EP3486570A1 - Sekundäre brennkammerphase für eine sequentielle gasturbinenbrennkammer - Google Patents

Sekundäre brennkammerphase für eine sequentielle gasturbinenbrennkammer Download PDF

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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
Application number
EP17201920.0A
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English (en)
French (fr)
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EP3486570B1 (de
Inventor
Andrea Ciani
Armando Alsina
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
Ansaldo Energia Switzerland AG
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Ansaldo Energia Switzerland AG
Priority date (The priority date 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 date listed.)
Filing date
Publication date
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Priority to EP17201920.0A priority Critical patent/EP3486570B1/de
Priority to CN201811359939.1A priority patent/CN110030579B/zh
Publication of EP3486570A1 publication Critical patent/EP3486570A1/de
Application granted granted Critical
Publication of EP3486570B1 publication Critical patent/EP3486570B1/de
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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)
EP17201920.0A 2017-11-15 2017-11-15 Sekundäre brennkammerstufe für eine sequentielle gasturbinenbrennkammer Active EP3486570B1 (de)

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)

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EP3486570A1 true EP3486570A1 (de) 2019-05-22
EP3486570B1 EP3486570B1 (de) 2023-06-21

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EP17201920.0A Active EP3486570B1 (de) 2017-11-15 2017-11-15 Sekundäre brennkammerstufe für eine sequentielle gasturbinenbrennkammer

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EP (1) EP3486570B1 (de)
CN (1) CN110030579B (de)

Families Citing this family (1)

* Cited by examiner, † Cited by third party
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)

* Cited by examiner, † Cited by third party
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)

* Cited by examiner, † Cited by third party
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

Patent Citations (6)

* Cited by examiner, † Cited by third party
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
EP3486570B1 (de) 2023-06-21
CN110030579A (zh) 2019-07-19

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