Abstract: The invention relates to a can-combustor for a can-annular combustor arrangement in a gas turbine. The can combustor includes an essentially cylindrical casing with an axially upstream front panel and an axially downstream outlet end. The can combustor further includes a number of premixed burners, extending in an upstream direction from said front panel and having a burner exit, supported by this front panel, for supplying a fuel/air mixture into a combustion zone inside the casing. Up to four premixed burners are attached to the front panel in a substantially annular array. Each burner has a conical swirl generator and a mixing tube to induce a swirl flow of said fuel/air mixture.

Abstract: A combustor arrangement with a front panel, a combustor liner, and a carrier structure element is provided for carrying the front panel and the combustor liner, wherein the combustor arrangement further includes a fastening system for connecting the front panel, the combustor liner, and the carrier structure element to one another. The fastening system includes at least one elastic connection element, the latter being fixedly connected to the carrier structure element and extending therefrom to the combustor liner and to the front panel. The elastic connection element is further fixedly connected to the combustor liner and/or the front panel such as to clamp the front panel, the combustor liner, and the carrier structure element to one another in a substantially fluid tight manner.

Abstract: A front panel for a combustor has a hot side and a cold side and at least one reception adapted for receiving a combustor part. The front panel has a double-wall design with a hot-side wall and a cold-side wall. The hot-side wall defines a hot-side downstream surface of the front panel. The cold-side wall defines a cold-side upstream surface of the front panel. The hot-side wall and the cold-side wall are axially spaced from one another, extend parallel to one another, and are connected to one another by an outer side wall.

Abstract: 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.

Abstract: The damper arrangement include two concentric hollow shapes, each having a wall, wherein the walls form an annular volume therebetween. The damper arrangement further includes one or more necks for connecting to a combustion chamber at corresponding one or more contact points. The one or more necks are connected to the annular volume.

Abstract: 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.

Abstract: A combustor arrangement with a front panel, a combustor liner, and a carrier structure element is provided for carrying the front panel and the combustor liner, wherein the combustor arrangement further includes a fastening system for connecting the front panel, the combustor liner, and the carrier structure element to one another. The fastening system includes at least one elastic connection element, the latter being fixedly connected to the carrier structure element and extending therefrom to the combustor liner and to the front panel. The elastic connection element is further fixedly connected to the combustor liner and/or the front panel such as to clamp the front panel, the combustor liner, and the carrier structure element to one another in a substantially fluid tight manner.

Abstract: A front panel for a combustor has a hot side and a cold side and at least one reception adapted for receiving a combustor part. The front panel has a double-wall design with a hot-side wall and a cold-side wall. The hot-side wall defines a hot-side downstream surface of the front panel. The cold-side wall defines a cold-side upstream surface of the front panel. The hot-side wall and the cold-side wall are axially spaced from one another, extend parallel to one another, and are connected to one another by an outer side wall.

Abstract: The damper arrangement include two concentric hollow shapes, each having a wall, wherein the walls form an annular volume therebetween. The damper arrangement further includes one or more necks for connecting to a combustion chamber at corresponding one or more contact points. The one or more necks are connected to the annular volume.

Abstract: The invention relates to a can-combustor for a can-annular combustor arrangement in a gas turbine. The can combustor includes an essentially cylindrical casing with an axially upstream front panel and an axially downstream outlet end. The can combustor further includes a number of premixed burners, extending in an upstream direction from said front panel and having a burner exit, supported by this front panel, for supplying a fuel/air mixture into a combustion zone inside the casing. Up to four premixed burners are attached to the front panel in a substantially annular array. Each burner has a conical swirl generator and a mixing tube to induce a swirl flow of said fuel/air mixture.

Abstract: 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.

Abstract: 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.