General Electric Company

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General Electric Company Polska Sp. Z O.o.

Michael Anthony Benjamin

Pradeep Naik

Steven Clayton Vise

Perumallu Vukanti

Clayton Stuart Cooper

Honggang Wang

Hiranya Kumar Nath

Manampathy Gangadharan Giridharan

Michael Terry Bucaro

Alfred Albert Mancini

Richard Louis Hart

Karthikeyan Sampath

R Narasimha Chiranthan

Ravindra Shankar Ganiger

Ajoy Patra

Hejie Li

Saket Singh

Joseph Zelina

Shai Birmaher

Nayan Vinodbhai Patel

Sibtosh Pal

Nicholas Ryan Overman

Brandon Wayne Miller

Duane Douglas Thomsen

Stefan Joseph Cafaro

Gregory Allen Boardman

Joshua Tyler Mook

Daniel J Kirtley

Eric Richard Westervelt

Marie Ann McMasters

Anil Raj Duggal

Joel Meier Haynes

Veeraraju Vanapalli

Radislav Alexandrovich Potyrailo

Arthur William Sibbach

Lawrence Bernard Kool

Kwanwoo Kim

Allen Michael Danis

Rimple Rakeshkumar Rangrej

Bangalore Aswatha Nagaraj

Randy M Vondrell

Craig Alan Gonyou

David Richard Barnhart

Narendra Digamber Joshi

Venkat Eswarlu Tangirala

Arthur Wesley Johnson

Kum Kang Huh

Jeffrey Eliot Chottiner

Hukam Chand Mongia

Erich Alois Krammer

Seung-Hyuck Hong

Sripathi Mohan

Gurunath Gandikota

Andrew Joseph Wickersham

Rajendra Mahadeorao Wankhade

Ramon Martinez

Robert Jon McQuiston

Varun Swaminathan Lakshmanan

Sean James Henderson

Krishnakumar Venkatesan

Paul Burchell Glaser

Eric John Stevens

Hari Ravi Chandra

Rameshkumar Muthuvel Chandrasekaran

Anquan Wang

Mark Richard Shaw

Shih-Yang Hsieh

Beverly Stephenson Duncan

Deepak Ghiya

Jayanth Sekar

Mayank Krisna Amble

Vishal Sanjay Kediya

Aaron J Glaser

Thomas George Holland

Brian Lewis Devendorf

David Justin Brady

Karthick Vilapakkam Gourishankar

Donald Lee Gardner

Ranganatha Narasimha Chiranthan

Sarah Marie Monahan

Mark David Durbin

Jeffrey Michael Martini

Richard Wade Stickles

Steven Joseph Lohmueller

William Thomas Ross

Dharmaraj Pachaiappan

George Chia-Chun Hsiao

Randall Charles Boehm

Hojjat Nasr

Owen James Sullivan Rickey

Bhavya Naidu Panduri

Pabitra Badhuk

Arundhati Sengupta

Aritra Chakraborty

Jacques Paul

Arvind Kumar Rao

Steven Marakovits

Mirko Gernone

Douglas Scott Jacobs

Koji Asari

A gas turbine engine includes a hydrogen fuel delivery assembly configured to deliver a hydrogen fuel flow, a compressor section configured to compress air flowing therethrough to provide a compressed air flow, and a combustor including a combustion chamber having a burner length and a burner dome height. The combustion chamber is configured to combust a mixture of the hydrogen fuel flow and the compressed air flow. The combustion chamber can be characterized by a combustor size rating between one inch and seven inches. In more detail, the combustion chamber can be characterized by the combustor size rating between one inch and seven inches at a core air flow parameter between two and one half kN and sixty kN, in which the combustor size rating is a function of the core air flow parameter.

Combustor size rating for a gas turbine engine using hydrogen fuel

A gas turbine engine fuel supply system can include a fuel delivery system, a thermal management system, a fuel manifold, and one or more sensors that identify one or more fuel parameters. A fuel control system is provided that adjusts parameters of the fuel based on data received from the sensors.

Methods of control for management of hot fuel

A diffuser for a combustor of a turbomachine engine. The diffuser includes a body having a first annular surface and a second annular surface that define a passage, the passage being configured to cause air to flow along the passage to a first component of the combustor. The first annular surface includes an opening that is configured to cause air to flow in a direction away from the passage to a second component of the combustor.

