General Electric Company

Symbol Technologies, LLC

Narendra Digamber Joshi

Venkat Eswarlu Tangirala

Nicholas William Rathay

Adam Rasheed

Joel Meier Haynes

Michael Solomon Idelchik

Nicholas Joseph Kray

Anthony John Dean

Sarah Marie Monahan

Manoj Ramprasad Shah

Kurt David Murrow

Douglas Carl Hofer

Christian Lee Vandervort

Richard Louis Hart

Brian Gene Brzek

Michael J Epstein

James Michael Caldwell

Edward E Ekstedt

Ronald Scott Bunker

Anil Raj Duggal

Honggang Wang

Lawrence Bernard Kool

Randy M Vondrell

Hendrik Pieter Jacobus De Bock

James William Bray

Timothy John Sommerer

Parag Prakash Kulkarni

Seung-Hyuck Hong

Paul V Heberling

James William Stegmaier

Brandon Wayne Miller

Adam Edgar Klingbeil

Thomas Michael Lavertu

Nitesh Jain

Krishan Lal Luthra

Andrei Tristan Evulet

Ke Liu

Konrad Roman Weeber

Joseph Zelina

Rui Zhou

Gregory Ronald Gillette

Zhi Zhou

Vladimir Zamansky

John Michael Koshoffer

William Paul Minnear

Kapil Kumar Singh

Matthew Patrick Boespflug

David Louis Burrus

Arnaldo Frydman

Jeffrey Anthony Hamel

Eric John Kress

Alan S Feitelberg

Michael Anthony Benjamin

Pradeep Naik

Daniel Alan Niergarth

William Dwight Gerstler

Miklos Stern

Mohamed Osama

Daniel Jason Erno

Russell Calvarese

Roy James Primus

Andrew Philip Shapiro

Timothy James Held

Manampathy Gangadharan Giridharan

Ian Francis Prentice

Seyed Gholamali Saddoughi

Brian Magann Rush

Mehmet Muhittin Dede

Vinayak Tilak

Ravi Rajamani

Patrick Michael Marrinan

Nikolai N Pastouchenko

Sandip Maity

Abhijeet Jayshingrao Yadav

Gary L Leonard

Jixian Yao

Chayan Mitra

Benjamin J Bekritsky

Joseph Douglas Monty

Gordon Raymond Smith

Francesco Papini

Keith Robert McManus

Andrew Joseph Wickersham

Omowoleola Chukwuemeka Akinyemi

Rachit Sharma

Michael Adam Bartlett

George Charles Goodman

Hongbo Cao

Gregory Alexander Natsui

Anoop Jassal

Gilbert Henry Badeer

Steven George Goebel

Shawn David Wehe

Gregg Anthony Deluga

Matthias Finkenrath

Robert Harold Weisgerber

Steven Keith Handelsman

Eric YuHang Fung

Hongwei Liao

Michael Colan Moscinski

Embodiments of a propulsion system are provided herein. In some embodiments, a propulsion system for an aircraft may include an electrical power supply; a motor coupled to the electrical power supply, wherein the electrical power supply provides power to the motor; and a fan disposed proximate a rear portion of an aircraft and rotatably coupled to the motor, wherein the fan is driven by the motor.

Propulsion system and methods of use thereof

Embodiments of a combustor for a gas turbine engine are provided herein. In some embodiments, a combustion chamber for a gas turbine engine comprising may include a combustor having an inner volume defined at least partially by a front wall, wherein the wall comprises a plurality of facets each having a through hole fluidly coupled to the inner volume, and wherein the plurality of facets are oriented such that an axis of each of the plurality of facets is offset from a central axis of the combustor by an angle.

Ultra compact combustor

A fuel mixer configured to provide a fuel-air mixture to a combustor of an engine. The fuel mixer may include a mixer body having a mixer outer wall, a center body, an annular passageway defined between the mixer outer wall and the center body, and a fuel tube assembly placed circumferentially about the mixer body. The fuel tube assembly may include at least one fuel channel for injecting a fuel flow into the annular passageway. The fuel tube assembly may be configured to cool a boundary layer flow present in the annular passageway. The fuel tube assembly may be configured to cool the mixer outer wall, the center body, or both the mixer outer wall and the center body. Heat from the mixer outer wall, the center body, or both the mixer outer wall and the center body, may pass to the fuel flow in the fuel tube assembly.

