General Electric Company

Mark Allan Hadley

Geoffrey David Myers

Jesse Ellis Barton

Jonathan Dwight Berry

Lucas John Stoia

Gregory Allen Boardman

Willy Steve Ziminsky

Kevin Weston McMahan

Gilbert Otto Kraemer

Ronald James Chila

Johnie Franklin McConnaughhay

Jun Cai

Jayaprakash Natarajan

Girard Albert Simons

Patrick Benedict Melton

Hendrik Pieter Jacobus De Bock

Joel Meier Haynes

Brian Gene Brzek

Sarah Lori Crothers

David Leach

Mark Stewart Schroder

Robert Joseph Rohrssen

Sam David Draper

John Joseph Lipinski

John Drake Vanselow

Thomas Edward Wickert

David Kenton Felling

Sergey Anatolievich Meshkov

Joel Meador Hall

Chetan Babu Velkur

Sergey Adolfovich Oskin

Rajani Kumar Akula

Jesse Blair Butler

Sergey Konstantinovich Yerokhin

Glenn David Nelson

Jayaprakesh Natarajan

A gas turbine includes a compressor and at least one combustor downstream from the compressor. The combustor includes a burner having an inner shroud extending axially along at least a portion of the burner, an outer shroud radially separated from the inner shroud and extending axially along at least a portion of the burner, and a plurality of stator vanes extending radially between the inner shroud and the outer shroud. The stator vanes have an inner end proximate the inner shroud and an outer end proximate the outer shroud. The burner further includes a vortex tip at one of either the inner end or the outer end of the stator vanes. The vortex tip provides a gap between the inner end and the inner shroud or the outer end and the outer shroud, and the vortex tip includes a plurality of fuel ports. The gas turbine further includes a turbine downstream from the combustor.

Gas turbine burner

A combustor and a method for reducing a temperature gradient of a combustor component are provided. The combustor includes a coating applied to at least a portion thereof with the coating serving to alter the emissivity of the at least a portion to which it is applied. The method includes applying a coating on at least one of a combustor liner and a flow sleeve, wherein the coating alters the emissivity exhibited where applied.

Gas turbine combustor including a coating having reflective characteristics for radiation heat and method for improved combustor temperature uniformity

A fuel injector is provided and includes a member defining a flowpath through which a first fluid flows, the flowpath having a cross-section with transverse elongate and short axes, a head defining a plenum storing a supply of a second fluid and a system fluidly coupled to the flowpath and the plenum to inject the second fluid from the plenum and into the flowpath at first and second locations along the elongate axis. The injected second fluid is formed into jets at the first and second locations, the first fluid entrains the jets such that the injected second fluid flows through the flowpath and mixes with the first fluid, and the short axis has a sufficient dimension such that the jets remain spaced from a sidewall of the member.

Fuel injector

A fuel nozzle is provided. The fuel nozzle includes a nozzle body, a plurality of swirler vanes, and at least one outlet. The nozzle body includes a back plate, a front plate, and a mixing zone defined therebetween. The back plate includes at least one inlet defined therein and the front plate includes at least one discharge defined therein. The plurality of swirler vanes are positioned between the back plate and the front plate and spaced circumferentially about the mixing zone. Each of the plurality of swirler vanes direct air obliquely into the mixing zone. The at least one outlet is defined within at least one of the nozzle body and the plurality of swirler vanes, the at least one outlet configured to inject fuel into said mixing zone.

Fuel nozzle and method of assembling the same

A combustor having a combustion chamber is provided with an external flow sleeve and a combustor liner surrounding the combustion chamber. A plurality of flow channels are provided on the combustor liner and a plurality of nozzles are disposed at predetermined locations on the flow channels. The locations of the nozzles are selected to provide different mixing times for fuel injected through the nozzles.

Systems and methods for suppressing combustion driven pressure fluctuations with a premix combustor having multiple premix times

Disclosed is a combustor including a baffle plate having at least one through baffle hole and at least one fuel nozzle extending through the at least one baffle hole. At least one shroud is secured to the baffle plate and includes at least one piston ring disposed at the shroud. The at least one piston ring is configured to meter a flow of diluent between the at least one shroud and the at least one fuel nozzle. Further disclosed is a method for providing diluent to a combustor including providing a piston ring gap defined by at least one piston ring disposed at a baffle plate and a fuel nozzle extending through a through hole in the baffle plate. The diluent is flowed through the piston ring gap toward at least one airflow hole in the fuel nozzle.

