General Electric Company

Ge Avio S.r.l.

Edge Prototype LLC

General Electric Company Polska Sp. Z O.o.

General Electric Deutschland Holding Gmbh

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

Andrew Hudecki

Ethan Patrick O'Connor

Tsuguji Nakano

Christopher James Kroger

John C Schilling

Scott Alan Schimmels

David Marion Ostdiek

Donald Albert Bradley

Egbert Geertsema

Thomas Ory Moniz

Ian Francis Prentice

Ryan T Roehm

Michael Anthony Benjamin

Andrea Piazza

Richard Alan Wesling

Frank Worthoff

Donald Scott Yeager

Nicholas Joseph Kray

Justin Paul Smith

David William Crall

Sridhar Adibhatla

Stefan Joseph Cafaro

Thomas Kupiszewski

Hejie Li

Brian Lewis Devendorf

Kevin Edward Hinderliter

Robert Jon McQuiston

Sean Christopher Binion

Timothy Richard Depuy

Ravindra Shankar Ganiger

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

John Carl Glessner

Matthew Robert Cerny

Kyle Robert Snow

Nicholas Taylor Moore

Brian Francis Nestico

Brian K Kestner

Christopher Charles Glynn

William Dwight Gerstler

Bhaskar Nanda Mondal

Alan Roy Stuart

Mohammed El Hacin Sennoun

Scott Hunter

Brian Gene Brzek

Craig William Higgins

Duane Howard Anstead

Thomas Lee Becker, Jr

Timothy Richard DePuy

Miriam Manzoni

Joseph Robert Dickman

Joel Francis Kirk

Lawrence Chih-hui Cheung

Nicholas Rowe Dinsmore

Trevor Wayne Goerig

Alexander Kimberley Simpson

Peter Allen Andrews, Jr

Steven Douglas Johnson

Alfred Albert Mancini

Michael Thomas Gansler

Ning Fang

Xiaohua Zhang

Kevin Robert Feldmann

Flavia Turi

Victor Moreno Patan

Robert R Rachedi

Allan George Van De Wall

Andrew J Hank

Eric Barre

Robert Eugene Dyson

Geoffrey Whitener

Narendra Digamber Joshi

Honggang Wang

Andrew Breeze-Stringfellow

Kirk Douglas Gallier

Kishore Ramakrishnan

Richard Louis Hart

Narayanan Payyoor

Aaron Michael Dziech

Craig Alan Gonyou

Syed Arif Khalid

Erich Alois Krammer

Weize Kang

Gary Willard Bryant, Jr

Arnab Sen

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 inlet cowl to the plurality of fan blades. The engine inlet defines a droop axis at the leading edge of the inlet cowl. The inlet cowl is tilted at a droop angle that is defined as an angle of the droop axis with respect to an axial plane of the turbine engine. The variable engine intake system adjusts the droop angle of the inlet cowl during operation of the turbine engine.

Turbine engine having a variable engine intake system

Variable diameter thrust link apparatus are disclosed. A thrust link for an aircraft engine includes a forward end coupled to the aircraft engine, the forward end having a first diameter, and an aft end coupled to the aircraft engine or a pylon, the aft end having the first diameter, wherein the thrust link defines a thrust link span that extends from the forward end to the aft end, wherein a second diameter greater than the first diameter is defined between the forward end and the aft end, wherein the second diameter spans a center diameter span that is equivalent to or between 60% and 90% of the thrust link span.

Variable diameter thrust link apparatus

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 operable in a cold start condition to prevent wear on components of a gearbox assembly of the turbine engine. The turbine engine includes a gearbox assembly, a pump for directing lubricant to the gearbox assembly, a supply line heating path comprising a heat exchanger, a recirculation bypass path, a valve being positionable between a first position to direct the flow of the lubricant into the supply line heating path and a second position to direct the flow of the lubricant into the recirculation bypass path, and an electronic control unit configured to position the valve between the first position and the second position based on a temperature of the lubricant.

