EP3026347A1 - Brûleur a corps non profilé annulaire - Google Patents

Brûleur a corps non profilé annulaire Download PDF

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Publication number
EP3026347A1
EP3026347A1 EP14194792.9A EP14194792A EP3026347A1 EP 3026347 A1 EP3026347 A1 EP 3026347A1 EP 14194792 A EP14194792 A EP 14194792A EP 3026347 A1 EP3026347 A1 EP 3026347A1
Authority
EP
European Patent Office
Prior art keywords
combustor
swirler
pilot
liner
passage
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.)
Withdrawn
Application number
EP14194792.9A
Other languages
German (de)
English (en)
Inventor
Ennio Pasqualotto
Douglas Anthony Pennell
Michael Duesing
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
Original Assignee
Alstom Technology 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
Application filed by Alstom Technology AG filed Critical Alstom Technology AG
Priority to EP14194792.9A priority Critical patent/EP3026347A1/fr
Priority to US14/950,601 priority patent/US20160146464A1/en
Priority to CN201510828187.9A priority patent/CN105627366A/zh
Publication of EP3026347A1 publication Critical patent/EP3026347A1/fr
Withdrawn legal-status Critical Current

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Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23M—CASINGS, LININGS, WALLS OR DOORS SPECIALLY ADAPTED FOR COMBUSTION CHAMBERS, e.g. FIREBRIDGES; DEVICES FOR DEFLECTING AIR, FLAMES OR COMBUSTION PRODUCTS IN COMBUSTION CHAMBERS; SAFETY ARRANGEMENTS SPECIALLY ADAPTED FOR COMBUSTION APPARATUS; DETAILS OF COMBUSTION CHAMBERS, NOT OTHERWISE PROVIDED FOR
    • F23M20/00—Details of combustion chambers, not otherwise provided for, e.g. means for storing heat from flames
    • F23M20/005—Noise absorbing 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
    • F23R3/02—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
    • F23R3/12—Air inlet arrangements for primary air inducing a vortex
    • F23R3/14—Air inlet arrangements for primary air inducing a vortex by using swirl vanes
    • 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
    • F23R3/002—Wall structures
    • 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
    • F23R3/286—Continuous combustion chambers using liquid or gaseous fuel characterised by the fuel supply having fuel-air premixing devices
    • 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
    • F23R3/343—Pilot flames, i.e. fuel nozzles or injectors using only a very small proportion of the total fuel to insure continuous 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
    • F23R3/42—Continuous combustion chambers using liquid or gaseous fuel characterised by the arrangement or form of the flame tubes or combustion chambers
    • F23R3/54—Reverse-flow combustion chambers
    • 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
    • F23R2900/00014—Reducing thermo-acoustic vibrations by passive means, e.g. by Helmholtz resonators

