Alstom Technology Limited

Ansaldo Energia Switzerland Ag

General Electric Company

Ansaldo Energia IP UK Limited

General Electric Technology Gmbh

Bruno Schuermans

Mirko Ruben Bothien

Nicolas Noiray

Christian Oliver Paschereit

Andreas Huber

Jaan Hellat

Peter Flohr

Valter Bellucci

Adnan Eroglu

Urs Benz

Andre Theuer

Majed Toqan

Martin Zajadatz

Hanspeter Zinn

Jost Imfeld

Luis Tay Wo Chong Hilares

Dirk Bueche

Rolf Dornberger

Petros Koumoutsakos

Peter Stoll

Lucas John Stoia

Madhavan Narasimhan Poyyapakkam

Eribert Benz

Stefano Bernero

Michael Düsing

Alexander Schnell

Felix Guethe

Timothy James Purcell

Fulvio Magni

Frank Graf

Dariusz Oliwiusz Palys

Sven Georg Bethke

Devis Tonon

Felix Michael Reinert

Alexander Ni

Ken-Yves Haffner

Naresh Aluri

Jeffrey Alan De Jonge

Diane Lauffer

Ghislain Singla

Michael Duesing

Pirmin Schiessel

Norbert Walter Emberger

Mauricio Ernesto Garay

Piotr Siewert

Alessandro Scarpato

Simone Roberto Walter Camponovo

Danda-Raj Pahari

Dejan Rajkovic

Zdenko Papa

Rainer Conzelmann

Luca Tentorio

Adrian Glauser

Francois Meili

Daniele Tabacco

Dominik Wassmer

Franklin Genin

Naresh Kumar Aluri

Piotr Siewart

The present application provides a variable frequency Helmholtz damper system for use with a combustor of a gas turbine engine. The variable frequency Helmholtz damper system may include one or more Helmholtz dampers and a purge medium temperature control unit for providing a flow of purge medium to the Helmholtz dampers. The purge medium temperature control unit may be in communication with a temperature control fluid flow such that the purge medium temperature control unit may vary the temperature of the flow of purge medium.

Variable frequency Helmholtz dampers

A fuel nozzle assembly includes a fuel plenum for delivering fuel to a combustion zone of a combustor. At least one narrow-band acoustic damper is provided in the fuel nozzle assembly. The narrow-band acoustic damper is tuned to dampen pressure oscillations within the combustor. The narrow-band acoustic damper may be a quarter-wave tube mounted to a cap plate within or adjacent to the fuel nozzle assembly.

Fuel nozzle with narrow-band acoustic damper

The invention concerns a damper arrangement for reducing combustion-chamber pulsation arising inside a gas turbine, wherein the gas turbine includes at least one compressor, a primary combustor which is connected downstream to the compressor, and the hot gases of the primary combustor are admitted at least to an intermediate turbine or directly or indirectly to a secondary combustor. The hot gases of the secondary combustor are admitted to a further turbine or directly or indirectly to an energy recovery, wherein at least one combustor is arranged in a can-architecture. At least one combustor liner includes air passages, wherein at least one of the air passages is formed as a damper neck. The damper neck being actively connected to a damper volume, and the damper volume is part of a connecting duct extending between a compressor air plenum and the combustor.

Damper arrangement for reducing combustion-chamber pulsation

The damper arrangement includes a plurality of interconnected volumes and a plurality of necks for connecting the damper to a combustion chamber at a plurality of contact points. The plurality of necks are connected to the plurality of volumes.

Simultaneous broadband damping at multiple locations in a combustion chamber

A sequential combustor arrangement and method are disclosed which can include a first burner, a first combustion chamber, a mixer for admixing a dilution gas to the hot gases leaving the first combustion chamber during operation, a second burner, and a second combustion chamber arranged sequentially in a fluid flow connection. The mixer can include at least one injection opening in the mixer wall for admixing the dilution gas to cool the hot flue gases leaving the first combustion chamber. Further, the mixer can include a damper with a damper volume and a neck connecting the damper volume to the mixer, for modulating and damping pressure pulsations inside the mixer.

Sequential combustor arrangement with a mixer

The invention refers to a sequential combustor arrangement including a first burner, a first combustion chamber, a mixer for admixing a dilution gas to the hot gases leaving the first combustion chamber during operation, a second burner, and a second combustion chamber arranged sequentially in a fluid flow connection. The mixer includes a plurality of injection pipes pointing inwards from the side walls of the mixer for admixing the dilution gas to cool the hot flue gases leaving the first combustion chamber with a low pressure drop. The disclosure further refers to a method for operating a gas turbine with such a combustor arrangement.

