EP0276397B1 - Brennkammer für Gasturbine - Google Patents

Brennkammer für Gasturbine Download PDF

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Publication number
EP0276397B1
EP0276397B1 EP87117059A EP87117059A EP0276397B1 EP 0276397 B1 EP0276397 B1 EP 0276397B1 EP 87117059 A EP87117059 A EP 87117059A EP 87117059 A EP87117059 A EP 87117059A EP 0276397 B1 EP0276397 B1 EP 0276397B1
Authority
EP
European Patent Office
Prior art keywords
burners
primary
combustion
combustion chamber
chamber according
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.)
Expired - Lifetime
Application number
EP87117059A
Other languages
German (de)
English (en)
French (fr)
Other versions
EP0276397A1 (de
Inventor
Jaan Dr. Hellat
Jakob Dr. Keller
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.)
BBC Brown Boveri AG Switzerland
Original Assignee
BBC Brown Boveri AG Switzerland
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
Application filed by BBC Brown Boveri AG Switzerland filed Critical BBC Brown Boveri AG Switzerland
Publication of EP0276397A1 publication Critical patent/EP0276397A1/de
Application granted granted Critical
Publication of EP0276397B1 publication Critical patent/EP0276397B1/de
Anticipated expiration legal-status Critical
Expired - Lifetime legal-status Critical Current

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Classifications

    • 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 
    • F23C6/00—Combustion apparatus characterised by the combination of two or more combustion chambers or combustion zones, e.g. for staged combustion
    • F23C6/04—Combustion apparatus characterised by the combination of two or more combustion chambers or combustion zones, e.g. for staged combustion in series connection
    • F23C6/042—Combustion apparatus characterised by the combination of two or more combustion chambers or combustion zones, e.g. for staged combustion in series connection with fuel supply in stages
    • 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 
    • F23C6/00—Combustion apparatus characterised by the combination of two or more combustion chambers or combustion zones, e.g. for staged combustion
    • F23C6/04—Combustion apparatus characterised by the combination of two or more combustion chambers or combustion zones, e.g. for staged combustion in series connection
    • F23C6/045—Combustion apparatus characterised by the combination of two or more combustion chambers or combustion zones, e.g. for staged combustion in series connection with staged combustion in a single enclosure
    • F23C6/047—Combustion apparatus characterised by the combination of two or more combustion chambers or combustion zones, e.g. for staged combustion in series connection with staged combustion in a single enclosure with fuel supply in stages
    • 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

