US20140053569A1

US20140053569A1 - Method for mixing a dilution air in a sequential combustion system of a gas turbine

Info

Publication number: US20140053569A1
Application number: US13/974,523
Authority: US
Inventors: Mirko Ruben Bothien
Original assignee: Alstom Technology AG
Current assignee: Ansaldo Energia Switzerland AG
Priority date: 2012-08-24
Filing date: 2013-08-23
Publication date: 2014-02-27
Also published as: EP2700879B1; CA2824124A1; SA113340804B1; KR20150132058A; RU2013138725A; EP2700879A2; CN103629660A; KR101712341B1; CA2824124C; JP6092047B2; EP2700879A3; JP2014044045A; RU2570480C2; KR20140026292A

Abstract

The invention concerns a method for mixing a dilution air with a hot main flow in 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, and 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. The hot gases of the second combustor are admitted to a further turbine or directly or indirectly to an energy recovery, wherein at least one combustor runs under a caloric combustion path having a can-architecture. At least one dilution air injection is introduced into the first combustor, and wherein the direction of the dilution air injection is directed against or in the direction of the original swirl flow inside of the first combustor.

Description

CROSS-REFERENCE TO RELATED APPLICATION

This application claims priority to European Application 12181748.0 filed Aug. 24, 2012, the contents of which are hereby incorporated in its entirety.

TECHNICAL FIELD

The invention refers to a method for mixing dilution air in a sequential combustion system of a gas turbine. The invention refers additionally to a dilution air injector for implementing the aforementioned method. Furthermore, the invention is related to mixing of dilution air with a hot main flow in a gas turbine or in a “CPSC” (Constant Pressure Sequential Combustion) for a can as well as annular combustor design in a reliable and uniform way.

