General Electric Company

Ge Avio S.r.l.

Edge Prototype LLC

General Electric Company Polska Sp. Z O.o.

Brandon Wayne Miller

Daniel Alan Niergarth

Arthur William Sibbach

Randy M Vondrell

Jeffrey Douglas Rambo

Jeffrey Donald Clements

Darek Tomasz Zatorski

Bugra Han Ertas

Patrick Michael Marrinan

Ethan Patrick O'Connor

Christopher James Kroger

Tsuguji Nakano

Scott Alan Schimmels

Donald Albert Bradley

Thomas Ory Moniz

Ian Francis Prentice

Michael Anthony Benjamin

Andrea Piazza

Richard Alan Wesling

David Marion Ostdiek

Justin Paul Smith

Egbert Geertsema

Nicholas Joseph Kray

David William Crall

Sridhar Adibhatla

Stefan Joseph Cafaro

Thomas Kupiszewski

Hejie Li

Brian Lewis Devendorf

Kurt David Murrow

Hendrik Pieter Jacobus De Bock

Paul Robert Gemin

Christian Xavier Stevenson

Joel Meier Haynes

Nikolai N Pastouchenko

Gert Johannes Van Der Merwe

Eric Richard Westervelt

Frank Worthoff

John Carl Glessner

Matthew Robert Cerny

Christopher Charles Glynn

William Dwight Gerstler

Bhaskar Nanda Mondal

Alan Roy Stuart

Mohammed El Hacin Sennoun

Scott Hunter

Brian Gene Brzek

Duane Howard Anstead

Thomas Lee Becker, Jr

Timothy Richard DePuy

Miriam Manzoni

Ravindra Shankar Ganiger

Steven Douglas Johnson

Michael Thomas Gansler

Ning Fang

Kevin Robert Feldmann

Xiaohua Zhang

Narendra Digamber Joshi

Honggang Wang

Andrew Breeze-Stringfellow

Kishore Ramakrishnan

Alfred Albert Mancini

Richard Louis Hart

Narayanan Payyoor

Aaron Michael Dziech

Craig Alan Gonyou

Craig William Higgins

Weize Kang

Erich Alois Krammer

Gary Willard Bryant, Jr

Arnab Sen

David Justin Brady

Raymond Floyd Martell

Douglas Duane Ward

William Joseph Bowden

Jeffrey Glover

Jeffrey Anthony Hamel

Michael Joseph Murray

Christopher Edward Wolfe

Trevor Howard Wood

David Raju Yamarthi

Kirk Douglas Gallier

Robert Proctor

Joseph Alfred Iannotti

Richard Ludwig Schmidt

Biao Fang

Nathan Evan McCurdy Gibson

Syed Arif Khalid

Monty Lee Shelton

Narendra Anand Hardikar

Mark Leonard Hopper

Jixian Yao

Scott Douglas Waun

Michele Gravina

Adam Joseph Wangler

Pranav R Kamat

Kevin Richard Graziano

Joseph George Rose

Mohan Kannaiah Raju

Santosh Kumar Pattnaik

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 aircraft fuel system includes a first fuel storage tank arranged to store a first fuel, a second fuel storage tank arranged to store a reduced oxygen fuel, and a fuel tank inerting system arranged to provide a reduced oxygen gas through a fuel contained within the second fuel storage tank to generate the reduced oxygen fuel within the second fuel storage tank. The fuel tank inerting system may include a fuel tank sparging system to inject the reduced oxygen gas directly into the fuel contained within the second fuel storage tank to generate the reduced oxygen fuel.

Aircraft fuel system

A gas turbine engine is provided. The gas turbine engine includes a turbomachine defining an engine inlet to an inlet duct, a fan duct inlet to a fan duct, and a core inlet to a core duct; a primary fan driven by the turbomachine; and a secondary fan located downstream of the primary fan within the inlet duct. The gas turbine engine defines a thrust to power airflow ratio between 3.5 and 100 and a core bypass ratio between 0.1 and 10, wherein the thrust to power airflow ratio is a ratio of an airflow through a bypass passage over the turbomachine plus an airflow through the fan duct to an airflow through the core duct, and wherein the core bypass ratio is a ratio of the airflow through the fan duct to the airflow through the core duct.

Gas turbine engine with third stream

A turbine engine for an aircraft includes a fuel delivery assembly for a hydrocarbon fuel to flow therethrough, a combustor combusting the fuel to generate combustion gases, and a core air exhaust nozzle exhausting the combustion gases from the turbine engine. The turbine engine also includes a contrail mitigation system having a heater and a fuel precipitate separator. The heater is selectively operable to heat the hydrocarbon fuel and to generate fuel precipitates in the hydrocarbon fuel, and the fuel precipitate separator separates the fuel precipitates generated by the heater from the fuel. A controller is coupled to the heater to operate the heater to heat the hydrocarbon fuel and to generate fuel precipitates in the hydrocarbon fuel in response to a contrail mitigation input.

