Alstom Technology Limited

General Electric Company

Ansaldo Energia Switzerland Ag

Ansaldo Energia IP UK Limited

Jan Cerny

Adnan Eroglu

Naresh Aluri

Jaan Hellat

Madhavan Narasimhan Poyyapakkam

Douglas Pennell

Fernando Biagioli

Franklin Marie Genin

Ennio Pasqualotto

Anurag Jhalani

Douglas Anthony Pennell

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

An axial swirler for a gas turbine burner includes a vane ring with a plurality of swirler vanes circumferentially distributed around a swirler axis. Each of the swirler vanes includes a trailing edge. In order to achieve a controlled distribution of the exit flow velocity profile and/or the fuel equivalence ratio in the radial direction. The trailing edge is discontinuous with the trailing edge having a discontinuity at a predetermined radius.

Axial swirler for a gas turbine burner

A burner () for a combustion chamber of a turbogroup includes a swirl generator (), a mixer (), and a lance () for introducing pilot fuel into a combustion space (). In order to stabilize combustion, the lance () is designed and/or arranged so that, at least in the pilot mode of the burner (), it extends far enough into the burner interior () for a flame front () of a combustion reaction, which takes place in the combustion space (), to extend at least partially into the burner interior ().

Burner fuel lance configuration and method of use

A method for operating a combustion chamber () for a gas turbine. The combustion chamber () has a combustion space () and a plurality of burners (), which are each arranged at an inlet () of the combustion space (). The burners () are divided into two burner groups (). A common oxidant supply () is provided for all the burners (). A common first fuel supply () is provided for the burners () belonging to the first burner group (). A common second fuel supply () and a common additional fuel supply () are provided for the burners () belonging to the second burner group (). A first fuel stream () is fed to the burners () belonging to the first fuel group (), in such a manner that a lean fuel/oxidant mix is established. A second fuel stream () is fed to the burners () belonging to the second burner group () in such a manner that a lean fuel/oxidant mix is established. The first fuel stream () for each burner () belonging to the first burner group () is selected to be greater than the second fuel stream () for each burner () belonging to the second burner group (). In addition, an additional fuel stream () is fed to the burners () belonging to the second burner group ().

Method for operating a gas turbine combustion chamber including a plurality of burners arranged in groups

Growing community of inventors

Fahrwangen, Switzerland

Jan Cerny