EP3486570A1 - Sekundäre brennkammerphase für eine sequentielle gasturbinenbrennkammer - Google Patents
Sekundäre brennkammerphase für eine sequentielle gasturbinenbrennkammer Download PDFInfo
- Publication number
- EP3486570A1 EP3486570A1 EP17201920.0A EP17201920A EP3486570A1 EP 3486570 A1 EP3486570 A1 EP 3486570A1 EP 17201920 A EP17201920 A EP 17201920A EP 3486570 A1 EP3486570 A1 EP 3486570A1
- Authority
- EP
- European Patent Office
- Prior art keywords
- combustor
- cross
- stage
- injection nozzles
- flow injection
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Granted
Links
- 238000004401 flow injection analysis Methods 0.000 claims abstract description 64
- 238000002485 combustion reaction Methods 0.000 claims abstract description 44
- 239000003381 stabilizer Substances 0.000 claims abstract description 42
- 230000007704 transition Effects 0.000 claims abstract description 20
- 239000007789 gas Substances 0.000 claims description 30
- 238000011144 upstream manufacturing Methods 0.000 claims description 13
- 239000000295 fuel oil Substances 0.000 claims description 12
- 239000002737 fuel gas Substances 0.000 claims description 8
- 230000008859 change Effects 0.000 claims description 7
- 230000003134 recirculating effect Effects 0.000 claims description 3
- 239000000446 fuel Substances 0.000 description 18
- 238000002156 mixing Methods 0.000 description 13
- 239000000203 mixture Substances 0.000 description 9
- 238000002347 injection Methods 0.000 description 8
- 239000007924 injection Substances 0.000 description 8
- 238000004519 manufacturing process Methods 0.000 description 5
- 238000012423 maintenance Methods 0.000 description 4
- 238000000034 method Methods 0.000 description 4
- 230000000712 assembly Effects 0.000 description 3
- 238000000429 assembly Methods 0.000 description 3
- 230000009286 beneficial effect Effects 0.000 description 3
- 230000008901 benefit Effects 0.000 description 3
- 238000009420 retrofitting Methods 0.000 description 3
- 238000010790 dilution Methods 0.000 description 2
- 239000012895 dilution Substances 0.000 description 2
- 238000000605 extraction Methods 0.000 description 2
- 230000010355 oscillation Effects 0.000 description 2
- 230000008569 process Effects 0.000 description 2
- 230000000087 stabilizing effect Effects 0.000 description 2
- XLYOFNOQVPJJNP-UHFFFAOYSA-N water Substances O XLYOFNOQVPJJNP-UHFFFAOYSA-N 0.000 description 2
- OKTJSMMVPCPJKN-UHFFFAOYSA-N Carbon Chemical compound [C] OKTJSMMVPCPJKN-UHFFFAOYSA-N 0.000 description 1
- 239000000654 additive Substances 0.000 description 1
- 230000000996 additive effect Effects 0.000 description 1
- 229910052799 carbon Inorganic materials 0.000 description 1
- 238000001816 cooling Methods 0.000 description 1
- 230000008878 coupling Effects 0.000 description 1
- 238000010168 coupling process Methods 0.000 description 1
- 238000005859 coupling reaction Methods 0.000 description 1
- 238000013016 damping Methods 0.000 description 1
- 230000003247 decreasing effect Effects 0.000 description 1
- 230000001934 delay Effects 0.000 description 1
- 238000010586 diagram Methods 0.000 description 1
- 239000003344 environmental pollutant Substances 0.000 description 1
- 239000012530 fluid Substances 0.000 description 1
- 239000007788 liquid Substances 0.000 description 1
- 230000000116 mitigating effect Effects 0.000 description 1
- 238000012986 modification Methods 0.000 description 1
- 230000004048 modification Effects 0.000 description 1
- 239000003921 oil Substances 0.000 description 1
- 230000003647 oxidation Effects 0.000 description 1
- 238000007254 oxidation reaction Methods 0.000 description 1
- 230000035515 penetration Effects 0.000 description 1
- 231100000719 pollutant Toxicity 0.000 description 1
- 238000010926 purge Methods 0.000 description 1
Images
Classifications
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
- F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
- F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
- F23R3/28—Continuous combustion chambers using liquid or gaseous fuel characterised by the fuel supply
- F23R3/36—Supply of different fuels
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
- F23C—METHODS OR APPARATUS FOR COMBUSTION USING FLUID FUEL OR SOLID FUEL SUSPENDED IN A CARRIER GAS OR AIR
- F23C99/00—Subject-matter not provided for in other groups of this subclass
- F23C99/001—Applying electric means or magnetism to combustion
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
