EP0059490A1 - Chambre de combustion annulaire avec brûleur annulaire pour turbines à gaz - Google Patents

Chambre de combustion annulaire avec brûleur annulaire pour turbines à gaz Download PDF

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Publication number
EP0059490A1
EP0059490A1 EP82200099A EP82200099A EP0059490A1 EP 0059490 A1 EP0059490 A1 EP 0059490A1 EP 82200099 A EP82200099 A EP 82200099A EP 82200099 A EP82200099 A EP 82200099A EP 0059490 A1 EP0059490 A1 EP 0059490A1
Authority
EP
European Patent Office
Prior art keywords
ring
gas
burner
channels
combustion chamber
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
EP82200099A
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German (de)
English (en)
Other versions
EP0059490B1 (fr
Inventor
Bernard Dipl.-Ing. Matt
Theo Woringer
Gerassime Dr.Dipl.-Ing. Zouzoulas
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 EP0059490A1 publication Critical patent/EP0059490A1/fr
Application granted granted Critical
Publication of EP0059490B1 publication Critical patent/EP0059490B1/fr
Expired legal-status Critical Current

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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
    • 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
    • F23R3/04—Air inlet arrangements
    • F23R3/10—Air inlet arrangements for primary air
    • 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

Definitions

  • the present invention relates to an annular combustion chamber with an annular burner for gas turbines according to the preamble of patent claim 1.
  • ring combustion chambers Compared to individual combustion chambers, ring combustion chambers have i.a. the advantage of a more compact design of the gas turbines.
  • Today's burners for ring combustion chambers consist of a relatively small number of individual burners distributed over the circumference of the ring combustion chamber, generally 10 to 20, in exceptional cases up to 48.
  • the temperature distribution in the gas stream at the inlet to the turbine is therefore, as already said, not as evenly as desired would be worth it, especially with a small number of individual burners.
  • these burners require a satisfactory flame stabilization, a large recirculation zone which is generated with swirl generators or flame holders with large blocking, which contribute to the pressure losses in the combustion chamber.
  • Another disadvantage of such conventional burners is that, at least in the ignition zone of the fuel / air mixture, there are stoichiometric conditions and thus locally high flame temperatures, which favors the formation of undesirable nitrogen oxides.
  • the total air flow through the burner apart from the cooling air flow for the combustion chamber wall, is divided into a primary air flow that flows through the combustion zone and one or more mixed air flows, which have to be mixed and swirled with the combustion gases further downstream after the burner outlet. for which high speeds with correspondingly large pressure losses are required.
  • the concept on which the invention is based is that very good, intimate mixing of the air with the gaseous and / or liquid fuel is to take place even before the ignition zone, which is a lower temperature peak, a more even temperature distribution in front of the gas turbine and reduced nitrogen oxide formation. A correspondingly selected air speed prevents reignitions. In addition, this eliminates the usual, strongly increasing elements for generating turbulence or a backflow, so that the associated pressure losses are avoided.
  • the ring combustion chamber with ring burner according to the invention should also be able to be set up for both gaseous and liquid fuels and for simultaneous operation with gaseous and liquid fuels.
  • combustion chamber 1 shows the arrangement of an annular combustion chamber according to the invention with an annular burner within an otherwise conventional gas turbine.
  • the combination of combustion chamber 2 and burner 3 is designated 1 here and has a common housing.
  • the combustion chamber 2 and the burner 3 will be separate components, in particular in the case of larger units, since in these, as will be explained below, the ring burner 3 is preferably composed of sectors.
  • the combustion air is, apart from a small amount of cooling air, which is branched off at taps 4, 5 and 6 for cooling the shaft channel and the housing, from the compressor 7 through an annular diffuser 8, which widens in front of the ring burner 3 to form a shock diffuser 9 Burner 3 promoted, where it is mixed intimately with the fuel gas or in addition to the fuel gas with an atomized liquid fuel over the entire channel cross section.
  • the combustion mixture ignites and the combustion gases pass through the combustion chamber 2, where the cooling air branched off in front of the burner is fed to the turbine 10 for work.
  • an annular trip bar 11 is provided, which serves as a turbulence generator and thus causes an approximately uniform speed distribution over the height of the diffuser channel.
  • the combustion chamber 2 can, as a result of the advantageous properties of the ring burner yet to be explained, be designed as an essentially smooth duct according to FIG. 1 without the usual internals for swirling the combustion mixture.