Turbine engine for an aircraft including a contrail mitigation system

A hydrogen fuel system including a fuel delivery assembly, a purge gas source, and a vent. The fuel delivery assembly is configured to receive hydrogen fuel from a hydrogen fuel source and to provide the hydrogen fuel from the hydrogen fuel source to a power generator. The purge gas source is fluidly coupled to the fuel delivery assembly and configured to provide a purge gas to the fuel delivery assembly. The vent is fluidly coupled to the fuel delivery assembly and configured to vent hydrogen fuel from the fuel delivery assembly when the purge gas is provided to the fuel delivery assembly.

Purge system for a hydrogen fuel system

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

A pump system to pressurize a fluid within a closed loop transport bus is disclosed herein. The example of the pump system disclosed herein includes a pump, including an impeller, to increase kinetic energy of a fluid flowing through the impeller; an electric motor, including a rotor shaft connected to the impeller, to provide torque to the rotor shaft; a first bearing to support the rotor shaft at a first operational speed range, the first bearing coupled to an inner race; a second bearing to support the rotor shaft at a second operational speed range, the rotor shaft coupled to an outer race; and one or more sprag elements configured to: engage the inner race with the outer race at the first operational speed range; and disengage the inner race from the outer race at the second operational speed range.

Integrated bearing systems for dynamically supporting shafts in pump systems

A gas turbine engine fuel supply system can include a fuel delivery system, a thermal management system, a fuel manifold, and one or more sensors that identify one or more fuel parameters. A fuel control system is provided that adjusts parameters of the fuel based on data received from the sensors.

Methods of control for management of hot fuel

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 mounting assembly for a gearbox assembly of a gas turbine engine includes at least one mounting member configured to mount a gear of the gearbox assembly to a component of the gas turbine engine, the at least one mounting member characterized by a lateral impedance parameter, a bending impedance parameter, and a torsional impedance parameter. A gas turbine engine includes the mounting assembly. The at least one mounting member may be a flex mount, a fan frame, or a flex coupling. The gear includes a first gear that is a split sun gear including a forward sun gear and an aft sun gear separate from the forward sun gear. The forward sun gear and the aft sun gear are each rotationally coupled to a rotating shaft of the gas turbine engine.

Mounting assembly for a gearbox assembly

A gas turbine engine includes a fan section having a fan rotatable with a fan shaft, a turbomachinery section having a turbine and a turbomachine shaft rotatable with the turbine, a power gearbox mechanically coupled to the fan shaft and the turbomachine shaft such that the fan shaft is rotatable by the turbomachine shaft across the power gearbox, a grounded structure coupled to and supporting the power gearbox, and a torque monitoring system. The torque monitoring system includes a gearbox sensor. The gearbox sensor is coupled to the grounded structure and the torque monitoring system configured to determine a torque across the power gearbox using the gearbox sensor.

System and method for detecting and reducing torque in a turbine engine

A turbine engine including a turbo-engine, a gearbox assembly, a propulsor, and a lubrication system. The turbo-engine includes a compressor section, a combustor, a turbine section, and an input shaft. The gearbox assembly includes a first gear, a plurality of second gears, and a third gear. The propulsor has an output shaft drivingly coupled to the input shaft through the gear assembly. The lubrication system is characterized by a Gearbox Lubrication System Parameter (GLSP) between 0.2 and 140 when a mass flow rate of the lubricant is linear with a lubricant pump speed, where the GLSP is given by: The GLSP is between 0.2 and 70 when the mass flow rate of the lubricant is modulated, where the GLSP is given by:

Turbine engine having a lubrication system with a gearbox lubrication system parameter

A turbine engine including a fan having a plurality of fan blades, a turbo-engine positioned downstream of the fan and having a core inlet, an engine intake that extends to the core inlet, and a variable engine intake system. Air enters the turbine engine through the engine intake. The variable engine intake system includes a plurality of wind condition sensors for sensing wind conditions about the turbine engine, and a controller. The plurality of wind condition sensors includes a first wind condition sensor on a first side of the turbine engine and a second wind condition sensor on a second side of the turbine engine. The controller adjusts the engine intake based on the wind conditions about the turbine engine from the first wind condition sensor and the second wind condition sensor.

Turbine engine having a variable engine intake system

A turbine engine including a fan having a plurality of fan blades, a nacelle that extends circumferentially about the fan, an engine intake including an engine inlet, and a variable engine intake system. The nacelle includes a fan cowl and an inlet cowl that is movable with respect to the fan cowl. The engine inlet is defined from a leading edge of the nacelle to the plurality of fan blades. The variable engine intake system adjusts the inlet cowl axially between a fully retracted position and a fully extended position to adjust an inlet length of the engine inlet. The inlet cowl maintains contact with the fan cowl when the inlet cowl is extended.

A gas turbine engine comprises a fan, a core turbine engine coupled to the fan, a fan case housing the fan and the core turbine engine, a plurality of outlet guide vanes extending between the core turbine engine and the fan case, and an acoustic spacing. The fan comprises a plurality of fan blades that define a fan diameter and a BEAL. The fan case comprises an inlet and an inlet length between the inlet and the fan. The acoustic spacing comprises a distance between the fan and the plurality of outlet guide vanes, and in combination with the BEAL determines an acoustic spacing ratio of the gas turbine engine. The combination of acoustic spacing and corrected specific thrust provide enhanced propulsive efficiency.

Gas turbine engine with acoustic spacing of the fan blades and outlet guide vanes

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Liberty Township, OH, United States of America

Brandon Wayne Miller