Definitions

  • the present invention relates generally to a system and method for improving combustion stability in a gas turbine combustor.
  • Diffusion type nozzles where fuel is mixed with air external to the fuel nozzle by diffusion, proximate the flame zone. Diffusion type nozzles have been known to produce high emissions due to the fact that the fuel and air burn stoichiometrically at high temperature to maintain adequate combustor stability and low combustion dynamics.
  • An enhancement in combustion technology is the utilization of premixing, such that the fuel and air mix prior to combustion to form a homogeneous mixture that burns at a lower temperature than a diffusion type flame and produces lower NOx emissions.
  • premixing fuel and air together before combustion allows for the fuel and air to form a more homogeneous mixture, which for a given combustor exit temperature will burn at lower peak temperatures, resulting in lower emissions.
  • Example of such a gas turbine flamesheet combustion system with reduced emissions and improved flame stability at multiple load conditions is disclosed in US patent application US2004/0211186A1
  • thermoacoustics of the flamesheet combustors could still lead to instability modes (such as pulsation), which could restrict the operation window.
  • aerodynamics of the burner allows occasional flame attachment in the mixing zone under certain circumstances, causing flashback and overheating risk.
  • current fuel staging strategies could cause asymmetrical heat load on the combustor liner, which could lead to creep problems.
  • measure which help against pulsation as for example the staging of 1/3-2/3 groups in the main fuel supply can lead to asymmetrical liner heat loading, as well as to non-uniformities in the combustor exit temperature profile.
  • a gas turbine combustor comprising a flow sleeve, a combustion liner located at least partially within the flow sleeve thereby creating a main passage between the flow sleeve and the combustor liner, and a dome located forward of the flow sleeve and encompassing at least a part of the combustion liner, the dome having a substantially rounded head end thereby forming a turning passage between the liner and the head end, and a swirler wall aligned along a centerline of the combustor, the swirler wall projecting into a space delimited by the liner, wherein the swirler wall and the rounded head end are connected, and wherein the connection forms an annular end face.
  • the combustor further comprises a center body positioned along the centerline and extending into the space delimited by the swirler wall, thereby forming a pilot passage between the swirler wall and the center body.
  • width of the pilot channel is substantially constant along the length of the pilot channel.
  • an area of the annular end face is 1.5 times to 5 times larger than an area of a cross section of the pilot passage.
  • a fuel lance is arranged in the center body.
  • the combustor further comprises a substantially cylindrical extension extending from a radially inner end of the rounded head end or the end face into the liner, wherein the extension is aligned with the centerline of the combustor.
  • the extension has substantially constant radius along the centerline of the liner, and/or the thickness of the extension is substantially equal to the thickness of the rounded head end.
  • a recess delimited by the central body, the annular end face and the rounded head end comprises a Helmholtz damper or/and means for pilot oil injection.
  • the combustion liner comprises a ring shaped rounded lip section and a curved middle section adapted to create a flame stabilization zone during operation.
  • the lip section comprises a Helmholtz damper and/or liquid fuel injection means.
  • the pilot passage comprises a pilot swirler in fluid communication with at least one pilot fuel injector, and the pilot swirler is an axial swirler or a radial swirler.
  • the main passage or the turning passage comprises a main swirler in a fluid communication with at least one main fuel injector, and wherein the main swirler is an axial swirler or a radial swirler.
  • the swirler wall is a part of a conical burner (e.g. EV burner or AEV burner).
  • the present application also provides for a gas turbine comprising the combustor described above.
  • the present application also provides for a method for operating the gas turbine combustor.