Definitions

  • the present invention relates to a combustion chamber of gas turbines for operation with liquid fuels. It also relates to a method for operating such a combustion chamber.
  • the low NOx emission values still tolerated by law can only be maintained in the case of a planar combustion if the time the gas particles stay in hot oxygen-rich zones is as short as possible, namely not more than a few milliseconds.
  • the temperature in the reaction area must not fall below a certain limit.
  • This grading can mean either a substoichiometric primary combustion zone with subsequent afterburning at low temperatures or the stepwise connection of over-stoichiometric burner elements. In any case, the grading also requires a powerful mixing mechanism.
  • premix combustion has proven to be the technically best measure for NO x reduction for the combustion of gaseous fuels.
  • a premix combustion can consist, for example, of a premix process with a large air ratio taking place within a number of tubular elements between the fuel and the compressor air before the actual combustion process takes place downstream of a flame holder.
  • the emission values of pollutants from combustion can be significantly reduced.
  • Combustion with the largest possible air ratio - given that the flame is still burning at all and furthermore that not too much CO is produced - not only reduces the amount of NO x pollutants, but also causes a consistent reduction of other pollutants, namely as already mentioned of CO and unburned hydrocarbons.
  • this optimization process can be driven in such a way that the space for combustion and post-reaction is kept much longer than would be necessary for the actual combustion.
  • a combustion chamber of a gas turbine which is designed for the combustion of a gaseous fuel.
  • This combustion chamber is of tubular construction, with a further combustion zone being connected downstream of the main combustion part at the beginning of the combustion chamber.
  • the main burning part consists of a short, hollow cone, placed in the middle of the combustion chamber, which is surrounded by a number of fuel nozzles.
  • the fuel jet from these nozzles is admixed with air which flows from the outside into the front part of the combustion chamber and from the inside via the cone cavity in the area of these nozzles.
  • this proposal comes with the types of combustion already recognized above, which belong to the prior art, namely there is partial combustion with substoichiometric conditions and a premixed main combustion, the support of which is accomplished by the rear combustion chamber.
  • the object of the invention is to achieve, in a combustion chamber of the type mentioned at the beginning, comparable low NOx emission values as in combustion chambers operated with gaseous fuels, without taking the risk of self-ignition of the liquid fuels outside the combustion chamber.
  • the advantage of the invention can be seen essentially in the fact that a system is provided in a simple manner that generates low NO x emissions, this system being able to achieve the premixing without the technology and infrastructure which are per se quite complex.
  • the basic idea is to provide a primary burner and afterburner system.
  • the liquid fuel is injected directly into the combustion chamber.
  • the injected fuel is shielded with an air jacket, which is a non-self-sufficient burner.
  • the afterburner which is placed in a central chamber at the end of the primary burner chamber, is used in combination with one or more primary burners.
  • the hot gases generated by the primary burners should not be able to ignite the mixture produced by the afterburner in the immediate vicinity of the fuel nozzle of the afterburner, in order to avoid combustion under near-stoichiometric conditions. This is ensured by the shielding air jacket, which is not swirled and which initially effectively shields the fuel mist emanating from the afterburner nozzle against the external hot gases. Ignition of the afterburner mixture should only be possible when the liquid fuel introduced from the afterburner nozzle has mixed sufficiently with the shielding jacket air and with the hot gas containing air, so that the combustion in the lean mixture takes place at low temperatures.
  • Fig. 1 shows a combustion chamber for gas turbines, which is housed in the GT ring housing 1. If the entire combustion chamber is embedded in a GT ring housing 1, it is connected to the compressed air 11 from the compressor 10 like a chamber.
  • the gas turbine ring housing wall is designed to withstand the compressor end pressure.
  • the geometric shape of the combustion chamber, as the axial section 12 wants to symbolize, is cylindrical and consists of two primary combustion chambers 5, 5a arranged at the ends, which are arranged symmetrically and V-shaped with respect to the central combustion chamber 6. Of course, the primary combustion chambers 5, 5a can lie in a horizontal plane with respect to the central axis of the central combustion chamber 6.
  • the primary combustion chambers 5, 5a themselves are equipped at their front ends in the circumferential direction with a number of axially parallel primary burners 2, 2a which are dependent on the performance of the combustion chamber. These essentially consist of a fuel line 3, 3a and a swirl body 8, 8a.
  • a continuous circular cylindrical primary combustion chamber 5, 5a several self-contained combustion chamber units can be provided distributed over the circumference, each consisting of a pair of twin burners with preferably twist bodies oriented in opposite directions. This means that an effective mixing process can be generated in the individual combustion chamber units, an annular cylindrical outlet channel collecting the hot gases escaping from the individual combustion chamber units in order to then lead them to the central combustion chamber 6.
