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 present invention relates to a burner for a combustion chamber of a gas turbine plant, with an injection device for introducing gaseous fuel into the burner. The injection device has a body which is arranged in the burner and which has at least one nozzle for introducing gaseous fuel into the burner. The body is configured as a streamlined body which has a streamlined cross-sectional profile and which extends with a longitudinal direction transversely with respect to a main flow direction prevailing in the burner. The at least one nozzle has its outlet orifice at an trailing edge of the streamlined body.

Abstract: In an SEV combustor and a method for reducing emissions in an SEV combustor of a sequential combustion gas turbine, an air/fuel mixture is combusted in a first burner and the hot gases are subsequently introduced into the SEV combustor (1) for further combustion. The SEV combustor (1) includes a chamber having a chamber wall (5) defining a mixing portion (8), for mixing the hot gases with a fuel, and a combustion region (9), at least one inlet (2) for introducing the hot gases into the mixing region (8), at least one inlet (12) for introducing a fuel into the mixing region (8) and at least one inlet (10, 13) for introducing steam into the mixing region.

Abstract: A combustion chamber arrangement is described for operating a gas turbine, with a combustion chamber wall which encloses the combustion chamber space and in the region of the combustion chamber outlet encloses a flow passage for hot gases which develop inside the combustion chamber, has a combustion chamber wall edge which freely terminates in the axial flow direction of the hot gases. The combustion chamber wall edge is formed with a profile which blocks or at least inhibits a diffuser effect when a cooling air flow, which is guided axially through the flow passages into the annular spatial area, flows over the combustion chamber wall edge.

Abstract: A combined flow straightener and mixer is disclosed as well as a burner for a combustion chamber of a gas turbine having such a mixing device. At least two streamlined bodies are arranged in a structure comprising the side walls of the mixer. The leading edge area of each streamlined body has a profile, which is oriented parallel to a main flow direction prevailing at the leading edge position, and with reference to a central plane of the streamlined bodies, the trailing edges are provided with at least two lobes in opposite transverse directions. The periodic deflections forming the lobes from two adjacent streamlined bodies are out of phase.

Abstract: A swirler including an annular housing with limiting walls. At least two vanes are arranged in the annular housing including the sidewalls of the swirler. The leading edge area of each vane has a profile, which is oriented parallel to a main flow direction prevailing at the leading edge position, wherein the profiles of the vanes turn from the main flow direction prevailing at the leading edge position to impose a swirl on the flow, and wherein, with reference to a central plane of the vanes the trailing edges are provided with at least two lobes in opposite transverse directions. A burner for a combustion chamber of a gas turbine including such a swirler and at least one nozzle having its outlet orifice at or in a trailing edge of the vane and to a method of operation of such a burner.

Abstract: A fuel lance (7) for introducing fuel into a gas flow in a combustor (1) of a gas turbine engine includes a region of the lance (7) through which the fuel is introduced into the gas flow having a generally helical formation (12).

Abstract: A combustor (1) for a gas turbine engine, particularly for a gas turbine having sequential combustion, includes a combustor wall (4) defining a mixing region (5) and a combustion region (6). The mixing region (5) has at least one first inlet (2) for introducing combustion air into the mixing region (5) and at least one second inlet for introducing fuel into the mixing region (5), the combustion region (6) extending downstream of the mixing region. The mixing region (5) crosses over to the combustion region (6) in a transition region (14). A baffle (9) extends from the transition region (14) generally in the downstream direction (15), forming at least one space (10) between the combustor wall (4) and the baffle (9).

Abstract: A method for assembling a gas turbine combustor includes providing a combustor case having a first end, a second end, and a centerline extending there between, coupling an end cover to the case first end, coupling a combustor liner within the case such that the liner is substantially coaxially aligned with respect to the case, and providing a streamline flow conditioner including a body that includes a radially outer surface and a radially inner surface, a deflection plate that extends from the body, and coupling the streamline flow conditioner to the case second end such that the streamline flow conditioner is coupled radially between the case and the combustor liner such that the deflection plate is adjacent the case second end. The deflection plate inner surface at least one of extends radially outward with respect to the centerline and defines a plurality of openings within the plate inner surface.