- F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
- F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
- F23R3/02—Continuous combustion chambers using liquid or gaseous fuel characterised by the air-flow or gas-flow configuration
- F23R3/16—Continuous combustion chambers using liquid or gaseous fuel characterised by the air-flow or gas-flow configuration with devices inside the flame tube or the combustion chamber to influence the air or gas flow
- F23R3/18—Flame stabilising means, e.g. flame holders for after-burners of jet-propulsion plants
- F23R3/20—Flame stabilising means, e.g. flame holders for after-burners of jet-propulsion plants incorporating fuel injection means
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
- F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
- F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
- F23R3/28—Continuous combustion chambers using liquid or gaseous fuel characterised by the fuel supply
- F23R3/34—Feeding into different combustion zones
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
- F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
- F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
- F23R3/42—Continuous combustion chambers using liquid or gaseous fuel characterised by the arrangement or form of the flame tubes or combustion chambers
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
- F23C—METHODS OR APPARATUS FOR COMBUSTION USING FLUID FUEL OR SOLID FUEL SUSPENDED IN A CARRIER GAS OR AIR
- F23C2900/00—Special features of, or arrangements for combustion apparatus using fluid fuels or solid fuels suspended in air; Combustion processes therefor
- F23C2900/07021—Details of lances
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
- F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
- F23R2900/00—Special features of, or arrangements for continuous combustion chambers; Combustion processes therefor
- F23R2900/03341—Sequential combustion chambers or burners
Definitions
- the present invention relates to a second-stage combustor for a sequential combustor of a gas turbine..
- burner assemblies i.e. devices devoted to fuel injection and mixing
- burner assemblies can be retracted form their housings for the purpose of maintenance or retrofitting without the need to disassemble large part of the combustors.
- burner assemblies are axially retractable for simple extraction from can combustors. Exactly on account of their complex structure, however, known burner assemblies cannot be retracted.
- a combustor of a gas turbine comprising:
- the combustor combines an axial lance injector with extremely simple design and a flame stabilizer device that allows to anchor the flame at one or more desired locations.
- the front flame may be set at an axially downstream location at full load, to reduce post-flame residence time and production of NOx.
- the axial lance injector is essentially defined by an elongated streamlined body that extends from the first-stage combustor into the second-stage combustion chamber and exploits cross-flow injection for providing effective supply and mixing even when large fuel flowrates are required.
- Cross-flow injection nozzles may be provided on the surface of the elongated streamlined body and there is no need for special components, creating complex fluid-dynamic structures.
- the axial lance injector may be manufactured by conventional manufacturing processes and it is not necessary to resort to additive techniques, to the advantage of cost.
- the axial lance injector is easily retractable in the axial direction, thus simplifying maintenance and retrofitting operations.
- Axial locations of the first and second cross-flow injection nozzles can be selected to exploit different time delays from injection locations and the flame front. This can be used for the purpose of mitigating thermoacoustic oscillations and for providing short mixing paths for highly reactive fuels.
- the second cross-flow injection nozzles are axially displaced downstream of the first cross-flow injection nozzles.
- the first cross-flow injection nozzles are fluidly coupled to a fuel gas supply line and the second cross-flow injection nozzles are fluidly coupled to a liquid fuel supply line.
- Separate supply paths for different fuels may be provided. Advantages of separate oil injection may thus be exploited. In particular, less strict purging requirements are allowed. Also, separate supply valves can be used and mixing paths can be separately optimized for different fuels.