  • the continuous circular-cylindrical primary combustion chamber 5 and 5a shown here is provided, the primary burners 2 or 2a arranged axially parallel to one another there can also be alternately equipped with swirl bodies 8, 8a oriented in opposite directions.
  • an afterburner 4 is provided in each case.
  • liquid fuel 15 is fed directly into the combustion chamber and shielded with an air jacket 14.
  • the afterburner 4 is designed in such a way that it is not self-sufficient, ie its ignition requires permanent ignition.
  • the hot gases 13 generated by the primary burners 2, 2a should not be able to ignite the mixture 14/15 produced by the afterburner 4 in the immediate vicinity of the fuel nozzle of the afterburner 4.
  • This is ensured by the shielding air jacket 14, which should preferably be untwisted and initially effectively shields the fuel mist 15 emanating from the afterburner nozzle against the hot gases 13 of the primary burners 2, 2a arriving there.
  • An ignition of the afterburner mixture 14/15 should only be possible when the liquid fuel 15 introduced by the burner nozzle is sufficiently strong with the shielding air coat 14 has mixed.
  • the air ratio related to the fuel supply of the afterburner 4 and the air jacket 14 is determined according to the same criteria as for a premix burner.
  • the rapid mixing in of the hot gases 13 after they have initiated the first spark ignition of the afterburner mixture 14/15 plays an important role in the stability of the combustion, which is why it must be ensured that the pulse density ratio between primary burner gases 13 and afterburner mixture 14/15 is very high high - well over 1 - is selected. It has been confirmed that an optimally designed afterburner 4 produces hardly more NO x than a premix burner, while the primary burners 2, 2a, which of course have to be self-sufficient, for example designed as diffusion burners, cause significantly higher NO x emissions.
  • the primary burners 2, 2a should therefore be planned as small as possible and they should be operated with high air ratios: both measures make it possible to keep the NO x emissions from the operation of the primary burners 2, 2a as low as possible. Consequently, for the operation of a gas turbine combustion chamber, this means that the primary burners 2, 2a and the afterburner 4 are operated in stages.
  • the afterburner 4 is preferably switched on at a load point near zero load of the gas turbines.
  • the load is regulated only via the fuel supply to the afterburner 4, it being possible for a gradual reduction in the fuel supply to the primary burner 2, 2a to be initiated as the afterburner load increases.
  • the lower limit for the reduction of the fuel supply to the primary burners 2, 2a is given on the one hand by the extinguishing limit of the primary burner and on the other hand by the necessity that the temperature of the exhaust gas of the primary burner must be high enough to initiate the burnout of the afterburner fuel.
  • the air jacket 14 shields the afterburner 4 and its liquid fuel spray cone 15 from the incoming hot gases 13 from the primary burners 2, 2a.
  • the mixture 14/15 produced by the afterburner 4 should not come to ignition in the immediate vicinity of the fuel nozzle 15 under near-stoichiometric conditions. Ignition of the afterburner mixture 14/15 should only be possible when the liquid fuel 15 introduced from the afterburner nozzle has mixed sufficiently with the shielding air jacket 14, that is to say downstream of the central combustion chamber 6. Further downstream is the mixing chamber 7, which ensures that that a vortex-free flow with uniform overall pressure and temperature profile can arise before the turbine 9 is acted upon.
  • the length of the mixing chamber 7 is heavily dependent on the strength of the mixing process: observations have shown that a vortex-free flow with a uniform pressure is achieved well after a length of about three diameters of the corresponding combustion chamber unit.
  • the optimal design of the primary burners 2, 2a reference is made to the description according to EP-0 193 029, in particular under FIG. 2.
  • the solution shown in FIG. 2 intends to further protect the afterburner 4 from the hot gases 13 of the primary burners 2, 2a flowing in.
  • the inlet 16 of the shielding air 14 into the combustion chamber is at least 'extended such that the liquid fuel spray cone 15 is also shielded.
  • the hot gases 13 only flow further downstream to the afterburner mixture 14/15; there the mixing of the liquid fuel 15 with the shielding jacket air 14 has progressed to such an extent that this mixture 14/15 can be ignited.
  • FIG. 3 shows a further variant of how the afterburner 4 and its liquid fuel spray cone 15 can be shielded from the incoming hot gases 13 in the region of the central combustion chamber 6.
  • the shielding air 14 flows on the one hand along the afterburner 4 and on the other hand laterally between a plurality of fins 17 into the central combustion chamber 6.
  • Such a provision has the advantage that the mixing between liquid fuel 15 and shielding air 14 in front of the mixing chamber 7 is thereby optimized.
  • this mixture 14/15 is then ignited by the hot gases 13 flowing there.
  • the entire length of the mixing chamber 7 thus remains available in order to provide a vortex-free flow with a uniform pressure and temperature profile for the turbine to be acted upon.

