Patent Grant
10352568
The present invention relates to a combustor arrangement for a gas turbine assembly, comprising a first burner, a first combustion chamber, a mixer for admixing a dilution gas to the hot gases leaving the first combustion chamber during operation, a second burner, and a second combustion chamber arranged sequentially in a fluid flow connection, wherein the first burner, the first combustion chamber, the mixer for admixing the dilution gas before the second burner and the second combustion chamber are arranged in a row to form a flow path extending between the first combustion chamber and the second burner.
Gas turbine assemblies are known from a number of prior art documents. WO 03/038253 provides a combustor arrangement for a gas turbine with sequential combustion via a plurality of common uniform annular combustion chambers.
WO 2012/136.787 A1 describes the use of a number of combustion chamber elements which are arranged individually around the rotor of a gas turbine assembly. Each combustion chamber element providing a combustor housing comprising a first and a second burner as well as an intermediate air supply has a tubular or quasi-tubular or shape-changing cross section and each combustion chamber element extends at a radial distance from a central axis of the gas turbine assembly. Fuel supply for the second burner as well as said air supply for the transfer duct are provided with specific ducts being radially oriented to the tubular combustion chamber element.
Based on this prior art, it is an object of the invention to provide an improved distribution of fuel and air for the two-staged combustor. A further object of the invention is an improved service and replacement approach for the combustor arrangement for a gas turbine assembly.
A combustor arrangement for a gas turbine assembly according to the invention comprises a central lance body arranged inside the flow path and extending from the first burner through the first combustion chamber into the mixer and optionally into the second burner, wherein the central lance body comprises at least one air duct, preferably a plurality of internal air ducts in the lance body, for providing air for at least one mixer between the first burner and the second burner, and optionally beyond, wherein the air is injected into the combustor through air supply elements.
The air ducts in the lance body can be arranged concentrically and they are annular-shaped passages when looked at in a cross section along the longitudinal axis of the combustor arrangement. Optionally they are arranged adjacent one to the other to the outer housing surface of the lance body trunk to provide inter alia thermal isolation to further elements inside the lance body as fuel ducts.
The annular-shaped passages have different lengths to connect with at least two different air supply elements along the lance body trunk. Especially one of the passages leads to the mixer and respective air supple elements and a further passage, usually of an inner annular air duct, leads to a zone beyond the mixer as well as to the free end of the lance body. Optionally further ducts lead to a mixing stage beyond the second burner.
The air supply elements at all stages can comprise annular passages, holes, slits or vents in the housing wall of the lance body or elements extending beyond the housing surface of the trunk of the lance body.
The housing of the combustor arrangement partially encompasses the lance body and is adapted to be connected to a housing of the second burner reaction zone of the turbine, wherein, in the connected position, the free end of the lance body extends into the housing of the second burner reaction zone. Then the central lance body is surrounded by the flow path and is arranged inside a combustor housing. Air supply channels as cavities between the different housing parts or as double walled housings are provided in counter flow direction of the flow path of the hot gases from the burners, i.e. outside of the combustor chamber towards a cavity around the first burner devices of the first burner to be introduced into the burner devices and/or into the air ducts of the lance body as well as optionally into the cavities of the double walled housing wall between the inner flow path of hot gases and the counter-flow air supply.
Within an embodiment of the combustor arrangement the fuel ducts are double line ducts adapted within the lance body to transport a first liquid fuel product and a second gaseous fuel product to the burners.
Each second burner can comprise fuel supply elements extending into the combustion cavity of the associated combustor housing. Such fuel supply elements are then connected with the fuel ducts and they can for example be lobed or micro VG injectors. The fuel supply elements can extend from the trunk of the lance body. They can extend radially from the trunk.
Each first burner can also comprise fuel supply elements extending into the combustion cavity of the associated combustion chamber element. Such fuel supply elements are then connected with the fuel ducts and they can for example be axial swirler injectors, flame sheet injectors, EV or AEV burners, wherein an EV burner is shown in EP 0 321 809 A1 and a so called AEV burner is shown in DE 195 47 913 A1.
The combustor housing can provide a cross-section increasing step of the combustion cavity between the first burner stage towards the first burner reaction zone for flame stabilization and to provide space for the expansion of the combustion gases.
The combustor housing can also provide a cross-section increasing step of the combustion cavity between the second burner stage towards a second burner reaction zone of the combustor arrangement for flame stabilization and to provide space for expansion of the combustion gases.
