Siemens Energy, Inc.

Siemens Westinghouse Power Corporation

Gas Turbine Efficiency Sweden Ab

General Electric Company

Siemens Power Generation, Inc.

Beijing Huatsing Gas Turbine & Igcc Technology Co., Ltd

Siemens Aktiengesellschaft

Robert James Bland

John L Battaglioli

William Richard Ryan

Clifford E Johnson

Samer P Wasif

John Luigi Battaglioli

Anil Gulati

Stephen A Ramier

Adam Bailey

John Barnes

Shanshan Zhang

Peter Tiemann

David James Wiebe

Bernd Prade

John Carl Glessner

Charles A Ellis

Richard Sterling Tuthill

Robert W Dawson

Guy Wayne DeLeonardo

William Theodore Bechtel

Stephen Hugh Black

Scott M Martin

Ramarao V Bandaru

Stefan Dahlke

Michael H Koenig

Bradley T Youngblood

Domenico Gambacorta

Dennis A Horazak

Jeffrey Arthur Benoit

Udo Schmitz

Karsten Jordan

Joseph Charles Taura

Stefan Martin Meyer

Xiaochen Zha

J Scott Markovitz

David Kargetta

Stefan Dahlk

Xiaoehen Zha

John Battaglioli

Domenlco Gambacorta

Stephen A Raimer

Scott J Markovitz

A method of optimizing a premix fuel nozzle for a gas turbine is provided including the step of providing a nozzle that provides, when the gas turbine is in operation, an axial flow field of an air and fuel mixture flows through the burner tube and around the nozzle tip, and at least two recirculation zones with at least two differing radial extents as part of a toroidal vortex generated on the nozzle tip to provide strong flame holding and flame propagation.

Method of optimizing premix fuel nozzles for a gas turbine

A nozzle tip for a premix fuel nozzle includes an outer body having an outer body external face facing the downstream end of the burner tube, the outer body external face having a smaller cross-sectional area than the cross-sectional area of the burner tube; and at least one segment radiating radially outwardly toward the internal wall of the burner tube from the outer body, wherein each segment has a proximal end disposed adjacent to the outer body external face and a distal end disposed in a direction toward the burner tube, wherein each segment has a segment downstream face angled relative to the longitudinal axis of the burner tube towards the downstream end of the burner tube. When the gas turbine is in operation, an axial flow field of an air and fuel mixture flows through the burner tube and around the nozzle tip, and two or more recirculation zones of differing radial extent are generated on the nozzle tip by the segments to provide strong flame holding and flame propagation.

Premix fuel nozzle for a gas turbine and combustor

A turbine engine having a plenum for passing fluids from an outlet of a compressor to an inlet of a combustor that may increase the efficiency of the turbine engine. The turbine engine may include a combustor, a compressor positioned upstream of the combustor, a transition channel extending from the compressor to the combustor, and a shell extending between the compressor and a combustor portal and positioned around the at least one transition channel. The turbine engine may also include an axial diffusor in the shell near the at least one transition channel, wherein the axial diffusor may include a fluid flow recess in a trailing edge of the axial diffusor. The turbine engine may also include a wave protrusion extending from a surface positioned radially inward of the axial diffusor. The fluid flow recess and the wave protrusion may reduce fluid flow loss within the shell.

Axial diffusor for a turbine engine

A pilot burner () for a gas turbine engine delivers an inner non-swirling fuel-oxidant mixture surrounded by an outer swirling fuel-oxidant mixture, thereby providing enhanced mixing with no recirculation zone. At least one fluid-restricting inlet port () provides an oxidant to an inner mixing passage (). The inner mixing passage () includes a plurality of fuel outlets (). An outer annular mixing passage () receives oxidant from an upstream port () surrounding the inner mixing passage and includes at least one swirler element () and fuel outlets ().

Pilot burner for gas turbine engine

Embodiments of the present invention provide resonators () that have lateral walls () disposed at non-square angles relative to the liner's longitudinal (and flow-based) axis () such that a film cooling of substantial portions of an intervening strip () is provided from apertures (A,B,) in a resonator box () adjacent and upstream from the intervening strip (). This film cooling also cools weld seams () along the lateral walls () of the resonator boxes (). In various embodiments the lateral wall angles are such that film cooling may be provided to include the most of the downstream portions of the intervening strips (). These downstream portions are closer to the combustion heat source and therefore expected to be in greater need of cooling.

Non-rectangular resonator devices providing enhanced liner cooling for combustion chamber

A gas turbine engine is provided comprising an outer shell, a compressor assembly, at least one combustor assembly, a turbine assembly and duct structure. The outer shell includes a compressor section, a combustor section, an intermediate section and a turbine section. The intermediate section includes at least one first opening and at least one second opening. The compressor assembly is located in the compressor section to define with the compressor section a compressor apparatus to compress air. The at least one combustor assembly is coupled to the combustor section to define with the combustor section a combustor apparatus. The turbine assembly is located in the turbine section to define with the turbine section a turbine apparatus. The duct structure is coupled to the intermediate section to receive at least a portion of the compressed air from the compressor apparatus through the at least one first opening in the intermediate section, pass the compressed air to an apparatus for separating a portion of oxygen from the compressed air to produced vitiated compressed air and return the vitiated compressed air to the intermediate section via the at least one second opening in the intermediate section.