Landscapes

  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Engine Equipment That Uses Special Cycles (AREA)
EP87117059A 1986-12-09 1987-11-19 Brennkammer für Gasturbine Expired - Lifetime EP0276397B1 (de)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
CH4892/86A CH672366A5 (ru) 1986-12-09 1986-12-09
CH4892/86 1986-12-09

Publications (2)

Publication Number Publication Date
EP0276397A1 EP0276397A1 (de) 1988-08-03
EP0276397B1 true EP0276397B1 (de) 1991-01-30

Family

ID=4284343

Family Applications (1)

Application Number Title Priority Date Filing Date
EP87117059A Expired - Lifetime EP0276397B1 (de) 1986-12-09 1987-11-19 Brennkammer für Gasturbine

Country Status (5)

Country Link
US (1) US4805411A (ru)
EP (1) EP0276397B1 (ru)
JP (1) JPS63156926A (ru)
CH (1) CH672366A5 (ru)
DE (1) DE3767873D1 (ru)

Families Citing this family (15)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4928481A (en) * 1988-07-13 1990-05-29 Prutech Ii Staged low NOx premix gas turbine combustor
US5199265A (en) * 1991-04-03 1993-04-06 General Electric Company Two stage (premixed/diffusion) gas only secondary fuel nozzle
US5259184A (en) * 1992-03-30 1993-11-09 General Electric Company Dry low NOx single stage dual mode combustor construction for a gas turbine
US5323604A (en) * 1992-11-16 1994-06-28 General Electric Company Triple annular combustor for gas turbine engine
DE19615910B4 (de) * 1996-04-22 2006-09-14 Alstom Brenneranordnung
GB2319078B (en) * 1996-11-08 1999-11-03 Europ Gas Turbines Ltd Combustor arrangement
EP0924470B1 (de) * 1997-12-19 2003-06-18 MTU Aero Engines GmbH Vormischbrennkammer für eine Gasturbine
US6430919B1 (en) * 2000-03-02 2002-08-13 Direct Propulsion Devices, Inc. Shaped charged engine
CN100529547C (zh) * 2002-01-14 2009-08-19 阿尔斯通技术有限公司 用于燃气轮机的环形燃烧室的燃烧器装置
US8387390B2 (en) 2006-01-03 2013-03-05 General Electric Company Gas turbine combustor having counterflow injection mechanism
US8178045B2 (en) * 2007-12-17 2012-05-15 University Of Louisville Research Foundation, Inc. Interchangeable preconcentrator connector assembly
US8448532B2 (en) * 2009-03-18 2013-05-28 The United States Of America As Represented By The Secretary Of The Navy Actively cooled vapor preconcentrator
RU2534189C2 (ru) * 2010-02-16 2014-11-27 Дженерал Электрик Компани Камера сгорания для газовой турбины(варианты) и способ эксплуатации газовой турбины
US20120304660A1 (en) * 2011-06-06 2012-12-06 Kupratis Daniel B Turbomachine combustors having different flow paths
US20230280035A1 (en) * 2022-03-07 2023-09-07 General Electric Company Bimodal combustion system

Family Cites Families (18)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR2221621B1 (ru) * 1973-03-13 1976-09-10 Snecma
US3872664A (en) * 1973-10-15 1975-03-25 United Aircraft Corp Swirl combustor with vortex burning and mixing
US4173118A (en) * 1974-08-27 1979-11-06 Mitsubishi Jukogyo Kabushiki Kaisha Fuel combustion apparatus employing staged combustion
US4052844A (en) * 1975-06-02 1977-10-11 Societe Nationale D'etude Et De Construction De Moteurs D'aviation Gas turbine combustion chambers
US4012904A (en) * 1975-07-17 1977-03-22 Chrysler Corporation Gas turbine burner
US4168609A (en) * 1977-12-01 1979-09-25 United Technologies Corporation Folded-over pilot burner
US4249373A (en) * 1978-01-28 1981-02-10 Rolls-Royce Ltd. Gas turbine engine
GB2013788B (en) * 1978-01-28 1982-06-03 Rolls Royce Gas turbine engine
GB2043868B (en) * 1979-03-08 1982-12-15 Rolls Royce Gas turbine
US4292801A (en) * 1979-07-11 1981-10-06 General Electric Company Dual stage-dual mode low nox combustor
GB2072827A (en) * 1980-03-29 1981-10-07 Rolls Royce A tubo-annular combustion chamber
GB2073400B (en) * 1980-04-02 1984-03-14 United Technologies Corp Fuel injector
JPS5755975U (ru) * 1980-09-16 1982-04-01
JPS57187531A (en) * 1981-05-12 1982-11-18 Hitachi Ltd Low nox gas turbine burner
JPS59202324A (ja) * 1983-05-04 1984-11-16 Hitachi Ltd ガスタ−ビン低ＮＯｘ燃焼器
JPS6057131A (ja) * 1983-09-08 1985-04-02 Hitachi Ltd ガスタ−ビン燃焼器の燃料供給方法
EP0169431B1 (en) * 1984-07-10 1990-04-11 Hitachi, Ltd. Gas turbine combustor
DE3661224D1 (en) * 1985-02-26 1988-12-22 Bbc Brown Boveri & Cie Gas turbine combustor

Also Published As

Publication number Publication date
JPS63156926A (ja) 1988-06-30
DE3767873D1 (de) 1991-03-07
EP0276397A1 (de) 1988-08-03
US4805411A (en) 1989-02-21
CH672366A5 (ru) 1989-11-15

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