Abstract: A combustion chamber arrangement is described for operating a gas turbine, with a combustion chamber wall which encloses the combustion chamber space and in the region of the combustion chamber outlet encloses a flow passage for hot gases which develop inside the combustion chamber, has a combustion chamber wall edge which freely terminates in the axial flow direction of the hot gases. The combustion chamber wall edge is formed with a profile which blocks or at least inhibits a diffuser effect when a cooling air flow, which is guided axially through the flow passages into the annular spatial area, flows over the combustion chamber wall edge.

Abstract: A fuel lance (7) for introducing fuel into a gas flow in a combustor (1) of a gas turbine engine includes a region of the lance (7) through which the fuel is introduced into the gas flow having a generally helical formation (12).

Abstract: A combustor (1) for a gas turbine engine, particularly for a gas turbine having sequential combustion, includes a combustor wall (4) defining a mixing region (5) and a combustion region (6). The mixing region (5) has at least one first inlet (2) for introducing combustion air into the mixing region (5) and at least one second inlet for introducing fuel into the mixing region (5), the combustion region (6) extending downstream of the mixing region. The mixing region (5) crosses over to the combustion region (6) in a transition region (14). A baffle (9) extends from the transition region (14) generally in the downstream direction (15), forming at least one space (10) between the combustor wall (4) and the baffle (9).

Abstract: In an SEV combustor and a method for reducing emissions in an SEV combustor of a sequential combustion gas turbine, an air/fuel mixture is combusted in a first burner and the hot gases are subsequently introduced into the SEV combustor (1) for further combustion. The SEV combustor (1) includes a chamber having a chamber wall (5) defining a mixing portion (8), for mixing the hot gases with a fuel, and a combustion region (9), at least one inlet (2) for introducing the hot gases into the mixing region (8), at least one inlet (12) for introducing a fuel into the mixing region (8) and at least one inlet (10, 13) for introducing steam into the mixing region.

Abstract: A lance is presented for introducing fuel into a second burner of a combustion chamber of a gas turbine installation with sequential combustion, having a first and a second combustion chamber. The lance includes a foot and a shank which projects from it. In an installed state of the lance the foot extends perpendicularly or at an angle to a main flow direction of the burner. The shank extends centrally in the burner and generally parallel to the main flow direction. At least one projecting arm, having at least one nozzle for introducing fuel, is arranged such that it extends from the shank. A respective end of the at least one arm is oriented in the main flow direction and extends generally parallel to the main flow direction.

Abstract: The present invention relates to a burner for a combustion chamber of a gas turbine plant. The burner includes a lance for introducing gaseous fuel into the burner. A shaft of the lance has at least one nozzle for introducing gaseous fuel into the burner. A main injection direction of the respective nozzle is oriented onto a portion of a burner wall. An introduction device for a diverting fluid is provided, which is designed for introducing a diverting fluid counteracting an impingement of the fuel flow on the burner wall.

Abstract: The present invention relates to a burner for a combustion chamber of a gas turbine plant, with an injection device for introducing gaseous fuel into the burner. The injection device has a body which is arranged in the burner and which has at least one nozzle for introducing gaseous fuel into the burner. The body is configured as a streamlined body which has a streamlined cross-sectional profile and which extends with a longitudinal direction transversely with respect to a main flow direction prevailing in the burner. The at least one nozzle has its outlet orifice at an trailing edge of the streamlined body.

Abstract: A method for assembling a gas turbine combustor includes providing a combustor case having a first end, a second end, and a centerline extending there between, coupling an end cover to the case first end, coupling a combustor liner within the case such that the liner is substantially coaxially aligned with respect to the case, and providing a streamline flow conditioner including a body that includes a radially outer surface and a radially inner surface, a deflection plate that extends from the body, and coupling the streamline flow conditioner to the case second end such that the streamline flow conditioner is coupled radially between the case and the combustor liner such that the deflection plate is adjacent the case second end. The deflection plate inner surface at least one of extends radially outward with respect to the centerline and defines a plurality of openings within the plate inner surface.