Sequential combustion with dilution gas mixer

The invention relates to a damper for reducing pulsations in a gas turbine, which includes an enclosure, a main neck extending from the enclosure, a spacer plate disposed in the enclosure to separate the enclosure into a first cavity and a second cavity and an inner neck with a first end and a second end, extending through the spacer plate to interconnect the first cavity and the second cavity. The first end of the inner neck remains in the first cavity and the second end remains in the second cavity. A flow deflecting member is disposed proximate the second end of the inner neck to deflect a flow passing through the inner neck. With the solution of the present invention, as a damper according to embodiments of the present invention operates, flow field hence damping characteristic in the second cavity constant regardless the adjustment of the spacer plate in the enclosure.

Damper for gas turbine

A method is disclosed for operating an axially staged mixer arrangement with dilution air or dilution gas injection in connection with a combustion arrangement of a gas turbine engine or turbo machinery, having at least one combustion chamber, at least one mixer arrangement for admixing air or gas portions to the hot gas flow leaving the combustion chamber. The mixer arrangement includes a plurality of injectors for admixing air portions to cool at least the hot gas flow leaving combustion chamber. The spacing between the last injection location of acting injectors and at least one subsequently arranged dilution air injection, inside of the mixer arrangement in the hot gas flow, where there is no node between the injection locations, corresponds to a distance equal or approximating to half of convective wave length.

Axially staged mixer with dilution air injection

A damping system and apparatus are disclosed for dampening acoustic pressure oscillations of a gas flow in a combustor of a gas turbine engine having at least one combustor with a combustor liner. A second inner cap portion is disposed on the at least one combustor inner liner. The second inner cap portion can have a hot surface, a cold surface, at least one burner opening protruding from the cold surface, and at least one neck ring having an internal opening and protruding from the cold surface. At least one resonating tube having a resonating tube neck is integrated with and protruding from the at least one neck ring. The at least one resonating tube is disposed between adjacent burner openings, and is configured such that the radial dimension is greater than or equal to the axial dimension.

System and apparatus for gas turbine combustor inner cap and resonating tubes

A damping system and apparatus are disclosed for dampening acoustic pressure oscillations of a gas flow in a combustor of a gas turbine engine having at least one combustor with a combustor liner. A second inner cap portion is disposed on the at least one combustor. The second inner cap portion can have a hot surface, a cold surface, at least one burner opening protruding from the cold surface, and at least one neck ring having an internal opening and protruding from the cold surface. At least one extended resonating tube having a resonating tube neck is integrated with and protruding from the at least one neck ring. The at least one extended resonating tube is disposed between adjacent burner openings, and is configured such that the radial dimension is less than the axial dimension.

System and apparatus for gas turbine combustor inner cap and extended resonating tubes

A method for determining the temperature of an aggressive and/or abrasive gas is described using an acoustic transmitter and an acoustic receiver. The transmitter emits an acoustic signal with varying frequencies and the receiver extracts from the acoustic input signal a frequency of the maximum. Based on this the frequency of this maximum the temperature of the gas between transmitter and receiver is calculated.

Method for determining the temperature inside a combustor

The invention relating to a A damper for combustion oscillation damping for a gas turbine, includes a resonator cavity with a box or cylinder shape, a neck in flow communication with the resonator cavity and a combustion chamber. The length Lof the neck is equal or greater than a maximum distance lthat a hot gas ingested from the combustion chamber reaches in the neck. The maximum distance lis determined according to the following equation:  With the solution of the present invention, the damper used to damping combustion oscillation may prevent hot gas ingestion during normal operations by uniquely determining the dimensions of the neck connecting the resonator cavity and the combustion chamber.

Damper for combustion oscillation damping in a gas turbine

Gas turbine damper with inner neck extending into separate cavities

A combustion device () for a gas turbine includes portions () having an inner and an outer wall () with an interposed noise absorption plate () having a plurality of holes (). The combustion device () further has first passages () connecting zones between the inner wall () and the plate () to the inside of the combustion device () and second passages () for cooling the inner wall (). The portions () also have an inner layer () between the inner wall () and the plate () defining inner chambers (), each connected to at least a first passage (), and an outer layer () between the outer wall () and the plate () defining outer chambers () connected to the inner chambers () via the holes () of the plate ().

Combustion device for a gas turbine

The invention relates to a method for operating a gas turbine which includes a compressor with annular inlet area, at least two burners, a combustion chamber and a turbine. According to the method, at least one first partial intake flow, consisting of oxygen-reduced gas which has an oxygen concentration which is lower than the average oxygen concentration of the compressor intake flow, and at least one second partial intake flow, consisting of fresh air, are fed to the compressor in an alternating manner in the circumferential direction of the inlet area. In addition, the invention relates to a gas turbine power plant with a gas turbine, the compressor inlet of which includes at least one first segment and at least one second segment which are arranged in an alternating manner around a compressor inlet in the circumferential direction, wherein a feed for an oxygen-reduced gas is connected to the first segment and a fresh air feed is connected to the second segment of the compressor inlet.

Gas turbine power plant with non-homogeneous input gas

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