  • the method comprising: supplying a first stream of fuel into the pilot channel or conical burner (e.g. EV burner or AEV burner) to mix with the first flow of air, and feeding the resulting first mixture into the combustion zone for providing pilot flame; supplying a second flow of air into the main passage; supplying a second stream of fuel into the main passage or turning passage to mix with the second flow of air, and feeding the resulting second mixture into the combustion zone for providing a main flame.
  • the pilot channel or conical burner e.g. EV burner or AEV burner
  • FIG. 1 An example of a premixing flamesheet combustor 100 for a gas turbine of the prior art is shown in Fig. 1 .
  • the combustor 100 is a type of reverse flow premixing combustor utilizing a pilot nozzle 102, a radial inflow mixer 104, and a plurality of main stage mixers 108.
  • the pilot portion of the combustor 100 is separated from the main stage combustion area by a center divider portion 110.
  • the center divider portion 110 separates the fuel injected by the pilot nozzle 102 from the fuel injected by the main stage mixers 108.
  • the air entering through the main and the pilot burner is separated by the divider 110.
  • a flame front 120 which might occur for an off-design case, is shown schematically indicating interaction of pilot and main flame, which might cause thermoacoustic instabilities.
  • Fig. 2a shows a cross section view of a gas turbine combustor 200 in accordance with an embodiment of the present invention.
  • the combustor 200 comprising a flow sleeve 202, a combustion liner 204 located at least partially within the flow sleeve 202 thereby creating a main passage 206 between the flow sleeve 204 and the combustor liner 204.
  • the combustor 200 also comprises a dome 208 located forward of the flow sleeve and encompassing at least a part of the combustion liner 204.
  • the dome 208 has a substantially rounded head end 210 thereby forming a turning passage 212 between the liner 204 and the head end 210.
  • the compressor 200 comprises also a swirler wall 214 aligned along a centerline 216 of the combustor 200, wherein the swirler wall 214 is projecting into the liner 204.
  • the swirler wall 214 and the rounded head end 210 are connected, wherein the connection forms an annular end face 218.
  • the structure and thickness of the end face 218 can vary, and in one embodiment the end face 218 is a thin plate, for example a sheet metal plate. In one embodiment the end face 218 has a flat surface substantially perpendicular to the centerline 216. In one embodiment of the present invention, the end face 218 is cooled via effusion and/or impingement cooling.
  • the combustor 200 further comprises a center body 220 positioned along the centerline 216 and extending into the space delimited by the swirler wall 214.
  • the swirler wall 214 and the center body 220 form a pilot passage 222.
  • the center body comprises a front surface 226 which can have different shapes, depending on the combustor design, such as bluff body shape.
  • the width of the pilot channel 222 can vary, and preferably is substantially constant along the length of the pilot channel 222.
  • the center body 220 could also comprise a fuel lance 608 (shown in Fig 6b ) to create a central pilot flame.
  • Fig. 2b shows an end view of a gas turbine combustor in accordance with an embodiment of the present invention.
  • the cross sections of different components are shown as a generally cylindrical, but they can have other shapes such as oval or elongated.
  • An area of the annular end face 218 can vary in respect of the size of the other components of the combustor 200. In one preferred and non-limiting example, the area of the annular end face 218 is 1.5 times to 5 times larger than an area of a cross section of the pilot passage 222.
  • the combustor 200 according to the invention in one embodiment can comprise main fuel supply 234, pilot fuel supply 230, main swirler with injectors 232 and pilot swirler with injectors 228 to create a pilot flame and a main flame during an operation of the combustor.
  • Figure 2c shows schematically flame fronts, inside a combustion zone 250, created during operation of the combustor 200 according to the present invention. Contrary to the prior art ( Fig.1 ) where the pilot flame and the main flame interacts, in the embodiment according to the invention a main flame 260 and a pilot flame 262 are clearly separated due to the advantageous design of the combustor 200 according to the invention.
  • Fig. 3a shows a cross section view of a gas turbine combustor 200 in accordance with another embodiment of the present invention which further comprises a substantially cylindrical extension 240 extending from a radially inner end of the rounded head end 210 into the liner 220.