BACKGROUND

Beforehand, some general considerations which allow a better understanding of the invention:
CO emissions of gas turbine engines need reductions for the sake of saving the environment. Such emissions are known to appear, when there is not sufficient time in the combustion chamber to ensure the CO to CO₂oxidation, and/or this oxidation is locally quenched due to contact with cold regions in the combustor. Since firing temperatures are smaller under part load conditions CO, and the CO to CO₂oxidation gets slower, thus CO emissions usually tend to increase under these conditions.
A reduction of CO emissions in turn might be invested in lowering the gas turbine load at the parking point of a gas turbine. This reduces the environmental impact due to reduced CO₂emissions and overall cost of electricity due to less fuel consumption during engine parking. Finally the CO emission reduction might be invested in a reduction of first costs due to savings on a CO catalyst. In this case a CO catalyst might be avoided (or at least reduced). At the same time losses, which appear due to a catalyst will be removed (or at least reduced), and thereby the overall efficiency of the power plant increased.
According to the US 2012/0017601 A1 the basic of this state of art is a method for operating the gas turbine, which keeps the air ratio λ of the operating burner of the second combustor below a maximum air ratio λ_maxduring part load operation. This method is characterized essentially by three new elements and also by supplementing measures which can be implemented individually or in combination.
The maximum air ratio λ_maxin this case depends upon the CO emission limits which are to be observed, upon the design of the burner and of the combustor, and also upon the operating conditions, that is to say especially the burner inlet temperature.
The first element is a change in the principle of operation of the row of variable compressor inlet guide vanes, which allows the second combustor to be put into operation only at higher part load. Starting from no-load operation, the row of variable compressor inlet guide vanes is already opened while only the first combustor is in operation. This allows loading up to a higher relative load before the second combustor has to be put in operation. If the row of variable compressor inlet guide vanes is opened and the hot gas temperature or turbine inlet temperature of the high-pressure turbine has reached a limit, the second combustor is supplied with fuel.
In addition, the row of variable compressor inlet guide vanes is quickly closed. Closing of the row of variable compressor inlet guide vanes at constant turbine inlet temperature TIT of the high-pressure turbine, without countermeasures, would lead to a significant reduction of the relative power.
In order to avoid this power reduction, the fuel mass flow, which is introduced into the second combustor, can be increased. The minimum load at which the second combustor is put into operation and the minimum fuel flow into the second combustor are therefore significantly increased.
As a result, the minimum hot gas temperature of the second combustor is also increased, which reduces the air ratio λ and therefore reduces the CO emissions.
The second element for reducing the air ratio λ is a change in the principle of operation by increasing the turbine exhaust temperature of the high-pressure turbine TAT1 and/or the turbine exhaust temperature of the low-pressure turbine TAT2 during part load operation. This increase allows opening of the row of variable compressor inlet guide vanes to be shifted to a higher load point.
Conventionally, the maximum turbine exhaust temperature of the second turbine is determined for the full load case and the gas turbine and possibly the downstream waste heat boiler are designed in accordance with this temperature. This leads to the maximum hot gas temperature of the second turbine not being limited by the TIT2 (turbine inlet temperature of the second turbine) during part load operation with the row of variable compressor inlet guide vanes closed, but by the TAT2 (turbine exhaust temperature of the second turbine). Since at part load with at least one row of variable compressor inlet guide vanes closed the mass flow and therefore the pressure ratio across the turbine is reduced, the ratio of turbine inlet temperature to turbine exhaust temperature is also reduced.
Correspondingly, with constant TAT2 the TIT2 is also reduced and in most cases lies considerably below the full load value. A proposed slight increase of the TAT2 beyond the full load limit, typically within the order of magnitude of 10° C. to 30° C., admittedly leads to an increase of the TIT2, but this remains below the full load value and can practically be achieved without service life losses, or without significant service life losses. Adaptations in the design or in the choice of material do not become necessary or can be limited typically to the exhaust gas side. For increasing the TIT2, the hot gas temperature is increased, which is realized by an increase of the fuel mass flow and a reduction of the air ratio λ, which is associated therewith. The CO emissions are correspondingly reduced.
A further possibility for reducing the air ratio λ of the burner in operation is the deactivating of individual burners and redistribution of the fuel at constant TIT2.
In order to keep the TIT2 constant on average, the burner in operation has to be operated hotter in proportion to the number of deactivated burners. For this, the fuel feed is increased and therefore the local air ratio λ is reduced.
For an operation which is optimized for CO emissions, in a gas turbine with split line, a burner (for example for the second combustor) which is adjacent to the split line is typically deactivated first of all. In this case, the plane in which a casing is typically split into upper and lower halves is referred to as the split line. The respective casing halves are connected in the split line by a flange, for example.
Its adjacent burners are subsequently then deactivated or a burner, which is adjacent to the parting plane on the opposite side of the combustor is deactivated and in alternating sequence the adjacent burners, which alternate on the two sides of the combustor, starting from the parting plane, are deactivated.
A burner which is adjacent to the split line is preferably deactivated first of all since the split line of a gas turbine is typically not absolutely leak proof and in most cases a leakage flow leads to a slight cooling and dilution (see below mentioned considerations) of the flammable gases and therefore to locally increased CO emissions. As a result of deactivating the burners which are adjacent to the parting plane, these local CO emissions are avoided.
The combustion instabilities, which are to be avoided by means of staging, typically no longer occur at low load, or are negligibly small, or at part load combustion occur. In one exemplary embodiment, it is proposed, therefore, to carry out the restricting not by means of a fixed restrictor but by means of at least one control valve. This at least one control valve is opened at low load so that all the activated burners can be operated virtually homogenously with a low air ratio λ. At high load, the at least one control valve is throttled in order to realize the staging.
Referring to the aforementioned aspects for an optimized operation for CO emissions and in connection with the currently proceeding, cooling air from the reheat combustor and any remaining air from the premix combustor, or fresh air from plenum can be supplied as dilution air to the main hot gas flow.
Existing solutions to solve this problems consists in an injection of secondary medium without swirl. Additionally burners generating swirl in opposite directions to minimize swirl of main flow.
Accordingly, the technical problem consists in a rapid and good mixing of hot gas products with fresh dilution air to obtain uniform inlet temperatures and flow field upstream of a reheat burner. Control of swirl of main flow is, additionally, mandatory.