  • the description of the invention is therefore limited to the ring burner alone, which, as already mentioned, is generally designed as a component separate from the ring combustion chamber.
  • the ring burner 3 is preferably composed of circular ring sectors, in particular in the case of larger units. The number of such sectors will generally depend on the size of the burner.
  • the sector 12 shown in FIGS. 2 and 3 in a view and in a radial section covers 22.5 ° , ie that the associated entire burner consists of 16 such sectors.
  • the outermost part of the sector is formed by the gas distributor box 13, which, as shown in FIG. 3, is divided by a partition 14 into a main gas chamber 15 and an ignition gas chamber 16, to which the gas is supplied through the gas supply lines 17 and 18. These two gas supply lines in turn branch off from a manifold, not shown.
  • Radial plate channels 19 and 20 branch off from the gas distribution box 13 from the two gas chambers 15 and 16 and are cut perpendicularly by plate channels 21 and 22 running in the circumferential direction.
  • the plate channels 19 to 22 form a grid-like channel network communicating with the gas distribution box 13, which limits honeycombs of approximately square cross section, into which gas flows during operation from nozzles 23, 24 which are provided in all channel walls.
  • Out 3 shows that a row of nozzles lying in one plane is provided for each honeycomb for both the main gas and the ignition gas.
  • two or more such rows of nozzles could also be provided, which can either be aligned in the flow direction or staggered one behind the other.
  • honeycomb gas is supplied only from two radial and from one circumferential plate channel.
  • the cross section of the flame holder is U-shaped, with baffle plates 29 being provided in the web of flame holder nozzles 28 and in the slot-shaped outlet channel in front of the flame holder nozzles, which are offset from one another in the manner shown in FIG to achieve escaping gas jet for the support flame.
  • the support flame is ignited, which then ignites the gas flowing out of the pilot gas nozzles 24 at the burner outlet. Since both the Zündgasdüsen 24 are fed gas chamber and the flame holder nozzles 28 ünd- of the Z, the flow of gas for the support flame is approximately proportional to the flow of gas from the Zündgasdüsen with which the turbine is idling and possibly operated at low load can. For higher performance 23 main gas is switched on from the main gas nozzles.
  • the gas Due to the many gas nozzles 23 and 24 distributed evenly over the inner circumference of the honeycomb channels in connection with the long mixing path up to the burner outlet, the gas is mixed very well with the air even before the burner outlet, without swirling, so that with a small pressure drop there is one very good combustion with a large excess of air takes place over the entire cross-section of the burner.
  • the temperature of the turbine is also correspondingly balanced by the combustion gases, to which the cooling air extracted at the tapping points 4, 5 and 6 in the combustion chamber is only supplied through slots 30 in the combustion chamber wall in the edge zone.
  • FIG. 4 shows a radial section through a sector of a dual burner that can be operated with liquid and gaseous fuel.
  • this has fuel nozzles 31 arranged in a radial row in front of the burner outlet, which are connected in rows or in groups via fuel lines 32, depending on the load condition, in addition to the gas or alone after the burner with the ignition gas has been ramped up to the idle power, that to the operation supply the combustion gas required under load, the pilot gas being able to be switched off since the flame stabilization is then brought about by the backflow zone prevailing at the burner outlet as a result of the eddies.
  • the axes of the fuel nozzles 31 are aligned with the intersection lines of the radial and the circumferentially extending plate channels, so that the fuel jet is atomized into the four honeycomb channels at the points of intersection of the plates.
  • FIG. 5 to 7 schematically show this distribution of the fuel jets and the fuel nozzles active under different load conditions, namely FIG. 5 at idle, FIG. 6 at partial load and FIG. 7 at full load.
  • Different combinations of active fuel nozzles are possible for part load in a known manner, depending on the individual case.
  • the ignition gas supplied via the ignition gas chamber 32, ignition gas channels 33 and longitudinal pipes 34 branching from the latter to a pipe grid provided at the burner outlet only serves to stabilize the flame.
  • the turbine is operated exclusively by main gas, which passes from the main gas chamber 35 into radial plate channels 36 and from these through main gas nozzles 37 into the air channels formed by adjacent plate channels 36.
  • the longitudinal pipes 34 parallel to the turbine axis open into the aforementioned pipe grid network at the nodes, which are formed by intersecting radial pipes 38 and ring pipes 39.
  • Both the radial tubes 38 and the ring tubes 39 are each with two Provided rows of flame holder nozzles 40 and 41, the axes of which are inclined at an acute angle to the flow direction of the burner.