Fuel mixer

The present disclosure is directed to a rotating detonation combustor that includes a forward wall, a radially inner wall, and a radially outer wall. The forward wall is disposed at an inlet end of the rotating detonation combustor. The radially inner wall surrounds a longitudinal axis and extends downstream from the forward wall to an outlet end of the rotating detonation combustor. The radially outer wall extends downstream from the forward wall to the outlet end and surrounds the radially inner wall to define at least one annular plenum between the radially inner wall and the radially outer wall. At least one partition is proximate to the inlet end and defines at least two mixing zones. A plurality of oxidizer inlets and a plurality of fuel inlets are disposed at the inlet end in fluid communication with the at least two mixing zones.

Rotating detonation combustor with discrete detonation annuli

An engine assembly includes a combustor, a fuel cell stack integrated with the combustor, and a pre-burner system fluidly connected to the fuel cell stack. The fuel cell stack is configured to direct fuel and air exhaust from the fuel cell stack into the combustor. The pre-burner system is configured to control a temperature of an air flow directed into the fuel cell stack. The combustor is configured to combust the fuel and air exhaust from the fuel cell stack into one or more gaseous combustion products that drive a downstream turbine. The engine assembly can further include a catalytic partial oxidation convertor that is fluidly connected to the fuel cell stack. The catalytic partial oxidation convertor is configured to develop a hydrogen rich fuel stream to be directed into the fuel cell stack.

Integrated fuel cell and engine combustor assembly

An engine assembly includes a combustor, a fuel cell stack integrated with the combustor, the fuel cell stack configured (i) to direct fuel and air exhaust from the fuel cell stack into the combustor and (ii) to generate electrical energy, a catalytic partial oxidation convertor that is fluidly connected to the fuel cell stack, the catalytic partial oxidation convertor being configured to optimize a hydrogen content of a fuel stream to be directed into the fuel cell stack, and one or more subsystems electrically connected with the fuel cell stack, the one or more subsystems being configured to receive the electrical energy generated by the fuel cell stack. The combustor is configured to combust the fuel and air exhaust from the fuel cell stack into one or more gaseous combustion products that drive a downstream turbine.

A rotating detonation combustor includes a combustion chamber configured for a rotating detonation process to produce a flow of combustion gas and an air plenum configured to contain a volume of air. The rotating detonation combustor also includes a flow passage coupled in flow communication between the combustion chamber and the air plenum and configured to channel an airflow from the air plenum. The rotating detonation combustor also includes a fuel inlet coupled in flow communication with the flow passage and configured to channel a fuel flow into the flow passage. The flow passage includes a plurality of fuel mixing mechanisms configured to mix the airflow and the fuel flow within the combustion chamber.

Turbine engine assembly including a rotating detonation combustor

A fuel cell unit that includes a support structure having a plurality of flow channels and an active layer membrane coupled with the support structure, the active layer membrane comprising at least one electrode layer. Each flow channel of the plurality of flow channels is configured to direct one of air and fuel across at least one electrode layer of an active layer membrane to create electric current. Each flow channel of the plurality of flow channels includes at least one enhancement feature that is configured to disrupt a formation of a boundary layer near a surface of the active layer membrane where reactions occur. The plurality of flow channels can be positioned in a zig-zag configuration to allow for an increase in power density of the fuel cell unit.

High performance fuel cells

A propulsion system includes at least one rotating detonation actuator comprising: a flow path extending from an inlet end to an outlet end; an inner wall defining a radially inner boundary of the flow path; an outer wall defining a radially outer boundary of the flow path; and at least one aircraft wing. The rotating detonation actuator is disposed in the aircraft wing. At least one rotating detonation wave travels through the flow path from the inlet end to the outlet end.