SUMMARY

The present invention is based on the object of proposing a method to improve mixing of dilution air and hot combustion products of first stage combustor by injecting dilution air with a swirl.
In addition, control of the existing main swirl is possible by either injecting in direction of the main swirl flow, to amplify existing swirl flow, or against it, to suppress integral or at various stages or levels the existing swirl flow.
Additionally, the present invention is based as example on the concept of constant pressure sequential combustion system. In this concept, hot combustion products from premix combustor are cooled down by a dilution air introduction and subsequently enter a reheat combustor.
Dilution air is responsible for mixing of premix and reheat cooling air with hot combustion products from the premix combustor. Primary requirements from such a dilution air are uniform temperature distribution at the inlet to reheat burner, as well as low pressure drop for performance reasons.
Accordingly, in the above identified gas turbine combustors, the main flow usually exhibits a swirling flow pattern. In an annular combustion chamber, this can be due that all burners generate a swirling flow in the same direction. In can combustors, usually more than one burner nozzle is used to inject the fuel and air into the combustion chamber. This also can result in a main swirl of the mean flow.
If downstream of this swirling flow air or fuel, and also dilution air, is injected, the challenge consists to obtain a good mixture with the hot gases as fast as possible. This is crucial to achieve uniform temperature and flow profile at the inlet of the reheat burner.
The present invention is in this sense related to mixing of dilution air with a hot main flow in a constant pressure sequential combustion system for a can as well as annular combustor designs in a reliable and uniform way.
In details, the invention describes below a procedure for mixing a dilution air with hot combustion products inside of first combustor, additionally, by injecting dilution air with a swirl, furthermore, control of the existing main swirl flow by either injecting in direction of the main swirl flow and finally amplify or suppress at various stages the existing main swirl flow.
Generic sketches of such gas turbines are shown in FIGS. 1 to 3, especially in FIG. 1.
Therein a compressor is followed by a combustor section, which can consist of a number of cans. Within these cans a first combustor is followed by a second combustor. Between or intermediate these two combustors dilution air might be injected in order to control the inlet temperature of the second combustor and therefore the self-ignition time of the fuel injected therein. Finally the hot combustion gases are fed into a turbine.
A can-architecture is also given, when an annular first and/or second combustion chamber having or comprising to each burner in flow direction an independent can or a separating flow combustion area which is wall-isolated from each other of the adjacent combustion areas or burners.
The basic idea of current invention is based on two basically concepts:
1. The gas turbine is equipped with two combustors in series with an injection of dilution air against direction of main swirl flow.
2. The gas turbine is equipped with two combustors in series with an injection of dilution air in direction of main swirl flow.
By injecting the dilution air with a defined swirl the following objectives can be achieved:
1. Enhance mixing between dilution air and hot gases from first burner.
2. Suppress integral or in part swirl of main flow by injecting the dilution air against the main swirl flow direction.
3. Amplify swirl of main flow by injecting the dilution air in direction of the main swirl flow direction.
Advantages associated with the present invention are as follows:

- Better and faster mixing of dilution air and hot gas from first combustor.
- Control of swirl of main flow in connection with inlet profile of the reheat combustor.
- Possible application for all gas turbine concepts including feature swirling main flows and injection of air downstream of the first burner.

