Siemens Energy, Inc.

Siemens Energy Global Gmbh & Co. Kg

Siemens Power Generation, Inc.

Siemens Aktiengesellschaft

Westinghouse Electric Corporation

Siemens Westinghouse Power Corporation

Walter Ray Laster

Juan Enrique Portillo Bilbao

Timothy A Fox

Andrew J North

Ramarao V Bandaru

Reinhard Schilp

Jacob William Hardes

Yuntao Chen

Robert H Bartley

Scott M Martin

Joseph Meadows

Joan Mae Baer

Lucky Tran

Gerald Joseph Bruck

David James Wiebe

David Allen Little

Douglas A Keller

Stephen A Ramier

Roy E Wootton

Samer P Wasif

Weidong Cai

Muzaffer Sutcu

David B Berrong

Domenico Gambacorta

Krishna C Miduturi

Peter Szedlacsek

Elvira Vladimirovna Anoshkina

Kyle L Landry

Richard A Evans

George W Sanford

Elizabeth R Colore

Mario Gaviria

R Michael Roidt

Juan Enrique Portilllo Bilbao

A finely distributed combustion system for a gas turbine engine is provided. A combustor body may extend along a longitudinal axis. A premixer space may be formed within the combustor body to premix air and fuel. The premixer space is in communication with an array of finely distributed perforations arranged in a wall of the combustor body to eject an array of premixed main flamelets throughout a contour of the combustor body between the upstream base of the combustor body and the downstream base of the combustor body. The array of finely distributed perforations&#x2014;potentially comprising thousands or even hundreds of thousands of perforations spatially distributed on a miniaturized scale&#x2014;for ejecting the premixed main flamelets is technically advantageous compared to conventional distributed combustion systems, where injection of relatively longer main flames occurs just at a few discrete axial locations.

Finely distributed combustion system for a gas turbine engine

A turbine engine has two combustion zones so as to operate in conditions where water scarcity is an issue. The secondary combustion zone is located downstream of the primary combustion zone. Fuel can be fed into an air scoop having air from a shell surrounding the primary and secondary combustion zone. The feeding of the fuel through the air scoop allows atomization of the fuel. The mixture can then enter the secondary combustion zone and mix with the products from the first combustion zone.

Fuel oil axial stage combustion for improved turbine combustor performance

Injector assemblies () and ducting arrangement () including such injector assemblies are provided. The injector assembly may include a reactant-guiding structure () arranged to convey a flow of reactants () into the combustion stage and means for injecting () a flow of air () into the combustion stage. The flow of air injected into the combustion stage may be arranged to condition interaction of the flow of reactants injected into the combustion stage with a cross-flow of combustion products (), as the flow of reactants is admitted into the combustion stage.

Ducting arrangement with injector assemblies configured to form a shielding flow of air injected into a combustion stage in a gas turbine engine

A fuel injector for injecting alternate fuels having a different energy density in a gas turbine is provided. A first fuel supply channel () may be fluidly coupled to a radial passage () in a plurality of vanes () that branches into passages () (e.g., axial passages) to inject a first fuel without jet in cross-flow injection. This may be effective to reduce flashback in fuels having a relatively high flame speed. A mixer () with lobes () for injection of a second fuel may be arranged at the downstream end of a fuel delivery tube (). A fuel-routing structure () may be configured to route the second fuel within a respective lobe so that fuel injection of the second fuel takes place radially outwardly relative to a central region of the mixer. This may be conducive to an improved (e.g., a relatively more uniform) mixing of air and fuel.

Fuel injector including a lobed mixer and vanes for injecting alternate fuels in a gas turbine

A cooling system for a fuel system in a turbine engine that is usable to cool a fuel nozzle is disclosed. The cooling system may include one or more cooling system housings positioned around the fuel nozzle, such that the cooling system housing forms a cooling chamber defined at least partially by an inner surface of the cooling system housing and an outer surface of the fuel nozzle. The fuel nozzle may extend into a combustor chamber formed at least in part by a combustor housing. The fuel nozzle may include one or more fuel exhaust orifices with an opening in an outer surface of the fuel nozzle and configured to exhaust fluids unrestricted by the housing forming the cooling system cooling chamber. The cooling system may provide cooling fluids to cool the fuel nozzle within the cooling system cooling chamber regardless of whether the fuel nozzle is in use.

Cooling system for fuel nozzles within combustor in a turbine engine

A flow conditioning device for a can annular gas turbine engine, including a plurality of flow elements disposed in a compressed air flow path leading to a combustor, configured such that relative adjustment of at least one flow directing element with respect to an adjacent flow directing element during operation of the gas turbine engine is effective to adjust a level of choking of the compressed air flow path.

Device for improved air and fuel distribution to a combustor

A ducting arrangement () in a combustion stage downstream of a main combustion stage of a gas turbine engine is provided. A duct () is fluidly coupled to receive a cross-flow of combustion gases from the main combustion stage. Duct () includes a duct segment () with an expanding cross-sectional area () where one or more injector assemblies () are disposed. Injector assembly () includes one or more reactant-guiding structures () arranged to deliver a flow of reactants into the downstream combustion stage to be mixed with the cross-flow of combustion gases. Disclosed injector assemblies are arranged in expanding cross-sectional area () to reduce total pressure loss while providing an effective level of mixing of the injected reactants with the passing cross-flow. Respective duct components or the entire ducting arrangement may be formed as a unitized structure, such as a single piece using a rapid manufacturing technology, such as 3D Printing/Additive Manufacturing (AM) technologies.