Landscapes

  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Gas Burners (AREA)
EP82200099A 1981-03-04 1982-01-27 Chambre de combustion annulaire avec brûleur annulaire pour turbines à gaz Expired EP0059490B1 (fr)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
CH1439/81 1981-03-04
CH143981 1981-03-04

Publications (2)

Publication Number Publication Date
EP0059490A1 true EP0059490A1 (fr) 1982-09-08
EP0059490B1 EP0059490B1 (fr) 1984-12-12

Family

ID=4210486

Family Applications (1)

Application Number Title Priority Date Filing Date
EP82200099A Expired EP0059490B1 (fr) 1981-03-04 1982-01-27 Chambre de combustion annulaire avec brûleur annulaire pour turbines à gaz

Country Status (5)

Country Link
US (1) US4455840A (fr)
EP (1) EP0059490B1 (fr)
JP (1) JPS57157936A (fr)
CA (1) CA1189330A (fr)
DE (1) DE3261484D1 (fr)

Cited By (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE19549143A1 (de) * 1995-12-29 1997-07-03 Abb Research Ltd Gasturbinenringbrennkammer
EP0818658A1 (fr) * 1996-07-11 1998-01-14 SOCIETE NATIONALE D'ETUDE ET DE CONSTRUCTION DE MOTEURS D'AVIATION -Snecma Chambre de combustion anti-nox à injection de carburant de type annulaire
EP1614963A1 (fr) * 2004-07-09 2006-01-11 Siemens Aktiengesellschaft Système de combustion à prémélange et procédé
EP1816400A2 (fr) * 2006-02-02 2007-08-08 Rolls-Royce Deutschland Ltd & Co KG Chambre de combustion de turbine à gaz dotée d'une injection de carburant sur la totalité de l'anneau de chambre de combustion
CN109028147A (zh) * 2017-06-09 2018-12-18 通用电气公司 环形喉道旋转爆震燃烧器和相应的推进***