To ensure this final purpose it is also necessary that the geometries and/or flow coefficients of the various components of the gas turbine are measured and components with high flow rates and components with low flow rates are combined inside the combustor cans or annular combustion chamber.
The gas turbine comprises essentially 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, for example to a steam generator.
Further advantages associated with the present invention are as follows:

- Reduced total combustor pressure drop, thus increased thermodynamic efficiency.
- Simple design of the injection of dilution air.
- Uniform temperature distribution at reheat burner inlet, thus a homogenous combustion process can act on the pulsations in the combustor and can act on an over-proportional increase of CO production of the reheat burner.
- Reliable operation without local backflow or overheating.

Based on these findings the concept can be expected to work for an engine, which runs under sequential combustion (with or without a high pressure turbine) in a can-architecture, but not only.
Referring to a sequential combustion the combination of combustors can be disposed as follows:

- At least one combustor is configured as a can-architecture, with at least one operating turbine.
- Both, the first and second combustors are configured as sequential can-can architecture, with at least one operating turbine.
- The first combustor is configured as an annular combustion chamber and the second combustor is built-on as a can configuration, with at least one operating turbine.
- The first combustor is configured as a can-architecture and the second combustor is configured as an annular combustion chamber, with at least one operating turbine.
- Both, the first and second combustor are configured as annular combustion chambers, with at least one operating turbine.
- Both, the first and second combustor are configured as annular combustion chambers, with an intermediate operating turbine.

Accordingly, in terms of injection of dilution air with a swirl flow for a can-architecture the interaction between individual cans is minimal or inexistent. Therefore for a can variant the described concept will be even more effective than for annular engine architecture.
In addition to the method, a gas turbine for implementing the method is a subject of the invention. Depending upon the concept of the injection of dilution air, the design of the gas turbine has to be adapted and/or the fuel distribution system and/or the cooling air system have to be adapted in order to ensure the feasibility depending on the used dilution air for reducing the locally combustor pressure drop. All the components of a gas turbine lie within the range of permissible tolerances. These tolerances lead to slightly different geometries and characteristics for each component and for the used concept of injection of dilution air.
This, especially, also leads to different pressure losses and flow rates during operation. The tolerances are selected so that they have practically no influence upon the operating behavior during normal operation, especially at high part load and full load. For this, the geometries and/or flow coefficients of the various injection of dilution air are measured with existing flow rates in connection with the operating dilution air swirls.
Additional advantages associated with this invention are as follows:
CO emissions are reduced especially at lower part-load conditions. Therefore, the gas turbine can be parked at lower values during such a period.

- Thereby the power plant operator can save fuel and therefore reduce the overall cost of electricity.
- Environmental benefit due to reduced CO emissions, lower parking point (thus less fuel consumption and CO₂production) or a combination of both advantages.
- Possibility of eliminating an expensive CO catalyst. Therefore first costs are reduced.

When using a setup including dilution air swirl between subsequent operating combustors further advantages arise:

- Further CO reduction, with all advantages described above, due to increased volume for CO oxidation with origin in the first combustor.
- Reduction of circumferential temperature gradients between the different can combustors. Therefore the turbine inlet profile is improved and lifetime of turbine parts is improved.

BRIEF DESCRIPTION OF THE DRAWINGS

The invention is shown schematically in FIGS. 1 to 3 (1, 2, 2 a, 2 b, 3, 3 a, 3 b) based on exemplary embodiments.

In the drawings:

FIG. 1 shows a gas turbine equipped with two combustors in series forming a sequential combustion;

FIG. 1 a shows a section of a can combustor with respect to FIG. 1;

FIG. 1 b shows a section of an annular combustor with respect to FIG. 1;

FIG. 2 shows a gas turbine equipped with two combustors in series and dilution air injection;

FIG. 2 a shows a section of a can combustor with respect to FIG. 2;

FIG. 2 b shows a section of an annular combustor with respect to FIG. 2;

FIG. 3 shows a gas turbine equipped with two combustors in series and dilution air injection;

FIG. 3 a shows a section of a can combustor with respect to FIG. 3;

FIG. 3 b shows a section of an annular combustor with respect to FIG. 3.