  • the extension 240 is extending from the end face 218.
  • the extension is substantially aligned with the centerline 216 of the combustor 200.
  • the extension 240 can vary in size, length, radius and width depending on operating parameters of the combustor 200.
  • the extension 240 is cylindrical and it has substantially constant radius along the centerline 216 of the liner.
  • the extension 240 and head end 210 have substantially same thickness.
  • the extension 240 and head end 210 could be made as two separate pieces or they can be made of a single piece of material. In one embodiment, the extension 240 and head end 210 are made of a sheet metal.
  • the cooling of the extension 240 may be done by near wall cooling using channels in axial direction.
  • Fig. 3b shows an end view of a gas turbine combustor in accordance with an embodiment of the present invention shown in Fig 3a .
  • an average thickness of extension 240 is smaller than average thickness of a cross section of the end face 218.
  • Fig. 4a shows a cross section view of a gas turbine combustor 200 in accordance with yet another embodiment according to the present invention wherein the combustion liner 204 comprises a ring shaped rounded lip section 420 and a curved middle section 430.
  • the liner 204 according to this embodiment could also comprise cooling holes 440.
  • the rounded lip section 420 is substantially hollow.
  • Fig. 4b shows an alternative embodiment, wherein the rounded lip section is made of thin material, substantially of the same thickness as the main portion of the liner 204, for example of a sheet metal. In this way, reducing the thickness of the rounded lip 420, there is advantageously more room for a stabilization zone.
  • Fig. 5 also shows a central pilot stabilization zone 530 and an outer pilot stabilization zone 520 created during operation of combustor 200 according to the invention.
  • the extension 420 advantageously makes possible effective separation of two pilot stabilization zones 520 and 530.
  • FIGS. 6a, 6b and 6c show additional embodiments of the present invention.
  • the lip section of the liner could comprise a Helmholtz damper 612 and/or liquid fuel injection means 606.
  • a recess 242 delimited by the central body 214, the annular end face 218 and the rounded head end 210 could comprises a Helmholtz damper 610 or/and a means for pilot oil injection 604.
  • Helmholtz damper is designed according to an individually determined or predetermined damping requirement against the thermoacoustic oscillation frequencies occurring in the combustion chamber.
  • the Helmholtz damper comprises a damper volume, a neck and a cooling channel.
  • the pilot swirlers (228, 618) and the main swirlers (232, 620) in general could be axial or radial swirlers.
  • the combustor 200 may comprise additional Helmholtz damper 602 and the fuel lance 608, both inside the center body 220, as shown in Fig. 6a and Fig. 6b .
  • the combustor 200 according to the invention could comprise a conical burner 702,704 device instead of the center body 220. Examples of these embodiments are shown in Fig. 7a and 7b , including EV burner (environmental burner from Alstom, disclosed in EP0321809 ) and AEV burner (advanced environmental burner from Alstom, disclosed in EP0704657 ) respectively.
  • the swirler wall 214 is a part of the conical burner 702,704.
  • Fig. 8a shows part of EV burner 702 wherein a conical column 5 of liquid fuel is formed in the interior 14 of the burner 702, which column widens in the direction of flow and is surrounded by a rotating stream 15 of combustion air which flows tangentially into the burner. Ignition of the mixture takes place at the burner outlet, a backflow zone 6 forming in the region of the burner outlet.
  • the burner itself consists of at least two hollow part-cone bodies 1, 2 which are superposed on one another and have a cone angle increasing in the direction of flow.
  • the part-cone bodies 1, 2 are mutually offset.
  • a nozzle 3 placed at the burner head ensures injection of the liquid fuel 2 into the interior 14 of the burner.
  • part cone body 1 of EV burner 702 corresponds to the swirler wall 212.
  • Fig. 8b shows part of AEV burner 704 comprising of at least part of the EV burner 702 and a mixing tube 802.
  • the mixing tube comprises a tube 804.
  • the tube 804 of AEV burner 704 corresponds to the swirler wall 212.
EP14194792.9A 2014-11-25 2014-11-25 Brûleur a corps non profilé annulaire Withdrawn EP3026347A1 (fr)