- At least one of the first cross-flow injection nozzles is axially displaced with respect to the other first cross-flow injection nozzles.
- the first cross-flow injection nozzles are staggered in the axial direction.
- At least one of the second cross-flow injection nozzles is axially displaced with respect to the other second cross-flow injection nozzles.
- the second cross-flow injection nozzles are staggered in the axial direction.
- the second cross-flow injection nozzles are oriented at such an angle that fuel oil is injected in the second-stage combustion chamber with a non-zero axial component of velocity.
- Inclination of the second cross-flow injection nozzles allows to reduce residence time of highly reactive fuel oil, because injected fuel leaves the nozzles with a non-zero axial component of velocity. In turn, lower residence time reduces the need for mixing additional water to fuel oil. In some cases, additional water may not be required at all, especially when inclination of the second cross-flow injection nozzles is supplemented by advanced flame location, as permitted by the flame stabilizer device.
- the combustor comprises vortex generators upstream of the second-stage combustion chamber.
- the vortex generators determine relatively slow mixing, i.e. longer distance and time are required. At full load, however, the flame stabilizer device may be used to set the flame location downstream. So, despite the large fuel flow rate at full load, efficient mixing is achieved. Early self-ignition of the air and fuel mixture is in any case avoided.
- the vortex generators may have quite simple shape (e.g. prismatic) and manufacturing thereof does not entail substantial problems in relation to both process complexity and cost.
- the flame stabilizer device comprises a full-load flame stabilizer at a downstream end of the elongated streamlined body.
- Stabilizing the full-load flame at a downstream location helps to enhance beneficial effects in terms of increased mixing distance and reduce post-flame residence time, especially at full-load. This is particularly beneficial for the purpose of maintaining low NOx emission at full-load.
- the flame stabilizer device comprises at least one partial-load flame stabilizer between the second cross-flow injection nozzles and the full-load flame stabilizer.
- Flame location may be adjusted during operation in accordance with load requirement. At partial load, flame temperature is low and long post-flame residence time is desired to achieve complete oxidation of carbon contained in the fuel flow and to reduce CO emissions. Thus, it may be of advantage to set the flame location at an upstream position. On the other hand, at full-load it is preferred to have the flame located as downstream as possible in the second-stage combustion chamber, to obtain good air-fuel mixing and reduce production of NOx.
- the flame stabilizer device comprises at least one of:
- Flame location can be thus effectively and precisely controlled during operation of the gas turbine.
- the first-stage combustor comprises an upstream end-cap and the elongated streamlined body is supported at the upstream end-cap.
- a flow channel is defined in the first-stage combustor and in the second-stage combustor around the elongated streamlined body and a cross section of the flow channel changes gradually along the flow direction in the transition region.
- the smooth transition between the first-stage combustor and the second-stage combustor prevents flow stagnation at the inlet of the second-stage combustion chamber. In this manner, stable flame anchorage is prevented and the flame location may be moved downstream as desired during operation using the flame stabilizer device.
- Figure 1 shows a simplified view of a gas turbine assembly, designated as whole with numeral 1.
- the gas turbine assembly 1 comprises a compressor section 2, a combustor assembly 3 and a turbine section 5.
- the compressor section 2 and the turbine section 3 extend along a main axis A.
- An airflow compressed in the compressor section 2 is mixed with fuel and is burned in the combustor assembly 3, possibly added with dilution air.
- the burned mixture is then expanded to the turbine section 5 and converted in mechanical power.
- the combustor assembly 3 is a two-stage sequential combustor and comprises a plurality of can combustors 10 arranged around the main axis A.
- Each of the can combustors 10, one of which is shown in Figure 2 comprises a first-stage combustor 12 and a second-stage combustor 13 sequentially arranged and defining a flow channel 15.
- An axial injector lance 16 extends from the first-stage combustor 12 into the second-stage combustor 13.
- the first-stage combustor 12 comprises a burner 17 and a first-stage combustion chamber 18.