DETAILED DESCRIPTION

FIGS. 1, 2 and 3 show a part of the gas- turbine group 100, 200, 300, namely the part which includes the sequential combustion, referring to a “CPSC” system (Constant Pressure Sequential Combustion).
Compressed air flows out of a compressor system (not shown) into a premixing burner 101, which can be operated with a fuel. The initial generation of hot gases takes place in a first combustion chamber 102 designed as a can combustor (see FIGS. 1 a, 2 a, 3 a) or as an annular combustion chamber (see FIGS. 1 b, 2 b, 3 b). The following generation of hot gases then take place in a second combustion chamber 104 designed as a can combustor (see FIGS. 1 a, 2 a, 3 a) or as an annular combustion chamber (see FIGS. 1 b, 2 b, 3 b). Typically, the gas turbine system includes a generator (not shown), which at the cold end of the gas turbine, that is to say at the compressor, is coupled to a shaft of the gas turbine.
Accordingly, FIGS. 1, 2 and 3 show gas turbine systems with sequential combustion for implementing the method according to the invention. The gas turbine system comprises a compressor (not shown), a first combustor 102, a second combustor 104 with a reheat burner and downstream of the second combustor a turbine (106, not shown).
FIG. 1 shows a first combustion chamber 102 having a premix burner 101 as disclosed for example by EP 0321 809 A1 or EP 0 704 657 A1. This publication forms an integral part of this description. The hot gases 103 generated in the first combustion chamber 102, designated as a can combustor or as an annular combustion chamber, stream to a second combustion chamber 104. The second combustion chamber 104 essentially has the form of a can (see FIG. 1 a) or an annular duct (see FIG. 1 b) through which flow occurs and in which preferably a gaseous fuel (not shown) is injected. A self-ignition of the injected fuel takes place starting at a temperature of the exhaust gases coming from the first combustion chamber 102 of at least 850 DEG C.
The second combustion chamber 104 has as burner 105, as discloses for example by EP 0 620 362 A1, a number of fuel lances roughly at the end of the premixing zone, which fuel lances are distributed over the periphery and assume the function of injecting the fuel. The entire configuration of the gas-turbine group, excluding the generator, is mounted on a single common rotor shaft.
The can architecture comprises a plurality of cans arranged in an annular array about the circumference of the turbine shaft (see FIG. 1 a), which enables an individual combustion operation of each can and which will be no harmful interactions among individual cans during the combustion process.
If premix burners 101 for the can's combustion or annular concept are provided, these should preferably be formed by the combustion process and objects according to the documents EP 0 321 809 A1 and/or EP 0 704 657 A2, wherein these documents forming integral parts of the present description.
In particular, said premix burners 101 can be operated with liquid and/or gaseous fuels of all kinds. Thus, it is readily possible to provide different fuels within the individual cans. This means also that a premix burner can also be operated simultaneously with different fuels.
The second or subsequent can combustor or annular combustor is preferably carried out by EP 0 620 362 A1 or DE 103 12 971 A1, wherein these documents forming integral parts of the present description.
Additionally, the following mentioned documents forming also integral parts of the present description:

- EP 0 321 809 A1 and B1 relating to a burner consisting of hollow part-cone bodies making up a complete body, having tangential air inlet slots and feed channels for gaseous and liquid fuels, wherein in that the centre axes of the hollow part-cone bodies have a cone angle increasing in the direction of flow and run in the longitudinal direction at a mutual offset. A fuel nozzle, which fuel injection is located in the middle of the connecting line of the mutually offset centre axes of the part-cone bodies, is placed at the burner head in the conical interior formed by the part-cone bodies.
- EP 0 704 657 A2 and B1, relating to a burner arrangement for a heat generator, substantially consisting of a swirl generator, substantially according to EP 0 321 809 A1 and B, for a combustion air flow and means for injection of fuel, as well of a mixing path provided downstream of said swirl generator, wherein said mixing path comprises transaction ducts extending within a first part of the path in the flow direction for transfer of a flow formed in said swirl generator into the cross-section of flow of said mixing path, that joins downstream of said transition ducts.

