Alstom Technology Limited

Ansaldo Energia Switzerland Ag

Abb Management Ag

Alstom (switzerland) Ltd

Abb Research Ltd.

General Electric Technology Gmbh

Ansaldo Energia IP UK Limited

Asea Brown Boveri Ag

Abb Alstrom Power (schweiz) Ag

Adnan Eroglu

Andrea Ciani

Peter Flohr

Urs Benz

Jaan Hellat

Ewald Freitag

Douglas Pennell

Michael Düsing

Johannes Buss

Richard Carroni

Yau-Pin Chyou

Klaus Knapp

Peter Stuber

Klaus Dobbeling

Madhavan Narasimhan Poyyapakkam

Michael Thomas Maurer

Wolfgang Polifke

Peter Senior

Christian Oliver Paschereit

Hans Peter Knopfel

Franz Joos

Felix Guethe

Peter Stuttaford

Timothy Albert Griffin

Rolf Althaus

Thomas Sattelmayer

Bruno Schuermans

Mirko Ruben Bothien

Bettina Paikert

Majed Toqan

Martin Zajadatz

Hans Peter Knoepfel

Robin McMillan

Diane Lauffer

Weiqun Geng

Jan Cerny

Oliver Riccius

Hans Peter Knöpfel

Michael Hutapea

Gregory John Kelsall

Andreas Schmid

Dieter Winkler

Eribert Benz

Jakob Keller

Stefano Bernero

Juergen Gerhard Hoffmann

Stephen Wilder Jorgensen

Roger Suter

Burkhard Schulte-Werning

Franklin Marie Genin

Robert Marmilic

Hartmut Haehnle

Thomas Zierer

Khawar Syed

Frank Graf

Jakob J Keller

Andreas Huber

Joshua R McNally

John Philip Wood

Naresh Aluri

Thiemo Meeuwissen

Verena Schmidt

Yang Yang

Mengbin Zhang

Alexander Beeck

Anton Winkler

Luis Tay Wo Chong Hilares

Nico Biagioli

Ennio Pasqualotto

Peter Novacek

Hans-Christoph Nienburg

Michael Duesing

Uwe Rüdel

Igor Baibuzenko

Rudolf Lachner

Klaus Döbbeling

Kaspar Loeffel

Nicolas Tran

Alexey Stytsenko

Sergey Mylnikov

Marcel Rieker

Pirmin Schiessel

Richard Smith

Hartmut Hähnle

Norbert Walter Emberger

Felix Güthe

Marcos Escudero Olano

Stephan Strueken

Uwe Ruedel

Afzal Pasha Mohammed

James Robertson

Thierry Lachaux

Jian-Xin Chen

Heinz Schaer

Douglas Anthony Pennel

Cristian Oliver Paschereit

Jens Policke

Zdenko Papa

Rainer Conzelmann

Luca Tentorio

A sealing arrangement at an interface between a combustor and a turbine of a gas turbine. The turbine can include deflecting vanes at its inlet, which deflecting vanes are each mounted within the turbine so as to define an inner or outer diameter platform and are in sealing engagement via an inner or outer diameter vane tooth with a seal arranged at the corresponding inner or outer diameter part of the outlet of the combustor. The seal is movable and is pressed on the inner or outer diameter vane tooth by a differential pressure such that the pressure of the mainstream hot gas flow is a lower pressure.

Sealing device arrangement at the interface between a combustor and a turbine of a gas turbine and gas turbine with such a sealing arrangement

The invention concerns a method for a part load CO reduction operation and a low-CO emissions operation of a gas turbine with sequential combustion. The gas turbine essentially includes at least one compressor, a first combustor which is connected downstream to the compressor. The hot gases of the first combustor are admitted at least to an intermediate turbine or directly or indirectly to a second combustor. The hot gases of the second combustor are admitted to a further turbine or directly or indirectly to an energy recovery. At least one combustor runs under a caloric combustion path having a can-architecture, and wherein the air ratio (&#x3bb;) of the combustion at least of the second combustor is kept below a maximum air ratio (&#x3bb;).

Method for a part load CO reduction operation for a sequential gas turbine

The invention relates to a gas turbine assembly which substantially includes at least one compressor, at least one first burner, at least one second burner that is connected downstream of the first burner, and at least one turbine that is connected downstream of the second burner. At least the first and second burner form a component of a tubular or quasi-tubular combustion chamber element in the flow direction of the combustion path of the burners. The combustion chamber element being closed or quasi-closed and extending between the compressor and the turbine. The combustion chamber elements are arranged around the rotor of the gas turbine assembly in the shape of a ring.

Gas turbine assembly and corresponding operating method

A dual fuel concentric nozzle such as for a dual fuel injector of a sequential burner of a sequential gas turbine, the dual fuel concentric nozzle having a nozzle downstream end, a liquid fuel duct, a gas fuel duct concentrically surrounding the liquid fuel duct and defining a gas fuel passage between the outer surface of the liquid fuel duct and the inner surface of the gas fuel duct, and a lateral surface concentrically surrounding the gas fuel duct and defining a carrier air passage, a downstream end edge of the gas fuel duct being recessed inside the nozzle downstream end with respect to a downstream end edge of the liquid fuel duct.