- the second-stage combustor 13 which is illustrated in greater detail in Figure 3 , is arranged downstream of the first-stage combustor 12 and includes a second-stage combustion chamber 20 extending along an axial direction and a transition duct 22 for coupling to the turbine section 5, here not shown. Moreover, a flame stabilizer device 23 is provided in the second-stage combustor 13.
- the second-stage combustion chamber 20 extends along an axial direction downstream of the first-stage combustor 12.
- the second-stage combustion chamber 20 comprises an outer liner 24 and inner liner 25.
- the outer liner 24 surrounds the inner liner 25 at a distance therefrom, so that a cooling channel 27 is defined between the outer liner 24 and the inner liner 25.
- the inner liner 25 delimits the flow channel 15 outwards in the second-stage combustion chamber 20 and forms a transition region 28 that joins the first-stage combustor 12 in such a way to define a smooth transition without steps and possibly sharp edges.
- the axial lance injector 16 comprises an elongated streamlined body 30 that extends in the axial direction from the first-stage combustor 12 into the second-stage combustion chamber 20 through the transition region 28 of the second-stage combustor 13. A downstream end of the elongated streamlined body 30 is arranged at an interface between the second-stage combustion chamber 20 and the transition duct 22.
- the first-stage combustor 12 comprises an upstream end-cap 31 and the elongated streamlined body 30 is supported at the upstream end-cap 31 together with the burner 17 (see figure 2 ).
- the elongated streamlined body 30 and the transition region 28 are configured to prevent gas flowing from the first-stage combustor 12 to the second-stage combustion chamber 20 from recirculating in the transition region 24.
- the elongated streamlined body 30 has a smooth ellipsoidal surface tapering towards a downstream end 30a.
- the surface of the elongated streamlined body 30 may have different smooth shape, however, such as generally oblong, conical or cylindrical.
- the downstream end 30a of the elongated streamlined body 30 may be truncated.
- a plurality of first cross-flow injection nozzles 32 and a plurality of second cross-flow injection nozzles 33 are provided on the elongated streamlined body 30 at respective axial locations.
- the first cross-flow injection nozzles 32 are all at a first axial location and the second cross-flow injection nozzles 33 are all at a second axial location.
- the second cross-flow injection nozzles 33 are axially displaced downstream of the first cross-flow injection nozzles 32.
- the second cross-flow injection nozzles 33 are arranged nearer to the outlet of the second-stage combustion chamber 20 than the first cross-flow injection nozzles 32.
- the first cross-flow injection nozzles 32 and the second cross-flow injection nozzles 33 are fluidly coupled to a fuel gas supply line 35 and to a fuel oil supply line 36, respectively. Terminal portions of the fuel gas supply line 35 and of a fuel oil supply line 36 are accommodated inside the elongated streamlined body 30. Accordingly, fuel gas and fuel oil may be separately fed to the second-stage combustion chamber 20. In addition, fuel oil is supplied at a location displaced axially downstream with respect to the fuel gas.
- Fuel gas is injected through the first cross-flow injection nozzles 32 in a direction substantially perpendicular to an axis B of the axial lance injector 16.
- the second cross-flow injection nozzles 33 may be inclined to inject fuel oil in an inclined direction, that form an injection angle ⁇ with the axis B of the axial lance injector 16.
- the angle ⁇ may be comprised between 30° and 90°.
- Radial and/or inclined sleeves (not shown) may be provided as desired to increase penetration of fuel gas and fuel oil, respectively.
- the flame stabilizer device 23 is arranged downstream of the first cross-flow injection nozzles 32 and of the second cross-flow injection nozzles 33 and is configured to anchor the flame selectively at one of a plurality of flame locations.
- the flame stabilizer device 23 is controlled by the controller 7 and comprises a full-load flame stabilizer 40 at a downstream end of the elongated streamlined body 30 and a at least one partial-load flame stabilizer 41 between the second cross-flow injection nozzles 33 and the full-load flame stabilizer 40.
- the full-load flame stabilizer 40 comprises a set of full-load electrodes 40a on the elongated streamlined body 30 and a full-load voltage supply line 40b running inside the elongated streamlined body 30.