Rotating detonation propulsion system

A gas turbine engine with a compressor section, a turbine section, and a combustion section located downstream from the compressor section and upstream from the turbine section, the combustion section including a dome inlet, a combustor outlet fluidly coupled to the turbine section, a liner and a dome assembly together at least partially defining a combustion chamber extending between the dome inlet and the combustor outlet, a fuel system fluidly coupled to the combustion section, the fuel system comprising a fuel supply, a primary fuel line fluidly coupling the fuel supply to the combustion section, and a reformer fluidly coupled to the fuel supply.

Turbine engine with fuel system including a catalytic reformer

A combustion section defines an axial direction, a radial direction, and a circumferential direction. The combustion section includes a casing that defines a diffusion chamber. A combustion liner is disposed within the diffusion chamber and defines a combustion chamber. The combustion liner is spaced apart from the casing such that a passageway is defined between the combustion liner and the casing. A fuel cell assembly is disposed in the passageway. The fuel cell assembly includes a fuel cell stack that has a plurality of fuel cells each extending between an inlet end and an outlet end. The inlet end receives a flow of air and fuel and the outlet end provides output products to the combustion chamber. The outlet end of the plurality of fuel cells extends through the combustion liner and partially defines the combustion chamber.

Gas turbine combustion section having an integrated fuel cell assembly

A system for removing heat from an aircraft component includes: a heat exchanger in proximity to an aircraft component and having an inlet, an outlet, and an internal surface coated with a catalyst; a source of hydrocarbon fuel in fluid communication with the inlet of the heat exchanger; a source of oxygen in fluid communication with the inlet of the heat exchanger; and a distribution system. A method of removing heat from an aircraft component includes providing a heat exchanger in proximity to an aircraft component, the heat exchanger being in fluid communication with a source of hydrocarbon fuel and a source of water and having an internal surface coated with a catalyst; introducing a hydrocarbon fuel into the heat exchanger; introducing oxygen into the heat exchanger; contacting the hydrocarbon fuel with the catalyst; and cracking the hydrocarbon fuel.

Systems and methods for removing heat from aircraft components

A flow control actuator includes a first side wall, a second side wall opposite and substantially parallel to the first side wall, an upstream wall mechanically coupled to upstream ends of the first and second side walls, and a downstream cap mechanically coupled to downstream ends of the first and second side walls. The first side wall, the second side wall, the upstream wall and the downstream cap collectively define an interior of the flow control actuator. An energy source is disposed in at least one of the first sidewall and the second sidewall. At least one fuel injector is disposed in the upstream wall, the first sidewall and/or the second sidewall for dispersing fuel into the flow control actuator. At least one air inlet is disposed in the upstream wall, the first sidewall and/or the second sidewall for introducing air into the flow control actuator. Fuel from fuel injector and air from the air inlet are ignited in the flow control actuator.

Combustion-powered flow control actuator with heated walls

A gas turbine engine includes; a compressor, a combustor, and a turbine in serial flow relationship; a heat exchanger, the heat exchanger having an inlet, an outlet, and an internal surface coated with a catalyst, the heat exchanger being located upstream of the compressor; a source of hydrocarbon fuel in fluid communication with the inlet of the heat exchanger; a source of oxygen in fluid communication with the inlet of the heat exchanger; and a distribution system for receiving reformed hydrocarbon fuel from the heat exchanger.

Gas turbine engine and method of operating

A gas turbine engine with a compressor section, a turbine section, and a combustion section located downstream from the compressor section and upstream from the turbine section, the combustion section including: a dome inlet, a combustor outlet fluidly coupled to the turbine section, a liner and a dome assembly together at least partially defining a combustion chamber extending between the dome inlet and the combustor outlet, a primary fuel injector fluidly coupled to the dome inlet, and a second fuel injector fluidly coupled to the combustion chamber.

Gas turbine engine combustor with primary and secondary fuel injectors

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Schenectady, NY, United States of America

Narendra Digamber Joshi