Furthermore, it is proposed fuel injector for use within a gas turbine reheat combustor, utilising auto-ignition of fuel, in order to improve the fuel air mixing for a given residence time. The specific embodiments of this injector are envisaged:

- The oscillating gaseous fuel is injected normal to the flow of oxidant in sense of a cross-flow configuration.
- The oscillating gaseous fuel is injected parallel to the flow of oxidant in sense of an in-line configuration.
- The oscillating gaseous fuel is injected at an oblique angle, between 0° and 90° to the flow of oxidant.
- EP 0 646 705 A1 and B1, relating to a method of establishing part load operation in a gas turbine group with a sequential combustion.
- EP 0 646 704 A1 and B1, relating to a method for controlling a gas turbine plant equipped with two combustor chambers.
- EP 0 718 470 A2 and B1, relating to method of operating a gas turbine group equipped with two combustor chambers, when providing a partial-load operation.

Other relevant published documents, which include one or more improvements of the above identified documents forming also integral parts of the present description.
Referring to a sequential combustion the combination of combustors can be disposed as follows:
At least one combustor is configured as a can-architecture, with at least one operating turbine.

- Both, the first and second combustors are configured as sequential can-can architecture, with at least one operating turbine.
- The first combustor is configured as an annular combustion chamber and the second combustor is built-on as a can configuration, with at least one operating turbine.
- The first combustor is configured as a can-architecture and the second combustor is configured as an annular combustion chamber, with at least one operating turbine.
- Both, the first and second combustor are configured as annular combustion chambers, with at least one operating turbine.
- Both, the first and second combustor are configured as annular combustion chambers, with an intermediate operating turbine.

In both cases, relating to can combustor 120 or annular combustion chamber 130, the azimuthal main flow 121, 131 is unitary in each system.
FIG. 2 shows a gas turbine, according to FIG. 1, having a first combustion chamber equipped with at least one dilution air injection 201 at appropriate place, downstream of the first burner system 101 and upstream of the second burner system 105, and having a second combustion chamber 104 downstream of the second burner system 105. More dilution air injections at different places along the first combustion chamber 102 are possible. Furthermore, the direction and the intensity of the single injected air along the first combustion chamber 102 can be regulated.
FIG. 2 a shows a can combustor 220 having tangential air inlet slots 222 forming a swirl flow 223 directed against the predominant direction of the original main swirl flow 221 from the operation of the first burner 101. The result of this impact consists in the fact that the existing swirl flow intensity from the first burner can be reduced or completely suppressed, depending on the intensity of the selected dilution air injection 222. FIG. 2 a shows a reduced resulting main swirl flow 224.
FIG. 2 b shows an annular combustion chamber 230 having tangential air inlet slots 232 forming a swirl flow 233 directed against the predominant direction of the original swirl flow 231 from the operation of the first burner 101. The result of this impact consists in the fact that the existing swirl flow intensity from the first burner can be reduced or completely suppressed, depending on the intensity of the selected dilution air injection 232. FIG. 2 b shows a reduced resulting main swirl flow 234.
FIG. 3 shows a gas turbine, according to FIG. 2, having a first combustion chamber equipped with a dilution air injection 301 at appropriate place, downstream of the first burner system 101 and upstream of the second burner system 105, and having a second combustion chamber 104 downstream of the second burner system 105. More dilution air injections at different places along the first combustion chamber 102 are possible. Furthermore, the direction and the intensity of the single injected air along the first combustion chamber 102 can be regulated.
FIG. 3 a shows a can combustor 320 having tangential air inlet slots 322 forming a swirl flow 323 in direction of the original main swirl flow 321 from the operation of the first burner 101. The result of this feeding consists in the fact that the existing swirl flow intensity from the first burner can be amplified, depending on the intensity of the selected dilution air injection 322. FIG. 2 a shows an amplified resulting main swirl flow 324.
FIG. 3 b shows an annular combustion chamber 330 having tangential air inlet slots 332 forming a swirl flow 333 directed against the predominant direction of the original swirl flow 331 from the operation of the first burner 101. The result of this impact consists in the fact that the existing swirl flow intensity from the first burner can be amplified, depending on the intensity of the selected dilution air injection 332. FIG. 3 b shows an amplified resulting main swirl flow 334.

