Abstract: 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 (?) of the combustion at least of the second combustor is kept below a maximum air ratio (?max).

Abstract: The lance of a burner includes a body that defines a first duct with first nozzles for injecting a liquid fuel and a second duct with second nozzles for injecting a gaseous fuel. Outlets of the first nozzles are spaced apart from outlets of the second nozzles. The body includes a third duct with third and fourth nozzles for injecting air. The third nozzles surround an axis of the first nozzles and the fourth nozzles surround an axis of the second nozzles.

Abstract: An exemplary burner of a gas turbine includes a tubular body with an inlet for an entrance of an air flow, downstream of inlet vortex generators, and a lance projecting into the tubular body and having a terminal portion extending along a longitudinal axis of the burner which is provided with nozzle groups for injecting fuel into the tubular body. The nozzle groups can lay in an injection plane perpendicular to the axis of the terminal portion of the lance. Downstream of the lance, the burner has an outlet. A ratio x/L between an axial distance x between the side trailing edge of the vortex generator and the injection plane, and the length L of the tubular body can be less than approximately 0.1052.

Abstract: The combustion device includes a burner, a combustion chamber downstream of the burner, a lance projecting into the burner for fuel and air injection, and a plenum that at least partly houses the burner. The plenum is connected to the inside of the lance to supply an oxidiser to it.

Abstract: The burner (1) for a gas turbine includes a duct (2) housing a plurality of tetrahedron shaped vortex generators (3) and a lance (4) to inject a fuel to be combusted. Within the duct (2), a plurality of vortex generators (3) are provided with a plurality of holes (9) for injecting cooling air. The cooling holes (9) define passing through areas that are non-uniformly distributed on a top wall (11) of the vortex generators (3). A method for locally cooling a hot gases flow passing through a burner includes non-uniformly injecting cooling air from a vortex generator into the hot gas flow in the duct, which can reduce the occurrence of flashback in the burner.

Abstract: A burner arrangement is disclosed with a conical swirler in the form of a double cone which is arranged concentric to a burner axis and which encloses a swirl chamber, and with a central fuel lance which lies in the burner axis and projects from the cone point of the swirler into the swirl chamber, wherein a first stage is provided for injecting premix fuel, in which the premix fuel is injected radially outwards into the swirl chamber through injection openings which are arranged on the fuel lance, and wherein a second stage is provided for injecting premix fuel, in which the premix fuel is injected into an air flow, which is guided in the double cone, through injection openings in the double cone. With such a burner arrangement, the gas pressure which is required in the first stage is reduced by the entire premix fuel being injected into the swirl chamber in the first stage through two oppositely-disposed injection openings with increased opening diameter.

Abstract: In a method for the low-CO emissions part load operation of a gas turbine with sequential combustion, the air ratio (?) of the operative burners (9) of the second combustor (15) is kept below a maximum air ratio (?max) at part load In order to reduce the maximum air ratio (?), a series of modifications in the operating concept of the gas turbine are carried out individually or in combination. One modification is an opening of the row of variable compressor inlet guide vanes (14) before engaging the second combustor (15). For engaging the second combustor, the row of variable compressor inlet guide vanes (14) is quickly closed and fuel is introduced in a synchronized manner into the burner (9) of the second combustor (15). A further modification is the deactivating of individual burners (9) at part load.

Abstract: A reheat combustor for a gas turbine engine includes a fuel/gas mixer for mixing fuel, air and combustion gases produced by a primary combustor and expanded through a high pressure turbine. Fuel injectors inject fuel into the mixer together with spent cooling air previously used for convectively cooling the reheat combustor. The fuel mixture is burnt in an annular reheat combustion chamber prior to expansion through low pressure turbine inlet guide vanes. The fuel/gas mixer and optionally the combustion chamber define cooling paths through which cooling air flows to convectively cool their walls. The fuel injectors are also convectively cooled by the cooling air after it has passed through the fuel/gas mixer cooling paths. The low pressure turbine inlet guide vanes may also define convective cooling paths in series with the combustion chamber cooling paths.

Abstract: A slurry formulation for production of a patch or a strip of a thermal barrier coating includes 20-60 dry weight percent of alkali metal silicate binder, 40-80 dry weight percent of ceramic fillers including at least one ceramic filler selected from the group consisting of yttria-stabilized zirconia and magnesia, 0-20 dry weight percent of additives, wherein the dry weight percent of the alkali metal silicate binder, the ceramic fillers and the additives add up to 100%, and one of a solvent and suspension agent.