Families Citing this family (42)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
JPS6082724A (ja) * 1983-10-13 1985-05-10 Agency Of Ind Science & Technol ガスタ−ビン燃焼器
DE3662462D1 (en) * 1985-07-30 1989-04-20 Bbc Brown Boveri & Cie Dual combustor
JPS63121310U (fr) * 1987-01-31 1988-08-05
JPS63121317U (fr) * 1987-01-31 1988-08-05
CH684963A5 (de) * 1991-11-13 1995-02-15 Asea Brown Boveri Ringbrennkammer.
FR2694624B1 (fr) * 1992-08-05 1994-09-23 Snecma Chambre de combustion à plusieurs injecteurs de carburant.
US5289685A (en) * 1992-11-16 1994-03-01 General Electric Company Fuel supply system for a gas turbine engine
US5323604A (en) * 1992-11-16 1994-06-28 General Electric Company Triple annular combustor for gas turbine engine
US5303542A (en) * 1992-11-16 1994-04-19 General Electric Company Fuel supply control method for a gas turbine engine
US5943866A (en) * 1994-10-03 1999-08-31 General Electric Company Dynamically uncoupled low NOx combustor having multiple premixers with axial staging
US5881756A (en) * 1995-12-22 1999-03-16 Institute Of Gas Technology Process and apparatus for homogeneous mixing of gaseous fluids
US6109038A (en) * 1998-01-21 2000-08-29 Siemens Westinghouse Power Corporation Combustor with two stage primary fuel assembly
SE514341C2 (sv) 1998-06-18 2001-02-12 Abb Ab Förfarande för start av en brännaranordning till en gasturbin
DE59810344D1 (de) * 1998-07-27 2004-01-15 Alstom Switzerland Ltd Verfahren zum Betrieb einer Gasturbinenbrennkammer mit gasförmigem Brennstoff
US6286298B1 (en) * 1998-12-18 2001-09-11 General Electric Company Apparatus and method for rich-quench-lean (RQL) concept in a gas turbine engine combustor having trapped vortex cavity
US6295801B1 (en) * 1998-12-18 2001-10-02 General Electric Company Fuel injector bar for gas turbine engine combustor having trapped vortex cavity
US6442939B1 (en) * 2000-12-22 2002-09-03 Pratt & Whitney Canada Corp. Diffusion mixer
US6427447B1 (en) * 2001-02-06 2002-08-06 United Technologies Corporation Bulkhead for dual fuel industrial and aeroengine gas turbines
JP2003074854A (ja) * 2001-08-28 2003-03-12 Honda Motor Co Ltd ガスタービン・エンジンの燃焼器
JP2003074856A (ja) * 2001-08-28 2003-03-12 Honda Motor Co Ltd ガスタービン・エンジンの燃焼器
JP2003074853A (ja) * 2001-08-28 2003-03-12 Honda Motor Co Ltd ガスタービン・エンジンの燃焼器
EP1507120A1 (fr) * 2003-08-13 2005-02-16 Siemens Aktiengesellschaft Turbine à gaz
US7249461B2 (en) * 2003-08-22 2007-07-31 Siemens Power Generation, Inc. Turbine fuel ring assembly
ES2400247T3 (es) * 2008-12-19 2013-04-08 Alstom Technology Ltd Quemador de una turbina de gas que tiene una configuración de lanza especial
US8234871B2 (en) * 2009-03-18 2012-08-07 General Electric Company Method and apparatus for delivery of a fuel and combustion air mixture to a gas turbine engine using fuel distribution grooves in a manifold disk with discrete air passages
AU2009352303B2 (en) 2009-09-13 2015-07-30 Lean Flame Inc. Method of fuel staging in combustion apparatus
US8276385B2 (en) 2009-10-08 2012-10-02 General Electric Company Staged multi-tube premixing injector
EA021650B1 (ru) * 2011-08-29 2015-08-31 Геннадий Борисович Варламов Многокамерная газовая горелка трубчатого типа
EA019766B1 (ru) * 2011-08-29 2014-06-30 Геннадий Борисович Варламов Низкоэмиссионная газовая горелка трубчатого типа с направленным воздушным потоком
US8429915B1 (en) * 2011-10-17 2013-04-30 General Electric Company Injector having multiple fuel pegs
US9341376B2 (en) * 2012-02-20 2016-05-17 General Electric Company Combustor and method for supplying fuel to a combustor
US9677766B2 (en) * 2012-11-28 2017-06-13 General Electric Company Fuel nozzle for use in a turbine engine and method of assembly
JP6239943B2 (ja) * 2013-11-13 2017-11-29 三菱日立パワーシステムズ株式会社 ガスタービン燃焼器
WO2015176908A1 (fr) * 2014-05-23 2015-11-26 Siemens Aktiengesellschaft Brûleur équipé d'une bague de distribution de combustible
USD791930S1 (en) 2015-06-04 2017-07-11 Tropitone Furniture Co., Inc. Fire burner
US10197291B2 (en) 2015-06-04 2019-02-05 Tropitone Furniture Co., Inc. Fire burner
US11226092B2 (en) 2016-09-22 2022-01-18 Utilization Technology Development, Nfp Low NOx combustion devices and methods
US11326521B2 (en) * 2020-06-30 2022-05-10 General Electric Company Methods of igniting liquid fuel in a turbomachine
US11333360B2 (en) * 2020-09-25 2022-05-17 General Electric Company Fuel injector for a turbomachine
EP4027059A1 (fr) * 2021-01-12 2022-07-13 Crosstown Power GmbH Bruleur, chambre de combustion et procede d'adaptation d'un appareil de combustion
GB202104885D0 (en) * 2021-04-06 2021-05-19 Siemens Energy Global Gmbh & Co Kg Combustor for a Gas Turbine
EP4411244A1 (fr) * 2023-02-02 2024-08-07 Pratt & Whitney Canada Corp. Chambre de combustion à mélange d'air et de carburant distribué

