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

Patrick Michael Marrinan

Bugra Han Ertas

Ethan Patrick O'Connor

Christopher James Kroger

Donald Albert Bradley

Thomas Ory Moniz

Scott Alan Schimmels

Ian Francis Prentice

Tsuguji Nakano

Richard Alan Wesling

Michael Anthony Benjamin

Nicholas Joseph Kray

Andrea Piazza

David William Crall

Sridhar Adibhatla

Thomas Kupiszewski

Justin Paul Smith

Kurt David Murrow

Hendrik Pieter Jacobus De Bock

Paul Robert Gemin

Christian Xavier Stevenson

Nikolai N Pastouchenko

Gert Johannes Van Der Merwe

Hejie Li

John Carl Glessner

David Marion Ostdiek

Brian Lewis Devendorf

Matthew Robert Cerny

Egbert Geertsema

Kyle Robert Snow

Brian Francis Nestico

Nicholas Taylor Moore

Christopher Charles Glynn

William Dwight Gerstler

Bhaskar Nanda Mondal

Alan Roy Stuart

Mohammed El Hacin Sennoun

Joel Meier Haynes

Scott Hunter

Brian Gene Brzek

Stefan Joseph Cafaro

Duane Howard Anstead

Eric Richard Westervelt

Thomas Lee Becker, Jr

Timothy Richard DePuy

Joseph Robert Dickman

Joel Francis Kirk

Steven Douglas Johnson

Michael Thomas Gansler

Ning Fang

Kevin Robert Feldmann

Frank Worthoff

Xiaohua Zhang

Miriam Manzoni

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

Arnab Sen

Gary Willard Bryant, Jr

Raymond Floyd Martell

Douglas Duane Ward

David Justin Brady

Jeffrey Anthony Hamel

William Joseph Bowden

Jeffrey Glover

Ravindra Shankar Ganiger

Christopher Edward Wolfe

Trevor Howard Wood

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

Pranav R Kamat

Joseph George Rose

David Vickery Parker

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

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. Relationships between acoustic spacing and a high-speed shaft rating allow for a gas turbine engine that reduces noise emissions while maintaining high performance.

Aircraft engines and high temperature anti-ice systems for aircraft engines are disclosed herein. An example aircraft engine includes: a fan including a plurality of fan blades; a turbomachine operably coupled to the fan for driving the fan, the turbomachine including a compressor section, a combustion section, and a turbine section; a supply duct to accept bleed air from the compressor section; and a heat exchange system to capture waste heat from the turbine section and convey the waste heat to the bleed air, the bleed air with the waste heat to be conveyed to at least one of an environmental control system or a wing of an aircraft.

High temperature anti-ice systems for aircraft engines

A lubrication system for a turbine engine that includes one or more rotating components. The lubrication system includes one or more tanks that store lubricant, 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 stable operating conditions of the lubrication system. The auxiliary lubrication system includes an auxiliary feed line and an auxiliary supply line. The auxiliary lubrication system receives the lubricant from the one or more tanks through the auxiliary feed line. The auxiliary lubrication system supplies the lubricant to the one or more rotating components through the auxiliary supply line when there is a potential lubricant interruption in the lubrication system.

Lubrication system for a turbine engine

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

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. Acoustic spacing comprises a distance between the fan and the plurality of outlet guide vanes, and in combination with the use of airfoils with leading edge protectors for both various stages of blades and vanes, the engine effectively reduces noise emissions, enhances aerodynamic efficiency, and improves structural durability.

A gas turbine engine including: a fan assembly comprising a fan; and a turbomachine drivingly coupled to the fan and including a compressor section, a combustion section, and a turbine section arranged in serial flow order and defining in part a working gas flowpath, the gas turbine engine defining a bypass passage over the turbomachine; the turbomachine further including a heat exchanger and defining an annular cooling passage extending between an inlet and an outlet, the inlet in airflow communication with the working gas flowpath at a location upstream of the compressor section and the outlet in airflow communication with the bypass passage, the heat exchanger in thermal communication with an airflow through the cooling passage.