The combustor arrangement can comprise a plurality of first burners arranged around the central lance, e.g. between two and ten first burners.
The combustor housing partially encompasses the lance body and is adapted to be connected to a housing of the second burner reaction zone of the turbine, wherein, in the connected position, the free end of the lance body extends into the housing of the second burner reaction zone.
The combustor arrangement preferably has a removable central lance body. The central lance body is removably mounted in the combustor arrangement. The combustor arrangement can be designed to allow an axial removal of the central lance body along the longitudinal axis of the combustor arrangement. The cross section of the flow path increases in counter flow direction such that the lance body and fuel injectors extending from the trunk of the lance body can be retracted in axial direction out of the flow path. The first burner typically has a smaller cross section that the first combustion chamber but the lance body shall be retractable together with the first burner, respectively a part of the front plate of the first combustion chamber shall be removable, preferably together with the lance body, to allow an axial retraction of the lance body.
For example the outer diameter of the hot gas flow path inside the sequential combustor arrangement remains constant or increases in counter flow direction from the position of the second burner to the mixer, and further to the first combustion chamber. The first burner is arranged such that it be removed separately before a removal of the central lance body or such that it can be removed together with the central lance body. The central lance body can be removed or withdrawn in counter flow direction of the hot gases in the sequential combustor arrangement.
The central lance provided according to the invention comprises inherently the fuel injection lances mounted within the housing. The central lance can be retrieved from the frame of the gas turbine in one single piece and can be replaced and serviced as such. This is far more effective than the replacement of the single fuel injection lances of WO 2012/136.787.
A further advantage is achieved through the distribution of fuel and air through the central lance body for both stages of the burner. Another advantage of a further embodiment of the invention is the better mixing because air can be injected from the outside housing wall as well as from the lance itself.
The invention provides a combustor arrangement for a gas turbine assembly having a central lance with axial swirlers, thus building a lower-cost and robust so-called constant pressure sequential combustor, which has the main advantage that the central lance is retractable comprising the fuel supply for all stages. The fuel injection for the first burner stage can be additionally staged in the radial, circumferential and axial direction.
It proved to improve the function of the combustor that a sudden expansion, i.e. an sudden increase in cross section, in the form of a backwards facing step or shoulder on both inner and outer side of the annulus follows the annular section of the first stage. Together with the swirl from the first stage, this step stabilizes the flame in the first burner stage in a wide operating range. For low load conditions, fuel can be predominantly supplied to the inner zone with radially staged fuel supply. At higher loads, fuel can be increased to outer stage.
After the first burner reaction zone, a dilution air mixer can be used to reduce the temperature of the hot gases to the level required by the second burner stage. Dilution air mixer can be supplied with air from both outside and inside, forming a double-sided, opposed wall jet mixer. A central-body type reheat burner follows the dilution air mixer. Fuel supply for the second stage burner is provided completely through the central lance body, both for gaseous and liquid fuels.
Burner configuration for the first burner stage can be inter alia axial swirler/injectors or so-called EV or AEV burners as disclosed in EP 0 321 809 A1 for the EV-burner and DE 195 47 913 A1 for the AEV-burner.
Further embodiments of the invention are laid down in the dependent claims.
Preferred embodiments of the invention are described in the following with reference to the drawings, which are for the purpose of illustrating the present preferred embodiments of the invention and not for the purpose of limiting the same. In the drawings,
The combustor 10 of the gas turbine assembly of
Each first burner device of the first burner 20 is arranged downstream of the compressor (not shown) and is acted upon by the air compressed there. The second burner 60 is arranged downstream of the reaction zone 21 belonging to the associated first burner 20 and is provided in an annular region around the lance body 50. The first reaction zone 21 is also called first combustion chamber. Each first burner device of the first burner 20 has a first fuel supply device 22 which supplies a gaseous and/or liquid fuel to said first burner device via a first fuel supply element 23 (here a lance extending into the first burner 20) provided on the longitudinal axis 24 of each first burner device.
The second burner 60 has autonomous second fuel supply elements 63 which likewise ensure the supply of a gaseous and/or a liquid fuel as will be explained later.
The first fuel supply device 22 can be connected (not in
The combustor 10 of the gas turbine assembly comprises the combustor housing 100 encompassing the plurality of first burner devices. Housing 100 can be a multi-part housing and being mounted in a flange area 101 to an exterior frame 102. It is also possible that the housing encompasses the exterior frame 102 entirely. Housing part 90 is usually also integrated into the combustor housing 100.