Claims

1. 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, and 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, wherein at least one combustor runs under a caloric combustion path having a can architecture, and wherein at least one dilution air injection is introduced into the first combustor, and wherein the resulting swirl flow through the dilution air injection is directed against or in the direction of the original swirl flow inside of the first combustor.

2. The method as claimed in claim 1, wherein the first and second combustor run under a caloric combustion path having a can-architecture.

3. The method as claimed in claim 1, wherein the first combustor runs under a caloric combustion path having an annular architecture, and the second combustor runs under a caloric combustion path having a can-architecture.

4. The method as claimed in claim 1, wherein the first combustor runs under a caloric combustion path having a can-architecture, and the second combustor runs under a caloric combustion path having an annular architecture.

5. The method as claimed in claim 1, wherein that at least one combustor runs under a caloric combustion path having an annular architecture.

6. A dilution air injector for implementing 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, and 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, wherein at least one combustor runs under a caloric combustion path having a can architecture, and wherein the first combustor comprising tangential air inlet slots forming a swirl flow directed against or in direction of the original main swirl flow inside of the first combustor.

7. The dilution air injector as claimed in claim 6, wherein the first combustor runs under a caloric combustion path having an annular architecture, and the second combustor runs under a caloric combustion path having a can-architecture.

8. The dilution air injector as claimed in claim 6, wherein the first combustor runs under a caloric combustion path having a can-architecture, and the second combustor runs under a caloric combustion path having an annular architecture.

9. The dilution air injector as claimed in claim 6, wherein that at least one combustor runs under a caloric combustion path having an annular architecture.

10. The dilution air injector as claimed claim 6 wherein the first combustor comprising at least one injector, wherein the direction and/or intensity of the injected air along the first combustion chamber are subject to regulation.

11. The dilution air injector as claimed in claim 6 wherein that the injector comprising means for regulating the intensity of the selected dilution air injection or for an additional supporting dilution air.

12. A combustor as claimed in claim 6 wherein at least one combustor comprising a burner consisting of hollow part-cone bodies making up a complete body, having tangential air inlet slots and feed channels for gaseous and liquid fuels, wherein in that the centre axes of the hollow part-cone bodies have a cone angle increasing in the direction of flow and run in the longitudinal direction at a mutual offset, wherein a fuel nozzle, which fuel injection is located in the middle of the connecting line of the mutually offset centre axes of the part-cone bodies, is placed at the burner head in the conical interior formed by the part-cone bodies.

13. A combustor as claimed in one of claims 6 to 11, characterized in that at least one combustor comprising a burner for a combustion air flow and means for injection of fuel, substantially consisting of a swirl generator, which substantially consisting of hollow part-cone bodies making up a complete body, having tangential air inlet slots and feed channels for gaseous and liquid fuels, wherein in that the centre axes of the hollow part-cone bodies have a cone angle increasing in the direction of flow and run in the longitudinal direction at a mutual offset, wherein a fuel nozzle, which fuel injection is located in the middle of the connecting line of the mutually offset centre axes of the part-cone bodies, is placed at the burner head in the conical interior formed by the part-cone bodies, and as well of a mixing path provided downstream of said swirl generator, wherein said mixing path comprises transaction ducts extending within a first part of the path in the flow direction for transfer of a flow formed in said swirl generator into the cross-section of flow of said mixing path, that joins downstream of said transition ducts.