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

Ian Francis Prentice

Scott Alan Schimmels

Richard Alan Wesling

Tsuguji Nakano

Michael Anthony Benjamin

Donald Scott Yeager

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

David Marion Ostdiek

Brian Lewis Devendorf

Matthew Robert Cerny

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

Thomas Lee Becker, Jr

Eric Richard Westervelt

John Carl Glessner

Joseph Robert Dickman

Joel Francis Kirk

Michael Thomas Gansler

Ning Fang

Kevin Robert Feldmann

Xiaohua Zhang

Timothy Richard DePuy

Narendra Digamber Joshi

Honggang Wang

Andrew Breeze-Stringfellow

Kishore Ramakrishnan

Steven Douglas Johnson

Alfred Albert Mancini

Richard Louis Hart

Aaron Michael Dziech

Narayanan Payyoor

Craig Alan Gonyou

Craig William Higgins

Erich Alois Krammer

Weize Kang

Frank Worthoff

Arnab Sen

Raymond Floyd Martell

Douglas Duane Ward

Gary Willard Bryant, Jr

David Justin Brady

Jeffrey Anthony Hamel

William Joseph Bowden

Jeffrey Glover

Adon Delgado, Jr

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

Joseph George Rose

Pranav R Kamat

David Vickery Parker

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

A gas turbine engine includes a steam generating system and a turbo-engine. The turbo-engine includes a first spool having a fan and a low-pressure turbine drivingly connected together, a second spool having an intermediate-pressure compressor and an intermediate-pressure turbine drivingly connected together, and a third spool having a high-pressure compressor and a high-pressure turbine drivingly connected together. The turbo-engine also includes a first steam turbine arranged to receive a flow of steam from the steam generating system, and arranged to be connected to, and to drive the first spool, and a second steam turbine arranged to receive a flow of steam from the steam generating system, and arranged to be connected to, and to drive one of the second spool or the third spool.

Gas turbine engine having multiple steam turbines

A gearbox assembly includes a gearbox and a gutter for collecting a gearbox lubricant scavenge flow from the gearbox. The gutter is characterized by a lubricant extraction volume ratio between 0.01 and 0.3, and given by Vis a gutter volume of the gutter and Vis a gearbox volume. A gas turbine engine includes a combustion section and the gearbox assembly. A fuel delivery system includes a fuel supply line for delivering fuel to the combustion section. A lubrication system includes a lubricant supply line for delivering lubricant to the gearbox assembly. A thermal management system includes a fuel-lubricant heat exchanger for cooling the lubricant with the fuel. The thermal management system selectively directs the fuel through fuel bypass lines or the lubricant through lubricant bypass lines to bypass the fuel-lubricant heat exchanger based on a fuel temperature or a lubricant temperature.

Gearbox assembly with lubricant extraction volume ratio

A turbine engine for an aircraft. The turbine engine includes a combustor fluidly coupled to a fuel delivery assembly to receive fuel from the fuel delivery assembly. The fuel is injected into the combustor and combusted in the combustor to generate combustion gases. A condenser is located downstream of a turbine to receive the combustion gases and to condense water. The fuel heat exchanger is thermally coupled to the condenser to receive heat from the water condensed by the condenser. The fuel heat exchanger is located in the fuel delivery assembly to receive the fuel and to transfer the heat received from the water to the fuel. The boiler is located downstream of the fuel heat exchanger. The boiler receives the water and is fluidly connected to the combustor to receive the combustion gases and to boil the water to generate steam.

Turbine engine including a steam system

A turbofan engine is provided. The turbofan engine includes a fan having a plurality of fan blades; a turbomachine operably coupled to the fan for driving the fan, the turbomachine having 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, the nacelle defining a radius and a longitudinal axis; and an inlet pre-swirl vane located upstream of the plurality of fan blades and defining a chord, the inlet pre-swirl vane coupled to the nacelle, wherein the inlet pre-swirl vane is angled at a first angle with respect to the radius of the nacelle, and wherein the chord of the inlet pre-swirl vane is angled at a second angle with respect to the longitudinal axis of the nacelle.

Turbofan engine having angled inlet pre-swirl vanes

A gas turbine engine includes a gearbox assembly that includes a gearbox and a gutter for collecting a gearbox lubricant scavenge flow from the gearbox. The gutter is characterized by a lubricant extraction volume ratio between 0.01 and 0.3, inclusive of the endpoints. The lubricant extraction volume ratio defined by: is a gutter volume of the gutter and Vis a gearbox volume. The gas turbine engine includes a lubricant flow control system that includes a variable flow lubricant pump that generates a pump variable flow of lubricant to the gearbox assembly. The gearbox assembly has a variable consumption demand for delivery of lubricant. A lubricant flow controller is configured to generate a pump control command for the variable flow lubricant pump to produce the pump variable flow of lubricant based on the variable consumption demand.

Fluid-filled thrust link apparatus and associated method are disclosed. A thrust link for an aircraft engine includes a first wall having a forward portion and an aft portion at opposite ends of the thrust link, the forward portion coupled to the aircraft engine, the aft portion coupled to the aircraft engine, a pylon, or an aircraft associated with the aircraft engine, and a second wall within an interior area surrounded by the first wall, the second wall spaced apart from the first wall, a space between the first wall and the second wall defining a channel within the interior area, the channel including a fluid, the fluid pressurized based on a damping ratio to withstand a resonant vibration frequency generated by the aircraft engine.

Fluid-filled thrust link apparatus and associated method

A turbine engine having a longitudinal centerline axis. The turbine engine including a fan comprising a plurality of fan blades that rotate about the longitudinal centerline axis and a rotational component coupled to the fan. The turbine engine including a fluid circuit for supplying fuel or lubricant to the turbine engine and a hydraulic fan brake coupled to the fluid circuit to prevent rotation of the rotational component, thus preventing rotation of the fan. The hydraulic fan brake including a hydraulic cylinder fluidly coupled to the fluid circuit and a valve coupled to the hydraulic cylinder and having a first valve position that disengages the hydraulic fan brake to allow rotation of the rotational component and a second valve position that engages the hydraulic fan brake to prevent rotation of the rotational component.

Gas turbine engine having a hydraulic fan brake

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Liberty Township, OH, United States of America

Brandon Wayne Miller