- the full-load flame stabilizer 40 produces sparks across the second-stage combustion chamber 20 and causes ignition of the mixture flowing through the second-stage combustion chamber 20 irrespective of temperature conditions and of the self-ignition time of the mixture.
- the self-ignition time of the mixture may be even so long that the mixture would not self-ignite within the second-stage combustion chamber 20, but the full-load flame stabilizer 40 is in any case capable of stabilizing the flame at the downstream end of the elongated streamlined body 30.
- the partial-load flame stabilizer 41 comprises a set of partial-load electrodes 41a on the streamlined body 30 and an upstream voltage supply line 41b.
- the partial-load electrodes 41a are arranged between the second cross-flow injection nozzles 33 and the full-load flame stabilizer 40.
- the full-load flame stabilizer 40 and the partial-load flame stabilizer 41 are selectively activated by the controller 7 on the basis of the load determined for the gas turbine assembly 1.
- the controller 7 activates the full-load flame stabilizer 40 and deactivates the partial-load flame stabilizer 41, thus setting a current flame location at the downstream end of the elongated streamlined body 30.
- the controller 7 activates the partial-load flame stabilizer 41 and deactivates the full-load flame stabilizer 40. Accordingly, the current flame location is moved upstream towards the crossflow injection nozzles 32, 33.
- the low load threshold does not exceed the high load threshold.
- the controller 7 controls an inlet gas temperature of hot gas flowing from the first-stage combustor 12 to the second-stage combustor 13.
- the controller 7 may act e.g. on a power split or power ratio of power delivered by the first-stage combustor 12 to power delivered by the second-stage combustor 13, and/or on a flow of dilution air admixed to the hot gas from the first-stage combustor 12 before entering the second-stage combustor 13.
- the controller 7 uses temperature control to help set a current flame location at an upstream region of the second-stage combustion chamber 20 (by increasing the inlet gas temperature at partial-load, with or without the aid of a flame stabilizer) or at a downstream region of the outlet of the second-stage combustion chamber 20 (by decreasing the inlet gas temperature at full-load; the full-load flame stabilizer 40 causes ignition of the mixture flowing through the second-stage combustion chamber 20 irrespective of temperature conditions, so the self-ignition time of the mixture may be even so long that the mixture would not self-ignite within the second-stage combustion chamber 20).
- Vortex generators 42 may be provided upstream of the second-stage combustion chamber 20 on the inner liner 25, for example in the transition region 32.
- the vortex generators 42 are configured to cause flow swirl by adding tangential component of velocity.
- the vortex generators 42 may be e.g. in the form of prismatic projections (see figure 4 by way of example), baffles, deflectors, lobes superficial roughness of the inner liner 25 or have any other suitable shape.
- At least one of the first cross-flow injection nozzles is axially displaced with respect to the other first cross-flow injection nozzles 132.
- the first cross-flow injection nozzles 132 may be staggered in the axial direction and arranged along a helical line on the elongated streamlined body 30 with uniform spacing in the circumferential direction.
- the second cross-flow injection nozzles here designated by numeral 133, is axially displaced with respect to the other second cross-flow injection nozzles 133.
- the second cross-flow injection nozzles 133 are staggered in the axial direction and arranged along a helical line on the elongated streamlined body 30 with uniform spacing in the circumferential direction.
- Figure 7 shows another embodiment of the invention.
- the flame stabilizer device here designate by numeral 223, comprises a change in cross section of the second-stage combustion chamber, here 220, in the axial direction.
- the change in cross section is configured to cause gas flowing through the second-stage combustion chamber to recirculate at the downstream flame location and cause flow stagnation.
- the change in cross section may be a sharp annular edge, as in the example illustrated in figure 7 .