Gas turbine engine

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 method of detecting an airflow fault condition in a gas turbine engine, the method including: operating the gas turbine engine with a thermal transport bus having an intermediary heat exchange fluid flowing therethrough; determining a performance characteristic of the intermediary heat exchange fluid in the thermal transport bus is outside of a predetermined range, wherein the performance characteristic includes a temperature, a pressure, a flowrate, or a combination thereof; and indicating an airflow fault condition in response to determining the performance characteristic is outside of the predetermined range.

System and method of detecting an airflow fault condition

A propulsion system includes (a) a turbine engine including a compressor section, a combustor, and a turbine section, (b) a ramjet engine, and (c) an inlet section including a turbine engine inlet portion and a ramjet inlet portion, the inlet section providing an inlet airflow to the turbine engine and to the ramjet engine. The turbine engine inlet portion includes an expansion turbine arranged upstream of the compressor section.

Aircraft propulsion system

A turbofan engine defining an axial direction and a longitudinal centerline along the axial direction is provided. The turbofan engine includes: a fan section having a fan, the fan comprising a plurality of fan blades; a turbomachine drivingly coupled to the fan, the turbomachine comprising a compressor section with a low pressure compressor, a turbine section with a low pressure turbine, a reduction gearbox, and an outer casing, the low pressure turbine drivingly coupled to the low pressure compressor across the reduction gearbox; an outer nacelle surrounding the fan and at least a portion of the turbomachine; an outlet guide vane extending between the turbomachine and the outer nacelle at a location downstream of the plurality of fan blades, the outlet guide vane defining a base and a tip and being forward swept from the base to the tip.

Gas turbine engine with forward swept outlet guide vanes

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 gas turbine engine also includes a heat exchanger including an inner peripheral wall and an outer peripheral wall extending between an inlet and an outlet. The inner peripheral wall and the outer peripheral wall define a flow channel therebetween. The heat exchanger includes a plurality of fins disposed in the flow channel and dividing the flow channel into a plurality of flow passages.

Mounting assembly for a gearbox assembly

A gas turbine engine is provided, having: a bifurcation; and a turbomachine further including a cooled cooling air (CCA) system, the CCA system having: a cold side bleed assembly defining an inlet positioned to be in fluid communication with an airflow over the bifurcation; a CCA heat exchanger in thermal communication with the cold side bleed assembly downstream of the inlet of the cold side bleed assembly; and a hot side bleed assembly defining an inlet in fluid communication with a working gas flowpath through a compressor section, at a compressor discharge cavity, or both, the hot side bleed assembly in thermal communication with the CCA heat exchanger to cool an airflow through the hot side bleed assembly, the hot side bleed assembly further in thermal communication with a hot component of the turbomachine to cool the hot component of the turbomachine.

Gas turbine engine having cooling systems

A turbine engine having an inner cowl that circumscribes at least a portion of an engine core, where the inner cowl is radially spaced from the engine core. An outer cowl circumscribes at least a portion of the inner cowl where the outer cowl includes a radially outer surface spaced from a radially inner surface to define an outer cowl space. A fire box is located within a portion of the outer cowl space. An accessory device is located in the fire box.

Turbine engine having a fire box

A turbine engine includes a first gearbox, lubricated by oil. A first sump, located below the first gearbox, collects the oil. A first scavenge element pumps the oil from the first sump in an operational state and does not pump the oil in a failure state. A second gearbox is also lubricated by oil. A second sump, located below the second gearbox, collects the oil. A second scavenge element pumps the oil from the second sump in both the operational state and in the failure state. A conduit connects the first sump to the second sump to transfer overflow oil from the first sump to the second sump in the failure state, the overflow oil combining with the oil in the second sump, upon the oil in the first sump reaching a threshold level.

Turbine engine including a lubrication system

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Liberty Township, OH, United States of America

Brandon Wayne Miller