Siemens Energy, Inc.

Siemens Aktiengesellschaft

Siemens Energy Global Gmbh & Co. Kg

Juan Enrique Portillo Bilbao

Walter Ray Laster

Andrew J North

Timothy A Fox

Bernd Prade

Scott M Martin

Rajesh Rajaram

Yuntao Chen

Jacob William Hardes

Danning You

Joseph Meadows

Lucky Tran

Olga Deiss

David M Ritland

Christian Beck

Kevin J Spence

Kevin M Spence

Jared M Pent

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

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 combustor and a method involving burner mains structurally configured to damp vibrational modes that can develop under high-frequency combustion dynamics are provided. The combustor may include a carrier (), and a plurality of mains () disposed in the carrier. Some of the mains (labeled with the letter X) include a body having a different structural feature relative to the respective bodies of the remaining mains. The mains with the different structural feature may be selectively grouped in the carrier to form at least one set of such mains effective to damp predefined vibrational modes in the combustor.

Combustor and method for damping vibrational modes under high-frequency combustion dynamics

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 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 combustor () including: a first premix main burner () comprising a first swirler airfoil section (); a second premix main burner () comprising a second swirler airfoil section (); and a supply air reversing region upstream of the premix burners (), (). The first swirler airfoil section () and the second swirler airfoil section () are effective to impart swirl to a first airflow and a second airflow characterized by a same swirl number as the airflows exit respective burners (), (). The combustor () is effective to generate a first airflow volume through the first premix main burner () that is different than a second airflow volume through the second premix main burner ().

Air biasing system in a gas turbine combustor

An acoustically dampened gas turbine engine having a gas turbine engine combustor with an acoustic damping resonator system is disclosed. The acoustic damping resonator system may be formed from one or more resonators positioned within the gas turbine engine combustor at an outer housing forming a combustor basket and extending circumferentially within the combustor. The resonator may be positioned in a head region of the combustor basket. In one embodiment, the resonator may be positioned in close proximity to an intersection between the outer housing and an upstream wall defining at least a portion of the combustor. The acoustic damping resonator system may mitigate longitudinal mode dynamics thereby increasing an engine operating envelope and decreasing emissions.

Acoustic damping system for a combustor of a gas turbine engine

A combustor assembly in a gas turbine engine includes a liner defining a combustion zone, at least one fuel injector for providing fuel, and a flow sleeve. An inner surface of the flow sleeve defines an outer boundary for an air flow passageway. Upon the air reaching a head end of the combustor assembly at an end of the air flow passageway the air turns 180 degrees to flow into the combustion zone where it is burned with the fuel. The combustor assembly further includes an inlet assembly positioned radially between the liner and the flow sleeve. The inlet assembly defines an inlet to the air flow passageway and includes a plurality of overlapping conduits that are arranged such that the air entering the air flow passageway passes through radial spaces between adjacent conduits.

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Oviedo, FL, United States of America

Juan Enrique Portillo Bilbao