Siemens Energy, Inc.

Siemens Aktiengesellschaft

Siemens Energy Global Gmbh & Co. Kg

Siemens Aktiengesell Schaft

Mikro Systems, Inc.

Timothy A Fox

Walter Ray Laster

Reinhard Schilp

Stephen A Ramier

Steven Williams

David James Wiebe

Juan Enrique Portillo Bilbao

Andreas Böttcher

James Bertoncello

William Richard Ryan

Anthony L Schiavo

Tobias Krieger

David M Ritland

Thomas Grieb

Jacob William Hardes

Robert H Bartley

Daniel W Garan

John Carl Glessner

Matthias Hase

Scott M Martin

Andrew J North

Abhijeet Tiwary

Weidong Cai

Jens Kleinfeld

Muzaffer Sutcu

Ulrich Wörz

Jaap Van Kampen

Vinayak V Barve

Jürgen Meisl

Birgit Grüger

Domenico Gambacorta

Holger Streb

Benjamin G Hettinger

Sachin Terdalkar

Kyle L Landry

Christin Schlensog

Siegfried Bode

Kurtis Dickerson

William F Van Nieuwenhuizen

Alexander Hyatt

Marcus Zurhorst

Sebastian Pfadler

Stephen A Raimer

Andreas BÖTTCHER

A gas turbine engine has a converging duct that has combustion products flow at low mach speeds through a first portion and a high mach speeds through a second portion. The converging duct has two types of cooling schemes formed. One type of cooling scheme is beneficial for the low mach speed combustion product flow and one type of cooling scheme is beneficial for the high mach speed combustion product flow. The two cooling schemes are blended together in order increase the efficiency of the cooling of the converging duct.

Converging duct with elongated and hexagonal cooling features

Method for constructing impingement/effusion cooling features in a component of a combustion turbine engine is provided. A pocketmay be arranged between an outer walland an inner wallof the component. A lasing deviceallows drilling through the component to form an effusion hole. The lasing device further allows welding closed an openingformed at outer wallof the component during the drilling with the lasing device through the component. Lasing devicefurther allows drilling through outer wallof the component to form an impingement holefor the impingement/effusion cooling feature. The proposed methodology in a multi-panel arrangement, for example, eliminates a need of having to pre-drill such holes in individual panels prior to the bonding and forming of the component, which overcomes various drawbacks commonly associated with such pre-drilling.

Method for constructing impingement/effusion cooling features in a component of a combustion turbine engine

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 fuel burner system () configured to inject a liquid fuel and a gas fuel into a combustor () of a turbine engine () such that the engine () may operate on the combustion of both fuel sources () is disclosed. The fuel burner system () may be formed from a nozzle cap () including one or more first fuel injection ports () in fluid communication with a first fuel source () of syngas and one or more second fuel injection ports () in fluid communication with a second fuel source () of natural gas. The fuel burner system () may also include an oil lance () with one or more oil injection passages () that is in fluid communication with at least one oil source () and is configured to emit oil into the combustor (). The oil lance () may include one or more fluid injection passages () configured to emit air to break up the oil spray and water to cool the combustor (), or both.

Flexible fuel combustion system for turbine engines

The present disclosure provides a gas turbine combustor liner () comprising an outer surface () and an inner surface (), a plurality of film cooling holes () through a thickness of the gas turbine combustor liner (), and a plurality of resonator boxes () affixed to the outer surface () of the gas turbine combustor liner (). The film cooling holes () extend circumferentially around the gas turbine combustor liner () and comprise a first set of holes () having a first axial row spacing X and a second set of holes () having a second axial row spacing X'. The second set of holes () is formed in the gas turbine combustor liner () in a downstream direction relative to the first set of holes (). The second axial row spacing X&#x2032; is greater than the first axial row spacing X.

Film cooling hole arrangement for acoustic resonators in gas turbine engines

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 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

An annular fluid distribution device () for distributing fluid into a gaseous flow (), that includes: a first fluid distribution manifold () having a first fluid inlet and first fluid outlets (), wherein the first fluid outlets inject a first fluid into the gaseous flow; and a second fluid distribution manifold (), having a second fluid inlet and second fluid outlets (), wherein the second fluid outlets inject a second fluid into the gaseous flow. The second fluid manifold is isolated from the first fluid distribution manifold, and the first fluid outlets and the second fluid outlets are disposed on a common fluid outlet plane ().

Circumferential biasing and profiling of fuel injection in distribution ring

A pilot fuel nozzle for a combustor includes an igniter forming a central body extending along a longitudinal center of the nozzle. A nozzle tip includes a plurality of circumferentially spaced fuel passages and a plurality of circumferentially spaced air passages extending to an outer side of the nozzle tip. The central body extends through a center of the nozzle tip for producing a spark to ignite a fuel/air mixture adjacent to the nozzle tip. A plurality of fuel tubes extend along the central body, each of the fuel tubes having an outlet end engaged on the nozzle tip for delivery of fuel from the nozzle tip into a combustion chamber of the combustor An outer sleeve surrounds the fuel tubes and defines an annular space in fluid communication with the air passages of the nozzle tip between the outer sleeve and the central body.

Nozzle with multi-tube fuel passageway for gas turbine engines

An assembly for a gas turbine is presented. The assembly includes a gas supply pipe passing through a bore in a flange of the gas turbine for supplying gas to a combustion chamber of the gas turbine and a sleeve surrounding the gas supply pipe, having a first end and a second end, wherein the first end is sealingly coupled to the gas supply pipe, and wherein the sleeve is adapted to be sealingly coupled to the flange at the second end such that the sleeve extends along a thickness of the flange.

Gas turbine assembly and method therefor

A combustor () for a gas turbine engine () has a circumferentially extending liner () defining at least a portion of an interior combustion chamber () and a hot gas path (). The liner includes a resonator section () including at least one resonator (A,B) having a resonator chamber (A,B) formed on an exterior of the liner. A thermal barrier coating () is disposed along an inner surface of the liner including an inner surface () of the resonator section. The resonator further includes a plurality of apertures () and each aperture extends through the liner and the thermal barrier coating at the resonator section and there is fluid flow communication between the combustion chamber and the resonator chamber.

Combustor resonator section with an internal thermal barrier coating and method of fabricating the same

A gas turbine combustor assembly including a combustor liner and a plurality of fuel nozzle assemblies arranged in an annular array extending within the combustor liner. The fuel nozzle assemblies each include fuel nozzle body integral with a swirler assembly, and the swirler assemblies each include a bellmouth structure to turn air radially inwardly for passage into the swirler assemblies. A radially outer removed portion of each of the bellmouth structures defines a periphery diameter spaced from an inner surface of the combustor liner, and an interface ring is provided extending between the combustor liner and the removed portions of the bellmouth structures at the periphery diameter.

Interface ring for gas turbine fuel nozzle assemblies

A casting () for a support housing (), including: a fuel manifold () comprising an a-stage gas gallery () and a b-stage gas gallery (); a-stage and a b-stage rocket bases () integrally cast with the fuel manifold (), the a-stage gas gallery () in fluid communication with the a-stage stage rocket bases and the b-stage gas gallery () in fluid communication with the b-stage rocket bases; and one oil tube passageway () for each fuel rocket base (), each oil tube passageway () spanning from an upstream end () of the fuel manifold () to an interior () of a respective fuel rocket base (). Each oil tube passageway () is disposed radially inward of an inner perimeter () of the b-stage gas gallery ().

Cast manifold for dry low NOx gas turbine engine

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Hamilton, Canada

Timothy A Fox