Gas turbine engine adapted for use in combination with an apparatus for separating a portion of oxygen from compressed air

A gas sleeve () for a combustor () of gas turbine engine () attaches to a support housing () of the combustor () to convey a fuel gas and to fit within a fuel rocket (). The gas sleeve () comprises a plurality of apertures () formed to provide impingement cooling. The apertures () comprise a tilt angle directed toward a structure in need of impingement cooling, for instance a weld joint () that attaches the fuel rocket () to the support housing (). The apertures () additionally may comprise a rotational angle effective to create a rotationally swirling flow of the portion of fuel gas that passes through the apertures (). A method of operation using this structure also is provided.

Extended life fuel nozzle

A curved diffuser () in a gas turbine engine () directs a primary portion of air flow from a compressor () through a curved discharge opening () into a plenum (). The curved diffuser () also comprises ports () through which a secondary portion of air passes into confined space () that is defined in part by a pressure boundary element that may be comprised of at least one plate () or at least one conduit (). The at least one plate () and the at least one conduit () respectively comprise apertures () through which pass the secondary portion of air to provide impingement-type cooling to transitions (). In various embodiments the velocity of the air between adjacent transitions () may flow at relatively uniform velocity along the longitudinal distance of the respective transitions ().

Gas turbine engine curved diffuser with partial impingement cooling apparatus for transitions

A combustion apparatus and duct structure for a gas turbine engine are provided. The combustion apparatus comprises a combustion system to receive fuel and air, ignite at least a portion of the fuel and air and output a stream of first combustion products and any remaining fuel and air. The combustion apparatus further comprises structure positioned adjacent to the combustion system for receiving and accelerating the first combustion products and any remaining fuel and air from the combustion system. The duct structure receives the first combustion products and any remaining fuel and air from the combustion apparatus, allows any remaining fuel and air to combust to generate second combustion products, accelerates the first and second combustion products and outputs the first and second combustion products to a first row blade assembly.

At least one combustion apparatus and duct structure for a gas turbine engine

A transition () for a gas turbine engine () comprises a transition wall () comprising cooling channels (L,R,L,R) that are adapted to pass a cooling fluid, such as compressed air from a compressor () during operation of the gas turbine engine () so as to cool combusted hot gases passing through the transition (). For each cooling channel (L,R,L,R), respective entry ports (L,L,L,R) and exit ports (L,R,L,R) are arranged so as to obtain a performance improvement based upon pressure differentials between the respective entry and exit ports. In various embodiments, a scoop () is associated with an entry port (L,R) so as to establish a more elevated pressure differential in the respective cooling channel (L,R). An entry port (-) may be positioned offset relative to a lower-positioned exit port (-) so as to so minimize or eliminate intake of heated airflow from a respective nearby exit port (-). Such offset positioning may be based on the airflow paths and cooling requirements at selected high-temperature operating conditions.

Gas turbine engine transitions comprising closed cooled transition cooling channels

A method and system for controlling combustion in a gas turbine engine () includes reducing an overall fuel flow provided to a stage () of burners (e.g.) of the gas turbine engine until reaching a predetermined dynamic operating condition of the first burner of the stage. The method also includes maintaining, while continuing to reduce the overall fuel flow (e.g.), a first portion (e.g.), of the overall fuel flow delivered to the first burner at a maintenance level so that the predetermined dynamic operating condition of the first burner is maintained.

Fuel flow tuning for a stage of a gas turbine engine

A burner () of a gas turbine engine () includes a cylindrical basket () comprising an air flow reversal region (). The flow reversal region ends at an air inlet plane () of the basket. The burner also includes a flow conditioner () disposed in the flow reversal region transecting an air flow () flowing non-uniformly through the flow reversal region, the flow conditioner being effective to mitigate variation of the air flow entering the basket across the inlet plane.

Air flow conditioner for a combustor can of a gas turbine engine

A combustor for a gas turbine is provided having a nozzle assembly located at one end and a combustion chamber defining a second end of the combustor. A venturi is positioned within the combustor, between the nozzle and the combustion chamber. The venturi defines a passageway therein having a first side facing the nozzle and a second side facing the combustion chamber. Compressed air is directed into an inlet in fluid communication with the first and second sides of the venturi passageway. The venturi passageway directs the compressed air from the inlet in opposite directions within the first and second sides of the passageway for cooling the venturi.

Venturi cooling system

A secondary nozzle is provided for a gas turbine. The secondary nozzle includes a flange and an elongated nozzle body extending from the flange. At least one premix fuel injector is spaced radially from the nozzle body and extends from the flange generally parallel to the nozzle body. At least one second nozzle tube is fluidly connected to the fuel source and spaced radially outward from the first nozzle tube with a proximal end fixed to the flange. The second nozzle tube has a distal end, spaced from the proximal end, with at least one aperture therein. A passageway extends between the proximal end and the distal end of the second nozzle tube, with the passageway fluidly connecting to the fuel source and the aperture.

Combustor nozzle

A combustor for a gas turbine is provided having a nozzle assembly located at one end and a combustion chamber defining a second end of the combustor. A venturi is positioned within the combustor, between the nozzle and the combustion chamber. The venturi defines an internal passageway therein having a first side facing the nozzle and a second side facing the combustion chamber. Compressed air is directed into an inlet in fluid communication with the first and second sides of the internal passageway. The internal passageway directs the compressed air from the inlet in opposite directions within the first and second sides of the internal passageway for cooling the venturi.

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Clifton Park, NY, United States of America

Robert James Bland