The different first burner devices are mounted within corresponding openings 103 of the housing 100. Each first burner device comprises a first burner housing 25 extending into the first burner reaction zone 21 and comprising at its free end 26 beyond the first burner reaction zone 21 a blocking and sealing area, especially a hula seal, against the housing part 90 of the combustor arrangement 10.
The number of combustor chambers arranged in this way depends on the size of the gas turbine assembly and on the power output to be achieved. The combustor chamber as accommodated in the housing 100 of a gas turbine assembly 10 is at the same time surrounded by an envelope of air 105, via which the compressed air flows to the first burner 20. The number of first burner devices of the first burner stage 20 can be predetermined to be between e.g. 3 and 10.
The combustion gas path is symbolized here by an arrow 27 and through which the combustion gases of the first burner 20 flow when the combustor of the gas turbine assembly is in operation.
The compressor generates compressed air which is supplied to the first burners 20. A substream of the compressed air may in this case serve as cooling gas or cooling air and be utilized for cooling various components of the combustor 10 of the gas turbine assembly. Here it flows between the housing parts 25 and 100 and provide a thermal isolation between these surfaces. The first fuel supply element 23 injects the fuels directly into the individual first burner device of the first burner 20, said burner device being acted upon by compressed air and being designed as a premix burner. Fuel injection and the respective premix burner are in this case coordinated with one another such as to establish a lean fuel/oxidizer mixture which burns within the first burner reaction zone 21 with favorable values for pollutant emission and efficiency. It is especially noted that the cross-section of the first reaction zone 21 behind the burner device is larger than the cross-section after the first burner 20 and approaching the second burner 60 at the end of zone 21. The combustion gases in this case occurring are supplied to the second burner 60.
The combustion gases from the first reaction zone 21 are cooled to an extent such that fuel injection into the combustion gases, which takes place via the second fuel supply device 63 at the second burner 60, does not lead to undesirable premature auto-ignition outside the second reaction zone 40. For example, the combustion gases are cooled to about 1100° C. or below with the aid of the elongated first reaction zone acting as a heat exchanger.
The fuel for the second stage is supplied from the center of the lance body 50 where on the input side a curled duct 162 provides elasticity when the device changes its dimension due to change of temperature. The spiral duct 162 for an axial compensation of the fuel duct line is then provided as longitudinal duct 62 along the axis 13 inside the lance body 50 of the combustor 10 until the second burner zone. There, an L-shaped outlet provides the liquid into the second burner area 60 through a number of second fuel supply devices 63 to distribute the fuel.
This additional fuel is then supplied in the second burner 60 with the aid of the second fuel supply device 63 comprising injectors. The fuel is added to the combustion gases of the first stage cooled in this way, here, too, the burners and fuel supply being configured so as to form a lean fuel/oxidizer mixture which burns in the second reaction zone 40 with favorable values in terms of pollutant emission and of efficiency.
The combustion gases formed in the second reaction zone 40 are then leaving the combustor arrangement and are led to the turbine. In this context, the central lance body 50 comprises a rounded free end 51, especially an aerodynamically shaped free end. The five first burner devices form a common ring-shaped transfer duct, so that the turbine acting directly downstream can be acted upon uniformly. It is noted that as beyond the first stage 20, the second burner reaction zone 40 is provided with a cross-section enlarging step providing space for the expansion of the fuel-gas mix. The second burner reaction zone 21 is also called second combustion chamber.
As an optional feature the central lance 50 can also provide cooling and process air in an air injection stage, also called mixer 30 between the first burner 20 and the second burner 60. The cooling air is distributed via air supply elements 33. These air supply elements 33 can be provided on both wall parts of the combustor casing, at the inner wall and at the outer wall, i.e. at the cylindrical inner wall of the lance 50 housing and at the cylindrical outer wall of the housing parts 90. To achieve this air ducts are provided within the housing part 90 or the entire housing part 90 comprises an air guiding cavity 91. On the inner side air ducts 52 and 53 are provided within the lance body 50.