Systems and methods for heating a gearbox prior to fan rotation

A fuel oxygen reduction unit for an engine is provided. The fuel oxygen reduction unit includes an inlet fuel line; a stripping gas source; a contactor selectively in fluid communication with the stripping gas source, the inlet fuel line, or both to form a fuel/gas mixture; and a separator that receives the fuel/gas mixture, the separator configured to separate the fuel/gas mixture into an outlet stripping gas flow and an outlet fuel flow; wherein a flow of stripping gas passes through the fuel oxygen reduction unit a single time.

Fuel oxygen reduction unit for prescribed operating conditions

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 acoustic spacing parameter, in combination with composite fan blades that include one or both of a fan leading edge to trailing edge compression factor (FLTCF) and a fan leading edge to trailing edge opening ratio (FLTOR), provide improved performance and reduced acoustic noise.

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

A lubrication system for a turbine engine. The turbine engine includes a fan having a fan shaft and one or more rotating components. The lubrication system includes one or more tanks that stores lubricant therein, a primary lubrication system, and an auxiliary lubrication system. The primary lubrication system supplies the lubricant from the one or more tanks to the one or more rotating components during normal operation of the turbine engine. The auxiliary lubrication system includes an auxiliary pump that is coupled to the fan shaft. The auxiliary lubrication system supplies the lubricant from the one or more tanks to the one or more rotating components based on a pressure of the lubricant in the primary lubrication system. Rotation of the fan shaft causes the auxiliary pump to pump the lubricant to the one or more rotating components.

Lubrication system for a turbine engine

A disclosed example non-transitory machine readable storage medium includes instructions to cause programmable circuitry to at least determine, based on output from at least one sensor, (i) a first parameter corresponding to a first position in a casing of a gas turbine engine, the first position at or downstream of a volume at which flows from respective ones of bleed offtakes are combined, and (ii) a second parameter corresponding to a second position in a casing of a gas turbine engine, the second position upstream from the first position, determine a status of at least one of the bleed offtakes or the at least one sensor based on the first and second parameters, and provide or indicate the status in response to the status indicating improper operation of at least one of the bleed offtakes or the at least one sensor.

Methods and apparatus to determine engine status with plenum measurements

Viscous damper apparatus and associated methods to control a response to a resonant vibration frequency are disclosed. An apparatus to support an aircraft engine includes a thrust link including a forward end and an aft end, the forward end of the thrust link coupled to the aircraft engine, and a damper including a piston rod coupled to the aft end of the thrust link, the piston rod including a piston, and a chamber including a fluid, the piston to move within the chamber.

Viscous damper apparatus and associated methods to control a response to a resonant vibration frequency

A gas turbine engine is provided, comprising: a turbomachine having a compressor section, a combustion section, and a turbine section in serial flow order; and a fan section having a fan drivingly coupled to the turbomachine and an airflow surface rotatable with the fan and exposed to a fan airflow provided to and through the fan during operation of the gas turbine engine, the airflow surface defining a plurality of boundary layer openings configured to ingest a boundary layer of the fan airflow over the airflow surface during operation of the gas turbine engine.

Gas turbine engine

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.

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 blades feature a low aspect ratio, reducing blade count while maintaining thrust and efficiency. Efficiency is enhanced through a determined relationship between fan blade count, aspect ratio, and specific flow.

A turbine engine experiences 1P vibrations. The turbine engine includes a turbo-engine, a compressor for compressing air, and a combustor for combusting fuel and the compressed air to generate combustion gases. A turbine receives the combustion gases and drives a fan assembly. The fan assembly includes a fan disk and a fan blade with a fan blade centerline axis. The fan blade is subjected to 1P loading, generating the 1P vibrations. A fan blade root fixed to the fan blade and connected to the fan disk, is aligned with the fan blade centerline axis, and conducts the 1P vibrations. A damper positioned radially between the fan blade root and the fan disk damps the 1P vibrations.

Turbine engine including a fan assembly having a damper

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

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Liberty Township, OH, United States of America

Brandon Wayne Miller