan blades, the inlet pre-swirl feature attached to or integrated into the nacelle; and a means for reducing ice buildup or ice formation on the inlet pre-swirl feature, the means in communication with the inlet pre-swirl feature.

Ice reduction mechanism for turbofan engine

Outlet guide vane mounts with an airfoil structure are disclosed herein. The airfoil structure includes an airfoil platform, a composite material, a load spreader block, and a shim material, the composite material and a load spreader block positioned on the airfoil platform, the composite material having a first stiffness and the load spreader block having a second stiffness, and a shim material positioned between the composite material and the load spreader block, the shim material having a third stiffness different from at least one of the first stiffness or the second stiffness.

Outlet guide vane mount

A turbine engine with an engine core defining an engine centerline and comprising a rotor and a stator. The turbine engine including a set of composite airfoils circumferentially arranged about the engine centerline and defining at least a portion of the rotor. An airfoil in the set of composite airfoils including a composite portion extending chordwise between a composite leading edge and a trailing edge and a leading edge protector coupled to the composite portion.

Turbine engine with composite airfoils

A turbine engine includes a fan and a turbomachine defining an engine centerline. The turbomachine includes a compressor section, a combustion section, and a turbine section in serial flow order. The turbine engine also includes a set of composite airfoils circumferentially arranged about the engine centerline. An airfoil in the set of composite airfoils including a composite portion extending chordwise between a composite leading edge and a trailing edge and a leading edge protector coupled to the composite portion. The leading edge protector includes at least one projection extended past a base of the airfoil to a root of the airfoil.

Collars to retain multiple fan blades for use with an aircraft engine are disclosed herein. An example gas turbine engine includes a plurality of fan blades and a collar coupled to the plurality of fan blades. The collar includes a first mating piece and a second mating piece coupled to the first mating piece. The first and second mating pieces define an opening configured to receive a first root of a first fan blade of the plurality of fan blades, and the first and second mating pieces define a slot to retain a second root of a second fan blade of the plurality of fan blades.

Pitch-controlled blade retention collar

Methods and apparatus to control a surface of an aircraft engine are disclosed. An example system to control a surface in an aircraft engine comprises a first valve to vary a flow of cold fluid from a thermal transfer bus (TTB) to an active surface control (ASC) system based on an operating condition of the aircraft engine, the ASC system positioned adjacent to the surface, the first valve positioned upstream from the surface, and a second valve to vary a flow of hot fluid from the TTB to the ASC system based on the operating condition, the second valve positioned downstream from the surface.

Methods and apparatus to control a surface of an aircraft engine

A blended wing aircraft is provided, including a body having a fuselage and a pair of wings extending outward from the fuselage; and an aircraft engine defining an outlet and including a thrust reverser assembly, the thrust reverser assembly including a deployable structure extending less than 360 degrees around the outlet.

Blended wing aircraft

A heat exchanger positioned within an annular duct of a gas turbine engine is provided. The heat exchanger extends substantially continuously along the circumferential direction and defining a heat exchanger height equal to at least 10% of a duct height. An effective transmission loss (ETL) for the heat exchanger positioned within the annular duct is between 5 decibels and 1 decibels for an operating condition of the gas turbine engine. The heat exchanger includes a heat transfer section defining an acoustic length (L), and wherein an Operational Acoustic Reduction Ratio (OARR) is greater than or equal to 0.75 to achieve the ETL at the operating condition.

Gas turbine engine having a heat exchanger located in an annular duct

Growing community of inventors

Boxford, MA, United States of America

Arthur William Sibbach