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Publication number Priority date Publication date Assignee Title
GB568681A (en) * 1943-04-15 1945-04-16 Fred Umpleby Improvements in and relating to internal combustion turbine plant
US2592748A (en) * 1944-02-17 1952-04-15 Rateau Soc Annular combustion chamber with hollow air guide vanes with radial gasiform fuel slots for gas turbines
US2635424A (en) * 1947-08-02 1953-04-21 Szczeniowski Boleslaw Combustor for high flow velocities
CH305532A (de) * 1951-06-25 1955-02-28 Parsons & Co Ltd C A Gasturbinenanlage.
FR1166520A (fr) * 1956-11-20 1958-11-12 Continental Aviat & Eng Corp Brûleur-veilleuse pour réacteurs
DE1120820B (de) * 1957-05-09 1961-12-28 Bristol Siddeley Engines Ltd Vorrichtung zur feinverteilten Zufuehrung einer Fluessigkeit in einen Gasstrom
US4100733A (en) * 1976-10-04 1978-07-18 United Technologies Corporation Premix combustor

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US2701444A (en) * 1950-01-26 1955-02-08 Solar Aircraft Co Burner for jet engines
US2712221A (en) * 1952-04-22 1955-07-05 Westinghouse Electric Corp Gas turbine afterburner apparatus
US3046731A (en) * 1955-10-07 1962-07-31 Edward Pohlmann Flame stabilization in jet engines
JPS4921515A (fr) * 1972-05-22 1974-02-26
GB1559779A (en) * 1975-11-07 1980-01-23 Lucas Industries Ltd Combustion assembly
US4158949A (en) * 1977-11-25 1979-06-26 General Motors Corporation Segmented annular combustor
GB2036296B (en) * 1978-11-20 1982-12-01 Rolls Royce Gas turbine

Patent Citations (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB568681A (en) * 1943-04-15 1945-04-16 Fred Umpleby Improvements in and relating to internal combustion turbine plant
US2592748A (en) * 1944-02-17 1952-04-15 Rateau Soc Annular combustion chamber with hollow air guide vanes with radial gasiform fuel slots for gas turbines
US2635424A (en) * 1947-08-02 1953-04-21 Szczeniowski Boleslaw Combustor for high flow velocities
CH305532A (de) * 1951-06-25 1955-02-28 Parsons & Co Ltd C A Gasturbinenanlage.
FR1166520A (fr) * 1956-11-20 1958-11-12 Continental Aviat & Eng Corp Brûleur-veilleuse pour réacteurs
DE1120820B (de) * 1957-05-09 1961-12-28 Bristol Siddeley Engines Ltd Vorrichtung zur feinverteilten Zufuehrung einer Fluessigkeit in einen Gasstrom
US4100733A (en) * 1976-10-04 1978-07-18 United Technologies Corporation Premix combustor

Cited By (8)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE19549143A1 (de) * 1995-12-29 1997-07-03 Abb Research Ltd Gasturbinenringbrennkammer
EP0818658A1 (fr) * 1996-07-11 1998-01-14 SOCIETE NATIONALE D'ETUDE ET DE CONSTRUCTION DE MOTEURS D'AVIATION -Snecma Chambre de combustion anti-nox à injection de carburant de type annulaire
FR2751054A1 (fr) * 1996-07-11 1998-01-16 Snecma Chambre de combustion anti-nox a injection de carburant de type annulaire
US5937653A (en) * 1996-07-11 1999-08-17 Societe Nationale D'etude Et De Construction De Moteurs D'aviation (Snecma) Reduced pollution combustion chamber having an annular fuel injector
EP1614963A1 (fr) * 2004-07-09 2006-01-11 Siemens Aktiengesellschaft Système de combustion à prémélange et procédé
EP1816400A2 (fr) * 2006-02-02 2007-08-08 Rolls-Royce Deutschland Ltd & Co KG Chambre de combustion de turbine à gaz dotée d'une injection de carburant sur la totalité de l'anneau de chambre de combustion
EP1816400A3 (fr) * 2006-02-02 2012-07-04 Rolls-Royce Deutschland Ltd & Co KG Chambre de combustion de turbine à gaz dotée d'une injection de carburant sur la totalité de l'anneau de chambre de combustion
CN109028147A (zh) * 2017-06-09 2018-12-18 通用电气公司 环形喉道旋转爆震燃烧器和相应的推进***

Also Published As

Publication number Publication date
US4455840A (en) 1984-06-26
DE3261484D1 (en) 1985-01-24
JPS57157936A (en) 1982-09-29
CA1189330A (fr) 1985-06-25
JPS6339812B2 (fr) 1988-08-08
EP0059490B1 (fr) 1984-12-12

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