Dual fuel concentric nozzle for a gas turbine

The invention refers to a sequential combustor arrangement comprising a first burner, a first combustion chamber, a mixer for admixing a dilution gas via a dilution gas inlet 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 sequential combustor arrangement further includes four cooling zones with a cooling channel. During operation a cooling gas flows through the cooling channels. The disclosure further refers to a method for operating a gas turbine with such a sequential combustor arrangement.

Sequential combustion with dilution gas mixer

A fuel injector device includes a body having a leading edge and a trailing edge and defining a streamwise direction from the leading edge to the trailing edge. The fuel injector device body includes a first surface and a second surface opposite the first surface, each surface extending between and including the leading edge and the trailing edge, and the surfaces conjoining each other at the leading edge and the trailing edge. The trailing edge, when seen in the streamwise direction, undulates along a trailing edge mean line and, along its extent, deviates in opposite directions from the mean line and includes at least one inflection point along its extent.

Fuel injector device

A sequential combustor arrangement including a first burner, a first combustion chamber, a mixer for admixing dilution air to hot gases leaving the first combustion chamber during operation, a second burner, and a second combustion chamber arranged sequentially in a fluid flow direction. The mixer guides combustion gases in a hot gas flow path extending from the first combustion chamber. The second burner including a duct having an inlet for connection to the first combustion chamber and an outlet for connection to the second burner. The mixer includes injection pipes pointing inwards from the side walls for admixing the dilution air to cool the hot gases leaving the first combustion chamber. The injection pipes are distributed circumferentially along the side wall of the mixer. The injection pipes have a conical or quasi-conical shape addressed to the center of the mixer.

Dilution gas or air mixer for a combustor of a gas turbine

A combustor arrangement for a gas turbine includes a first burner, a first combustion chamber, a mixer for admixing a dilution gas to the gases leaving the first combustion chamber during operation, a second burner, and a second combustion chamber arranged sequentially in a fluid flow connection. These elements of the combustor arrangement are arranged in a row to form a flow path extending between the first combustion chamber and the second burner. The arrangement includes a central lance body in the flow path, extending from the first burner into the second burner, which lance body includes at least one air duct for providing air for the mixer, wherein the air is injected into the combustor through air supply elements.

Combustor arrangement for a gas turbine

A combustor arrangement for a gas turbine includes a first burner, a first combustion chamber, a mixer for admixing a dilution gas to the gases leaving the first combustion chamber during operation, a second burner, and a second combustion chamber arranged sequentially in a fluid flow connection. These elements of the combustor arrangement are arranged in a row to form a flow path extending between the first combustion chamber and the second burner. The combustor arrangement includes acentral lance body arranged inside the flow path and extending from the first burner through the first combustion chamber into the mixer and into the second burner, wherein the lance body includes a fuel duct for providing fuel for the first burner and/or for the second burner.

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 Helmholtz damper for a combustor of a gas turbine includes an enclosure defining a damping volume from which a neck portion extends and which has a flow path (F) for cooling and purging air with an inlet opening and an outlet opening to the enclosure. The outlet opening is formed in the neck portion. A seal is arranged at the neck portion adjacent to the outlet opening for cooling and purging air such that a cooling effect of the seal is provided.

Helmholtz damper for gas turbine with cooling air flow

A gas turbine combustor part of a gas turbine includes a wall, containing a plurality of near wall cooling channels extending essentially parallel to each other in a first direction within the wall in close vicinity to the hot side and being arranged in at least one row extending in a second direction. The near wall cooling channels are each provided at one end with an inlet for the supply of cooling air, and on the other end with an outlet for the discharge of cooling air. The inlets open into a common feeding channel for cooling air supply, and the outlets open into a common discharge channel for cooling air discharge. The feeding channel and the discharge channel extend in the second direction.

Gas turbine part comprising a near wall cooling arrangement

The invention concerns a gas turbine combustion system, including a gas turbine. The gas turbine includes at least one compressor, at least one combustion chamber for generating working gas, wherein the combustion chamber connected to receive compressed air from the compressor, at least one turbine connected to receive working gas from the combustion chamber. The combustion chamber consists of an individual can-combustor or comprising a number of can-combustors arranged in an annular can-architecture, wherein the can-combustor having at least one premixed burner. The ignition of the mixture starts at the premixed burner outlet and the flame is stabilized in the region of the premixed burner outlet by means of a backflow zone. The can-combustor comprising a number of premixed burners arranged uniformly or divided at least in two groups within the can-combustor.

Gas turbine combustion system

A combustion chamber of a gas turbine including first and second premixed fuel supply devices connected to a combustion device having first zones connected to the first premixed fuel supply devices and second zones connected to the second premixed fuel supply devices. The second fuel supply devices are shifted along a combustion device longitudinal axis with respect to the first fuel supply devices. The first zones are axially upstream of the second premixed fuel supply devices.

Combustion chamber and method for operating a combustion chamber

Growing community of inventors

Untersiggenthal, Switzerland

Adnan Eroglu