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- Engineering & Computer Science (AREA)
- Chemical & Material Sciences (AREA)
- Combustion & Propulsion (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
Priority Applications (2)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
EP17201920.0A EP3486570B1 (de) | 2017-11-15 | 2017-11-15 | Sekundäre brennkammerstufe für eine sequentielle gasturbinenbrennkammer |
CN201811359939.1A CN110030579B (zh) | 2017-11-15 | 2018-11-15 | 用于燃气涡轮的连续燃烧器的第二级燃烧器 |
Applications Claiming Priority (1)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
EP17201920.0A EP3486570B1 (de) | 2017-11-15 | 2017-11-15 | Sekundäre brennkammerstufe für eine sequentielle gasturbinenbrennkammer |
Publications (2)
Publication Number | Publication Date |
---|---|
EP3486570A1 true EP3486570A1 (de) | 2019-05-22 |
EP3486570B1 EP3486570B1 (de) | 2023-06-21 |
Family
ID=60327229
Family Applications (1)
Application Number | Title | Priority Date | Filing Date |
---|---|---|---|
EP17201920.0A Active EP3486570B1 (de) | 2017-11-15 | 2017-11-15 | Sekundäre brennkammerstufe für eine sequentielle gasturbinenbrennkammer |
Country Status (2)
Country | Link |
---|---|
EP (1) | EP3486570B1 (de) |
CN (1) | CN110030579B (de) |
Families Citing this family (1)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
CN112695267A (zh) * | 2020-12-28 | 2021-04-23 | 郑州立佳热喷涂机械有限公司 | 气体点火式内孔超音速喷枪 |
Citations (6)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US5802854A (en) * | 1994-02-24 | 1998-09-08 | Kabushiki Kaisha Toshiba | Gas turbine multi-stage combustion system |
US6895759B2 (en) * | 2001-02-02 | 2005-05-24 | Alstom Technology Ltd | Premix burner and method of operation |
EP2072899A1 (de) * | 2007-12-19 | 2009-06-24 | ALSTOM Technology Ltd | Kraftstoffeinspritzsystem |
EP2116769A1 (de) * | 2008-05-09 | 2009-11-11 | ALSTOM Technology Ltd | Brennstofflanze für einen Brenner |
EP2647911A2 (de) * | 2012-04-05 | 2013-10-09 | General Electric Company | Brennkammer und Verfahren zur Versorgung einer Brennkammer mit Brennstoff |
EP3015772A1 (de) * | 2014-10-31 | 2016-05-04 | Alstom Technology Ltd | Brennkammeranordnung für eine gasturbine |
Family Cites Families (2)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
GB0515034D0 (en) * | 2005-07-21 | 2005-08-31 | Rolls Royce Plc | Method and system for operating a multi-stage combustor |
US8176739B2 (en) * | 2008-07-17 | 2012-05-15 | General Electric Company | Coanda injection system for axially staged low emission combustors |
-
2017
- 2017-11-15 EP EP17201920.0A patent/EP3486570B1/de active Active
-
2018
- 2018-11-15 CN CN201811359939.1A patent/CN110030579B/zh active Active
Patent Citations (6)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US5802854A (en) * | 1994-02-24 | 1998-09-08 | Kabushiki Kaisha Toshiba | Gas turbine multi-stage combustion system |
US6895759B2 (en) * | 2001-02-02 | 2005-05-24 | Alstom Technology Ltd | Premix burner and method of operation |
EP2072899A1 (de) * | 2007-12-19 | 2009-06-24 | ALSTOM Technology Ltd | Kraftstoffeinspritzsystem |
EP2116769A1 (de) * | 2008-05-09 | 2009-11-11 | ALSTOM Technology Ltd | Brennstofflanze für einen Brenner |
EP2647911A2 (de) * | 2012-04-05 | 2013-10-09 | General Electric Company | Brennkammer und Verfahren zur Versorgung einer Brennkammer mit Brennstoff |
EP3015772A1 (de) * | 2014-10-31 | 2016-05-04 | Alstom Technology Ltd | Brennkammeranordnung für eine gasturbine |
Also Published As
Publication number | Publication date |
---|---|
CN110030579B (zh) | 2023-03-21 |
CN110030579A (zh) | 2019-07-19 |
EP3486570B1 (de) | 2023-06-21 |
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