General Electric Company

General Electric Deutschland Holding Gmbh

Ge Avio S.r.l.

David Marion Ostdiek

Brandon Wayne Miller

Randy M Vondrell

Stefan Joseph Cafaro

William Joseph Solomon

Craig William Higgins

David Alexander Hiett

Alexander Kimberley Simpson

Darek Tomasz Zatorski

Syed J Khalid

Robert Jon McQuiston

Mohamed Osama

Jeffrey S Spruill

Arthur William Sibbach

Daniel Alan Niergarth

Jeffrey Douglas Rambo

Andrea Piazza

Alan Roy Stuart

Gert Johannes Van Der Merwe

Syed Arif Khalid

Daniel Lawrence Tweedt

Juraj Hrubec

Kevin Edward Hinderliter

Michael Scott McCambridge

Trevor Wayne Goerig

Drew Michael Capps

Stephen William Freund

Timothy M Kasberg

Logan Daniel Fritz

Matthew J Drodofsky

Matthew Vincent Gettemy

Matthew Truong

Alexander Kimberely Simpson

Jason Holden

Craig Williams Higgins

Jeffrey Spruill

Lance Blakeman

Michael Mcclain

Andrew Dreikosen

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 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 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 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, a secondary fan located downstream of the primary fan within the inlet duct, and a booster located downstream of the secondary fan and comprising a booster rotor blade, an inlet guide vane, and booster cowl, the booster cowl separating an upstream portion of the fan duct into an upper fan duct having an upper fan duct inlet and a lower fan duct having a lower fan duct inlet, the upper fan duct inlet and lower fan duct inlet collectively forming the fan duct inlet, the inlet guide vane located forward of the booster rotor blade.

A control system and schemes for controlling a three-stream gas turbine engine are disclosed. In one aspect, a three-stream engine is architecturally arranged so as to define a primary bypass flowpath, a secondary bypass flowpath, and a core flowpath that may each output propulsive thrust. The three-stream engine includes one or more effectors that can be controlled to adjust a thrust contribution provided by the secondary bypass flowpath to the net propulsive thrust as well as a thermal contribution provided by the secondary bypass flowpath to an associated thermal management system. Competing demands, limits, and priorities can be considered in controlling the effector. In some embodiments, a secondary effector can be ganged or controlled in conjunction with the effector to assist with adjustment of the contributions provided by the secondary bypass flowpath.

Three-stream gas turbine engine control

An engine with sound-based active controls is provided. The engine includes a sound sensor coupled to the engine system configured to capture sound from the engine system, one or more thrust effector devices, and an engine controller. The engine controller is configured to: determine a sound profile based on signals from the sound sensor, determine a modified control parameter for at least one of the one or more thrust effector devices based on the sound profile and a nominal schedule; and control the at least one of the one or more thrust effector devices according to the modified control parameter, deviating from the nominal schedule.

Turbine engine with sound-based active control

An engine for aeromechanical instability abatement includes a sensor configured to capture data from rotating blades of the engine system, a airflow effector device, and an engine controller. The engine controller is configured to control the airflow effector device according to a nominal schedule, detect, based on a signal from the sensor indicating a vibration amplitude of the rotating blades within a frequency band, an incipient instability condition, in response to the incipient instability condition being present, determine a modified control parameter for at least one of the airflow effector device, and control the airflow effector device according to the modified control parameter, deviating from the nominal schedule.

Aircraft engine aeromechanical instability detection

A method for reducing a clearance gap between a plurality of rotor blades and a shroud assembly of an engine includes determining, with a flight control system, that an airplane is in a first flight condition. The method also includes adjusting the clearance gap to a first clearance gap distance associated with the first flight condition. Further, the method includes receiving, with the flight control system, a demand for a second flight condition. During the second flight condition, the method includes adjusting at least two independently controllable parameters, the at least two independently controllable parameters comprising, at least, a first parameter for optimizing the clearance gap and a second parameter for satisfying a thrust demand of the engine, the first parameter having a first impact on the clearance gap, the second parameter having a second impact on the clearance gap, the first impact being greater than the second impact.

System and method for reducing a clearance gap in an engine

A method is provided for operating a hybrid-electric propulsion system having a first engine, a second engine, a first electric machine coupled to the first engine, and a second electric machine coupled to one of the first engine or the second engine. The method includes: receiving data indicative of a first engine operating parameter, a second engine operating parameter, or both; determining a first engine operating parameter margin, a second parameter operating margin, or both; determining a load share for the first engine, the second engine, or both, or between the first engine and the second engine based on the first engine operating parameter margin, the second engine operating parameter margin, or both; and transferring a first amount of power to or from the first electric machine and a second amount of power to or from the second electric machine in response to the determined load share.

Hybrid electric propulsion system load share

A method for energy conversion for a vehicle is provided. The method including extracting a flow of compressed fluid from a compressor section of a propulsion system; flowing the flow of compressed fluid to a turbine operably coupled to a driveshaft, in which the driveshaft is operably coupled to a load device; expanding the flow of compressed fluid through the turbine to generate an output torque at the driveshaft to operate the load device; and flowing the expanded flow of compressed fluid from the turbine to thermal communication with a thermal load.

Vehicle with energy conversion system

An electric propulsor is provided. The electric propulsor includes a core cowl and an outer cowl. A first air flowpath is defined radially outward of the outer cowl and a second air flowpath is defined between the core cowl and the outer cowl. The electric propulsor also includes one or more electric machines, a fan rotatably drivable by at least one electric machine, and a booster having a plurality of airfoils disposed at least in part in the second air flowpath, the booster being rotatably drivable by at least one electric machine for compressing air flowing along the second air flowpath. The electric propulsor further includes a heat exchanger disposed within the second air flowpath downstream of the booster, the heat exchanger being in thermal communication with at least one of the one or more electric machines and/or a gearbox mechanically coupled with the fan.

Electric propulsor with electric machine in thermal communication with a heat exchanger

A gas turbine is provided, the gas turbine engine including a turbomachine having an inlet splitter defining in part an inlet to a working gas flowpath and a fan duct splitter defining in part an inlet to a fan duct flowpath. The gas turbine engine also includes a primary fan driven by the turbomachine defining a primary fan tip radius R, a primary fan hub radius R, and a primary fan specific thrust rating TP; and a secondary fan downstream of the primary fan and driven by the turbomachine, the secondary fan defining a secondary fan tip radius R, a secondary fan hub radius R, and a secondary fan specific thrust rating TS; wherein the gas turbine engine defines an Effective Bypass Area, and wherein a ratio of Rto Requals

A propulsion system, the propulsion system comprising a rotating element, a stationary element, and an inlet between the rotating element and the stationary element, wherein the inlet passes radially inward of the stationary element; wherein the inlet passes radially inward of the stationary element; wherein the inlet leads to an inlet duct containing a ducted fan having an axis of rotation and a plurality of blades; and wherein the inlet duct divides into a first duct and a second duct, separate from the first duct. A method of operating a propulsion system, comprising the steps of: operating a first rotating fan assembly to produce a first stream of air; directing a portion of the first stream of air into a second ducted rotating fan assembly; operating the second ducted rotating fan assembly to produce a second stream of air; dividing the second stream of air into a core stream and a fan stream; and directing the core stream into a gas turbine engine core.

Propulsion system architecture

Growing community of inventors

Liberty Township, OH, United States of America

David Marion Ostdiek