Diffuser with passlets

A turbine engine with a compressor section, a combustion section, and a turbine section in serial flow arrangement along an engine centerline. The combustion section including a primary combustor liner having an inner liner and an outer liner. A dome wall and a dome inlet are located in the dome wall. At least one opening is located in the outer liner downstream from the dome inlet. A primary combustion chamber and a set of secondary combustors are fluidly coupled to the primary combustion chamber at the at least one opening.

Combustion section with a primary combustor and a set of secondary combustors

A gas turbine engine includes a steam generating system, and a combustion section having a first combustor and a second combustor in fluid communication with a secondary combustion zone of the first combustor. The second combustor is operable to receive compressor bleed air and fuel, along with steam from the steam generating system. During a non-idle operating state and a power-augmentation operating state, when the steam system is operable, steam is provided to the second combustor and then to the secondary combustion zone of the first combustor. When the steam generating system is inoperable, the fuel and bleed air are provided to the second combustor and combusted, and combustion products are provided to the first combustor. When the steam generating system is partially operable, the steam, the bleed air, and the fuel are provided to the second combustor, and combustion products are provided to the first combustor.

Gas turbine engine having a combustion section with steam provided to second combustor plenum

A gas turbine engine comprising a compressor section, a combustion section, and a turbine section in a serial flow arrangement, with the combustion section comprising a combustor liner that at least partially defines a combustion chamber; and a gaseous fuel nozzle assembly, comprising: a steam supply to supply steam; a hydrogen fuel supply to supply gaseous hydrogen fuel; and a fuel nozzle body fluidly coupled with the steam supply and the hydrogen fuel supply.

Gas turbine and fuel nozzle assembly for mixing hydrogen fuel and steam to control flame speed

A turbine engine includes a compressor section for providing a compressed air flow, a fuel system for providing a primary combustor fuel supply, and a secondary combustor fuel supply. A primary combustor, located downstream of the compressor section, combusts the compressed air flow and the primary combustor fuel supply to generate primary combustion products. A turbine section, located downstream of the primary combustor, includes a turbine and a secondary combustor that further combusts the primary combustion products and the fuel from the secondary combustor fuel supply. A steam system provides a steam supply to the secondary combustor. The secondary combustor further combusts the primary combustion products and the secondary combustor fuel supply to generate secondary combustion products, and the turbine section is rotated by the secondary combustor steam supply and the inter-turbine combustion products.

Sequential combustion gas turbine engine including a steam system for injecting steam to each combustor

A turbine engine can utilize a combustor to combust fuel to drive the turbine, which drives 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. Increasing efficiency and carbon-containing emission needs benefit from the use of alternative fuels, which combust at higher temperatures than traditional fuels, requiring improved fuel introduction without the occurrence of flame holding or flashback.

Engine fuel nozzle and swirler

A fuel supply system for a combustion section of a turbine engine. The turbine engine including a heat source. The fuel supply system comprising a first fuel supply having a liquid fuel, a second fuel supply having a first gaseous fuel, and a heat exchanger fluidly coupled to the first fuel supply and the heat source. The heat exchanger producing a heat sufficient to change the liquid fuel to a second gaseous fuel that is the same type of fuel as the first gaseous fuel.

Fuel supply system for a combustor

A method for operating a propulsion system for an aircraft, the propulsion system including a gas turbine engine and a fuel cell assembly, the fuel cell assembly including a fuel cell stack having a fuel cell defining an outlet positioned to remove output products from the fuel cell during operation, the method including: executing a startup sequence for the fuel cell assembly, wherein executing the startup sequence for the fuel cell assembly includes initiating the startup sequence for the fuel cell assembly; executing a startup sequence for the gas turbine engine, wherein executing the startup sequence for the gas turbine engine comprises initiating the startup sequence for the gas turbine engine subsequent to initiating the startup sequence for the fuel cell assembly; and operating the fuel cell assembly to provide output products to a combustion section of the gas turbine engine.

Systems and method of operating a fuel cell assembly

A turbine engine has a compression section, a combustion section, and a turbine section in serial flow arrangement. The combustion section has a combustion liner and dome wall collectively forming at least a portion of a combustion chamber. The combustion section has a fuel nozzle assembly. The fuel nozzle assembly has a compressed air tube, and a body. The body defining a gaseous fuel channel exhausting into the combustion chamber.

Turbine engine having a combustion section with a fuel nozzle assembly

A gas turbine engine fuel supply system can include a fuel delivery system, a thermal management system, a fuel manifold, and one or more sensors that identify one or more fuel parameters. A fuel control system can be controlled to adjust one or more parameters of the fuel based on data received from the sensors.

Growing community of inventors

Cincinnati, OH, United States of America

Michael Anthony Benjamin