Abstract: 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.

Abstract: The burner of a gas turbine includes a swirl generator and, downstream of it, a mixing tube. The swirl generator is defined by at least two walls facing one another to define a conical swirl chamber and is provided with nozzles arranged to inject a fuel and apertures arranged to feed an oxidiser into the swirl chamber. The burner includes a lance which extends along a longitudinal axis of the swirl generator and is provided with side nozzles for ejecting a fuel within the burner. The side nozzles have their axes inclined with respect to the axis of the lance and can be positioned along the axis of the burner.

Abstract: The invention discloses a method for operating a gas turbine with sequential combustion, which gas turbine includes a compressor, a first combustor with a first combustion chamber and first burners, which receives compressed air from the compressor, a second combustor with a second combustion chamber and second burners, which receives hot gas from the first combustor with a predetermined second combustor inlet temperature, and a turbine, which receives hot gas from the second combustor. The CO emission for part-load operation is reduced by reducing the second combustor inlet temperature for base-load operation of the gas turbine, and increasing the second combustor inlet temperature when decreasing the gas turbine load (RLGT) from base-load to part-load.

Abstract: A burner, such as for a secondary combustion chamber of a gas turbine with sequential combustion having first and second combustion chambers, includes an injection device for introducing at least one gaseous fuel into the burner. The injection device has at least one body which is arranged in the burner with at least one nozzle for introducing the 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 perpendicularly or at an inclination to a main flow direction prevailing in the burner. The at least one nozzle has its outlet orifice at or in a trailing edge of the streamlined body. The body has two lateral surfaces substantially parallel to the main flow direction. At least one vortex generator is located on at least one lateral surface upstream of the at least one nozzle.

Abstract: The invention relates a method for mixing a dilution air with a hot main flow in a sequential combustion system of a gas turbine, wherein the gas turbine essentially comprises at least one compressor, a first combustor which is connected downstream to the compressor The hot gases of the first combustor are admitted to at least one intermediate turbine or directly or indirectly to at least one second combustor, wherein the hot gases of the second combustor are admitted to a further turbine or directly or indirectly to an energy recovery.

Abstract: 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.

Abstract: An exemplary combustor includes at least a portion having an inner liner and an outer cover plate, which together form an interposed cooling chamber. A plurality of hollow elements extend from the liner and protrude into the cooling chamber. Each hollow element defines a damping volume connected to an inner volume of the combustion chamber via a calibrated duct. During operation, the hollow elements damp pressure pulsations and, in addition, also transfer heat.

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: A method is provided for fuel injection in a sequential combustion system comprising a first combustion chamber and downstream thereof a second combustion chamber, in between which at least one vortex generator is located, as well as a premixing chamber having a longitudinal axis downstream of the vortex generator, and a fuel lance having a vertical portion and a horizontal portion, being located within said premixing chamber. The fuel injected is an MBtu-fuel. In said premixing chamber the fuel and a gas contained in an oxidizing stream coming from the first combustion chamber are premixed to a combustible mixture. The fuel is injected in such a way that the residence time of the fuel in the premixing chamber is reduced in comparison with a radial injection of the fuel from the horizontal portion of the fuel lance.

Abstract: The burner (1) of a gas turbine includes a swirl generator (2) and, downstream of it, a mixing tube (3). The swirl generator (2) is defined by at least two walls facing one another to define a conical swirl chamber (5) and is provided with nozzles (6) arranged to inject a fuel and apertures (7) arranged to feed an oxidizer into the swirl chamber (5). The burner (1) includes a lance (9) which extends along a longitudinal axis of the swirl generator (2) and is provided with side nozzles (11) for ejecting a fuel within the burner (1). The side nozzles (11) have their axes (12) inclined with respect to the axis of the lance (9) and can be positioned along the axis of the burner.

Abstract: A burner (1) for a gas turbine includes a duct (2) enclosing a plurality of vortex generators (3) and, downstream of them, a lance (5) provided with nozzles (6) for injecting a gaseous fuel. The burner (1) also includes a mixer (7) for diluting and mixing the gaseous fuel with air to form a mixture. The mixer (7) is connected to the nozzles (6) for feeding them with the mixture. A method includes feeding the gaseous fuel in the burner (1).