Method and apparatus for introducing diluent flow into a combustor

A system for augmenting gas turbine power output includes a compressed air supply, and a compressed air storage plenum in fluid communication with the compressed air supply. The compressed air storage plenum is configured to store a compressed air from the compressed air supply for later use. The system further includes an inlet plenum sealingly coupled to an inlet of the gas turbine. The inlet plenum is in fluid communication with the compressed air storage plenum so as to route the compressed air from the compressed air storage plenum into the inlet of the compressor during augmented operation of the gas turbine.

System and method for augmenting gas turbine power output

A fuel injector having an injector body, a mixing circuit, and at least one injector is provided. The injector body has a plurality of manifolds, an inlet, and an outlet. The manifolds are configured for receiving fuel, and the inlet is configured for receiving air. The mixing circuit is positioned within the injector body. The mixing circuit is configured for receiving fuel from at least one of the manifolds, and air from the inlet to create an air-fuel mixture that exits the outlet. The least one fuel injector is positioned radially outwardly from the mixing circuit. The at least one injector receives fuel from at least one of the plurality of manifolds and injects fuel to the outlet.

Fuel injector with mixing circuit

A fuel injector is provided and includes a first tube, having first and second opposing ends, which is supplied with fuel, and one or more second tubes disposed within the first tube, each of the one or more second tubes being supplied with air and having sidewalls defining injection holes through which the fuel enters the one or more second tubes to mix with the air, and an outlet end of the sidewalls corresponding to the second end of the first tube.

Methods and systems are provided for inducing combustion dynamics within turbine engines to remove combustion deposits within the turbine engine during operation of the turbine engine.

Methods and systems for inducing combustion dynamics

A fuel nozzle tip for use with a combustor and a method of assembling the fuel nozzle tip are provided. The fuel nozzle tip includes a fuel tube and an air collar coupled to the fuel tube. The fuel tube includes a first plurality of circumferentially-spaced fuel openings and a second plurality of circumferentially-spaced fuel openings. The fuel tube is configured to channel fuel into a mixing zone defined within the combustor. The air collar includes a plurality of circumferentially-spaced air openings configured to discharge air into the mixing zone.

Airblown syngas fuel nozzle with diluent openings

Disclosed is a combustor including a baffle plate having at least one through baffle hole and at least one fuel nozzle extending through the at least one baffle hole. A plurality of injection holes extend through the at least one fuel nozzle and are configured to meter a flow of diluent into the combustor. Further disclosed is a method for providing diluent to a combustor including providing a plurality of openings located at at least one fuel nozzle extending through a through hole in a baffle plate. The diluent is flowed through the plurality of openings toward at least one airflow opening in the at least one fuel nozzle.

Method and apparatus of introducing diluent flow into a combustor

A method of regulating the Wobbe number of a multi-composition gas fuel supplied to one or more combustors of a gas turbine includes: (a) providing a control system for regulating fuel and air flow to the one or more combustors; and (b) reforming a fraction of the gas fuel upstream of the one or more combustors to form hydrogen and carbon monoxide to be supplied to the one or more combustors with a remaining fraction of the fuel; wherein the fraction of fuel reformed is adjusted to maintain the Wobbe number of the fuel supplied to the one or more combustors within a predetermined range.

Wobbe control and enhanced operability through in-line fuel reforming

A fuel injector is provided, including a fuel injector body, a plurality of fuel vanes, and a plurality of fuel pegs. The injector body includes a manifold and an inlet. The manifold is configured for receiving fuel, and the inlet is configured for receiving air. The fuel vanes are located within the injector body and are positioned in a direction that is generally parallel with a longitudinal axis of the injector body to orient the air flowing from the inlet. The plurality of fuel pegs are fluidly connected to the manifold and are arranged within the plurality of fuel vanes. The plurality of fuel pegs are each spaced at a distance that is about equal between each of the plurality of fuel pegs.

Injector having multiple fuel pegs

A gas turbine combustor includes a combustion chamber defined by a combustor liner, the combustor liner having an upstream end cover supporting one or more nozzles arranged to supply fuel to the combustion chamber where the fuel mixes with air supplied from a compressor. A transition duct is connected between an aft end of the combustion chamber liner and a first stage turbine casing, the transition duct supplying gaseous products of combustion to the first stage turbine nozzle. One or more additional fuel injection nozzles are arranged at an aft end of the transition duct for introducing additional fuel into the transition duct upstream of the first stage turbine nozzle.

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Greer, SC, United States of America

Mark Allan Hadley