General Electric Company

Krishnakumar Venkatesan

Michael Anthony Benjamin

Michael Terry Bucaro

Pradeep Naik

Steven Clayton Vise

Clayton Stuart Cooper

Perumallu Vukanti

Manampathy Gangadharan Giridharan

Ajoy Patra

R Narasimha Chiranthan

Narendra Digamber Joshi

Christian Lee Vandervort

Nayan Vinodbhai Patel

Parag Prakash Kulkarni

Fei Han

Duane Douglas Thomsen

Ertan Yilmaz

Shih-Yang Hsieh

Owen James Sullivan Rickey

Ahmed Mostafa ElKady

Ali Ergut

Arin Elspeth Lastufka Cross

Kevin Vandevoorde

Yuxin Zhang

Ramal Janith Samarasinghe

Michel Bucaro

A combined cycle propulsion system for a flight vehicle includes a compressor-fed combustion engine, and a multi-mode supersonic engine. The multi-mode supersonic engine includes an adjustable inlet section, a combustion section arranged downstream of the adjustable inlet section and including a first combustor portion having at least one rotating detonation combustor and a second combustor portion having a supersonic combustion type combustor, and an adjustable exhaust nozzle section arranged downstream of the combustion section. The at least one rotating detonation combustor functions as a pilot for the supersonic combustion type combustor.

Combined cycle propulsion system for hypersonic flight

An engine can utilize a combustor to combust fuel to drive the engine. A fuel nozzle assembly can supply fuel to the combustor for combustion or ignition of the fuel. The fuel nozzle assembly can include a swirler and a fuel nozzle to supply a mixture of fuel and air for combustion. The fuel nozzle can include both a primary and secondary fuel passage, and an additional air passage to provide for greater flame control, fuel provision, or local fuel and air mixing prior to combustion.

Gas turbine fuel nozzle having an inner air passage and plural outer fuel passages

Gas turbine nozzle having an inner air swirler passage and plural exterior fuel passages

A turbine engine comprising a compressor section, a combustion section having a combustor, and a turbine section in serial flow arrangement. The combustor having a combustion chamber, at least one fuel injector, at least one compressed air passage, and at least one igniter. The at least one igniter can be provided within a portion of the at least one fuel injector or the at least one compressed air passage.

Combustor with an igniter provided within at least one of a fuel injector or a compressed air passage

A fuel injection system for use in a combustor of a turbine engine includes a mixer assembly including a mixer housing and a fuel nozzle assembly positioned radially inward of the mixer housing. The fuel nozzle assembly includes a substantially annular fuel injection housing and a substantially annular main fuel injector coupled to the fuel injection housing. The main fuel injector includes a body, a fuel delivery passage defined in the body, a swirl chamber defined in the body downstream of the fuel delivery passage, and a plurality of circumferentially-spaced fuel metering slots defined in the body and coupled in flow communication with and between the fuel delivery passage and the swirl chamber.

Turbine engine fuel injection system and methods of assembling the same

A method and system of feeding fuel into a gasifier are provided. The feed injector system includes a first injector port assembly including a plurality of annular channels substantially concentric about a longitudinal axis that define corresponding fluid flow paths that direct a flow of fluid substantially axially therethrough from a respective source to a reaction zone and a second injector port assembly including a flow port surrounded by a plurality of auxiliary ports spaced about a circumference of the flow port, the plurality of auxiliary ports communicatively coupled to a toroidal passage configured to receive a flow of fluid and channel the flow of fluid to the auxiliary ports such that the flow of fluid is discharged from the auxiliary ports having an axial flow component, a radially inward flow component, and a circumferential flow component.

Method and system for mixing reactor feed

A system and method of reducing gas turbine nitric oxide emissions includes a first combustion stage configured to burn air vitiated with diluents to generate first combustion stage products. A second combustion stage is configured to burn the first combustion stage products in combination with enriched oxygen to generate second combustion stage products having a lower level of nitric oxide emissions than that achievable through combustion with vitiated air alone or through combustion staging alone.

System and method using low emissions gas turbine cycle with partial air separation

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Clifton Park, NY, United States of America

Krishnakumar Venkatesan