Ducting arrangement with injector assemblies arranged in an expanding cross-sectional area of a downstream combustion stage in a gas turbine engine

Method and computer-readable model for additively manufacturing a ducting arrangement () for a gas turbine engine are provided. Ducting arrangement () may include a duct () to be fluidly coupled to receive a cross-flow of combustion gases from a main combustion stage. Duct () includes a duct segment () with an expanding cross-sectional area () where one or more injector assemblies () are disposed. Injector assembly () includes one or more reactant-guiding structures () arranged to deliver a flow of reactants to be mixed with the cross-flow of combustion gases. The ducting arrangement is effective to reduce total pressure loss while providing an effective level of mixing of the injected reactants with the passing cross-flow. Respective duct components or the entire ducting arrangement may be formed as a unitized structure, such as a single piece using a rapid manufacturing technology, such as 3D Printing/Additive Manufacturing (AM) technologies.

Method and computer-readable model for additively manufacturing ducting arrangement for a gas turbine engine

A secondary fuel stage () of a combustor of a gas turbine engine. The combustor has a main combustion zone () upstream of the secondary fuel stage to ignite working gas The secondary fuel stage includes a nozzle () with dual outlets () oriented with a circumferential component to inject an air-fuel mixture () into the combustor. The nozzle is effective to entrain the air-fuel mixture with the working gas () such that a peak temperature () of the working gas at a location downstream of the secondary fuel stage is less than a peak temperature () of working gas if the air-fuel mixture were injected into the combustor with a single outlet nozzle ().

Dual outlet nozzle for a secondary fuel stage of a combustor of a gas turbine engine

Method and computer-readable model for additively manufacturing a ducting arrangement () in a combustion stage are provided. The ducting arrangement may include a combustor wall () in a combustion stage fluidly coupled to receive a cross-flow of combustion products (). An injector assembly () may be in fluid communication with cooling fluid conduits () in the combustor wall to receive cooling fluid that passes through the cooling fluid conduits. The injector assembly may include means for injecting () a flow of the cooling fluid () arranged to condition interaction of a flow of reactants () injected with the cross-flow of combustion products. Respective duct components or the entire ducting arrangement may be formed as a unitized structure, such as a single piece using a rapid manufacturing technology, such as 3D Printing/Additive Manufacturing (AM) technologies.

Method and computer-readable model for additively manufacturing ducting arrangement with injector assemblies forming a shielding flow of air

A fuel injector for injecting fuels having a different energy density in a gas turbine is provided. A first fuel supply channel () and a second fuel supply channel () may be coaxially arranged in a fuel delivery structure (). A first set of vanesincludes a radial passage () in fluid communication with the first channel () to receive a first fuel. Passage () branches into passages () each having an aperture () to inject the first fuel without jet in cross-flow injection. A second set of vanes () includes a radial passage () in fluid communication with the second channel () to receive a second fuel. Passage () branches into passages () each having an aperture () arranged to inject the second fuel also without jet in cross-flow injection. This arrangement may be effective to reduce flashback that otherwise may be encountered in fuels having a relatively high flame speed.

Fuel injector including tandem vanes for injecting alternate fuels in a gas turbine

An apparatus and method for lean/rich combustion in a gas turbine engine (), which includes a combustor (), a transition () and a combustor extender () that is positioned between the combustor () and the transition () to connect the combustor () to the transition (). Openings () are formed along an outer surface () of the combustor extender (). The gas turbine () also includes a fuel manifold () to extend along the outer surface () of the combustor extender (), with fuel nozzles () to align with the respective openings (). A method () for axial stage combustion in the gas turbine engine () is also presented.

Lean-rich axial stage combustion in a can-annular gas turbine engine

A flow conditioning device () for a can annular gas turbine engine, including a plurality of flow elements () disposed in a compressed air flow path () leading to a combustor (), configured such that relative adjustment of at least one flow directing element () with respect to an adjacent flow directing element () during operation of the gas turbine engine is effective to adjust a level of choking of the compressed air flow path ().

A fuel gas cooling system for a combustion basket spring clip seal support is disclosed. The fuel gas cooling system may be formed from one or more fuel gas supply channels terminating proximate to a spring clip at the intersection between a combustor basket and a transition section such that fuel gas may be supplied to the hot gas path proximate to the intersection between the combustor basket and the transition section. The fuel gas supply channel may create an intermediate fuel gas burn at this intersection, which may reduce the firing temperature at the fuel nozzles and reduce NOx emissions.

Fuel gas cooling system for combustion basket spring clip seal support

An emissions control system for a gas turbine engine including a flow-directing structure () that delivers combustion gases () from a burner () to a turbine. The emissions control system includes: a conduit () configured to establish fluid communication between compressed air () and the combustion gases within the flow-directing structure (). The compressed air () is disposed at a location upstream of a combustor head-end and exhibits an intermediate static pressure less than a static pressure of the combustion gases within the combustor (). During operation of the gas turbine engine a pressure difference between the intermediate static pressure and a static pressure of the combustion gases within the flow-directing structure () is effective to generate a fluid flow through the conduit ().

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Oviedo, FL, United States of America

Walter Ray Laster