It is an advantage of feeding the air from the outer surface housing 90 and from the inner surface housing, especially in air injection stage 30, but also at the end of the lance body 50 with ducts 53 and opposite distribution vents in housing 90 in the lower second burner stage 61, that the air has only to travel half the diameter of the combustor in area 30 (or 61) to thoroughly mix with the combustion gas in the mixing stage 31 (or the mixing stage 61) when travelling to the second burner 60 or to the second burner reaction zone 40. The combustion process can be further enhanced, if short tubes are provided radially or slightly oriented in the direction of the gas flow as air supply elements 33 to inject the air even more evenly distributed within the process cavity between the stages 21 and 31.
It is an advantage of the principle of use of the single central lance body 50, incorporating a plurality of first burner devices, that it is independent from the embodiment chosen for the fuel injection lance with its first and second burners 20 and 60. Although a specific first burner stage 20 from the applicant (GT13E2 AEV Burner by Alstom) is schematically shown in the drawing of
On the other side, it is possible that gas turbine assembly 10 is run with only a part of the autonomously operated first burner devices of first burner 20 for part-load operation. Then, there is not necessarily a reduction in operation to the five first burners devices, but the number of first burner devices which are fully in operation can be reduced, here from five to a reduced number. Flexibility, the gain in efficiency and minimization of pollutant emissions in the gas turbine assembly 10 according to the invention can thus be maximized in any operating state.
The combustor arrangement 10 is shown with simplified main parts. The combustor arrangement has an encompassing housing 100 wherein the housing parts 90 of the embodiment of
Air from ducts within the central lance 50 and from the encompassing housing cavity 191 are injected at the mixing stage 30 according to the air flow 35 indicating arrows to be mixed within the mixing stage 31. The introduction of this additional air can be provided through simple bores, slits or vents in the housing walls as air supply elements 33.
Then additional fuel is injected at the second burner stage 60 as described in connection with the embodiment of
It can be seen from
In one embodiment, which can of course be combined with the features of the embodiment of
The annular passage 211 splits off into an burner area 213 around the different first burner devices and around the burner device housings 95 as well as into an device housing passage 215. The respective arrows are gas flow path arrow 212 and 214. The gas in the device housing passage 215 flows in a counter flow compared to main burn flow path 27.
Gas around the burner devices enters the burner devices at arrow 216 and are guided into the combustor reaction zone 21. A further gas flow 218 enters the lance body 50 and divides up in cavity space 219 inside the trunk of lance body 50 into an outer annular space 221 and an inner annular space 223. Both cavities guide gas inside the trunk to the respective outlets in the mixing stage 30 and the second burner stage 60.
Reference numeral 224 at the mixer 30 shows an injection arrow 224 directed radially to inject the gas as dilution gas into the mixer chamber. A further gas portion is guided along the lance body trunk 50 in an annular passage 225 towards the end of the mixing stage.
On the opposite housing 90 side, gas entering through the liner 41 in space 233 is guided through similar holes, vents or annular passages according to the referenced arrow 234 into the mixing stage. Further gas from the space 233 is guided according to arrow 266 as second burner gas into the second burner zone opposite to the fuel injection as explained in connection with
Inside the trunk of the lance body 50 at the rounded free end 51 similar gas from the annular passage 221 is injected into the lower zone 61 of the second burner through slits, holes or annular passages in the rounded free end 51 of the lance body 50 according to the arrow with the reference numeral 226.
Furthermore, it is possible that additional gas it injected into the second combustor area or zone 40 at the end surface 55 of the lance body 50 facing this second combustor area 40. The respective arrow has the reference numeral 228. The final gas passages 228 are oriented to inject the gas in an angle of 30 to 60 degrees from the longitudinal axis 13 of the combustor arrangement 10.
| Number | Date | Country | Kind |
|---|---|---|---|
| 14191329 | Oct 2014 | EP | regional |
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| 0 321 809 | Jun 1989 | EP |
| 2 116 766 | Nov 2009 | EP |
| 2 397 763 | Dec 2011 | EP |
| 2 400 216 | Dec 2011 | EP |
| 2 551 598 | Jan 2013 | EP |
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| WO 2012136787 | Oct 2012 | WO |
| Entry |
|---|
| Extended European Search Report dated Mar. 23, 2016, by the European Patent Office in corresponding European Patent Application No. 15188256.0-1605. (9 pages). |
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| Brochure, “GT13E2 AEV Burner”, An evolutionary gas turine solution, Alstom. |
| Office Action dated Nov. 6, 2017, by the European Patent Office in corresponding European Patent Application No. 15188256.0. |
| Number | Date | Country | |
|---|---|---|---|
| 20160123595 A1 | May 2016 | US |