Priority Applications (3)

Application Number Priority Date Filing Date Title
EP14194792.9A EP3026347A1 (fr) 2014-11-25 2014-11-25 Brûleur a corps non profilé annulaire
US14/950,601 US20160146464A1 (en) 2014-11-25 2015-11-24 Combustor with annular bluff body
CN201510828187.9A CN105627366A (zh) 2014-11-25 2015-11-25 具有环状非流线型主体的燃烧器

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
EP14194792.9A EP3026347A1 (fr) 2014-11-25 2014-11-25 Brûleur a corps non profilé annulaire

Publications (1)

Publication Number Publication Date
EP3026347A1 true EP3026347A1 (fr) 2016-06-01

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EP14194792.9A Withdrawn EP3026347A1 (fr) 2014-11-25 2014-11-25 Brûleur a corps non profilé annulaire

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US (1) US20160146464A1 (fr)
EP (1) EP3026347A1 (fr)
CN (1) CN105627366A (fr)

Families Citing this family (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP3026346A1 (fr) * 2014-11-25 2016-06-01 Alstom Technology Ltd Chemise de chambre de combustion
EP3361161B1 (fr) * 2017-02-13 2023-06-07 Ansaldo Energia Switzerland AG Ensemble brûleur pour chambre de combustion d'une centrale électrique à turbine à gaz et chambre de combustion comprenant ledit ensemble brûleur
KR101889542B1 (ko) * 2017-04-18 2018-08-17 두산중공업 주식회사 연소기 노즐 조립체 및 이를 포함하는 가스터빈
US20190301408A1 (en) * 2018-04-02 2019-10-03 Caterpillar Inc. Combustion system for an internal combustion engine
JP7393262B2 (ja) * 2020-03-23 2023-12-06 三菱重工業株式会社 燃焼器、及びこれを備えるガスタービン
US11692709B2 (en) * 2021-03-11 2023-07-04 General Electric Company Gas turbine fuel mixer comprising a plurality of mini tubes for generating a fuel-air mixture
US11859819B2 (en) 2021-10-15 2024-01-02 General Electric Company Ceramic composite combustor dome and liners

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US4199935A (en) * 1975-11-28 1980-04-29 The Secretary Of State For Defence In Her Britannic Majesty's Government Of The United Kingdom Of Great Britain And Northern Ireland Combustion apparatus
EP0321809A1 (fr) 1987-12-21 1989-06-28 BBC Brown Boveri AG Procédé pour la combustion de combustible liquide dans un brûleur
EP0704657A2 (fr) 1994-10-01 1996-04-03 ABB Management AG Brûleur
US6430933B1 (en) * 1998-09-10 2002-08-13 Alstom Oscillation attenuation in combustors
US20040211186A1 (en) 2003-04-28 2004-10-28 Stuttaford Peter J. Flamesheet combustor
US20060260316A1 (en) * 2005-05-23 2006-11-23 Power Systems Mfg., Llc Flashback Suppression System for a Gas Turbine Combustor
US20140090389A1 (en) * 2012-10-01 2014-04-03 Peter John Stuttaford Variable length combustor dome extension for improved operability

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JP2544470B2 (ja) * 1989-02-03 1996-10-16 株式会社日立製作所 ガスタ―ビン燃焼器及びその運転方法
JPH06272862A (ja) * 1993-03-18 1994-09-27 Hitachi Ltd 燃料空気混合方法およびその混合装置
IT1273369B (it) * 1994-03-04 1997-07-08 Nuovo Pignone Spa Sistema perfezionato combustione a basse emissioni inquinanti per turbine a gas
US6986254B2 (en) * 2003-05-14 2006-01-17 Power Systems Mfg, Llc Method of operating a flamesheet combustor
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US7237384B2 (en) * 2005-01-26 2007-07-03 Peter Stuttaford Counter swirl shear mixer
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JP5172468B2 (ja) * 2008-05-23 2013-03-27 川崎重工業株式会社 燃焼装置および燃焼装置の制御方法
US9151500B2 (en) * 2012-03-15 2015-10-06 General Electric Company System for supplying a fuel and a working fluid through a liner to a combustion chamber
US10060630B2 (en) * 2012-10-01 2018-08-28 Ansaldo Energia Ip Uk Limited Flamesheet combustor contoured liner
EP3026346A1 (fr) * 2014-11-25 2016-06-01 Alstom Technology Ltd Chemise de chambre de combustion

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* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4199935A (en) * 1975-11-28 1980-04-29 The Secretary Of State For Defence In Her Britannic Majesty's Government Of The United Kingdom Of Great Britain And Northern Ireland Combustion apparatus
EP0321809A1 (fr) 1987-12-21 1989-06-28 BBC Brown Boveri AG Procédé pour la combustion de combustible liquide dans un brûleur
EP0704657A2 (fr) 1994-10-01 1996-04-03 ABB Management AG Brûleur
US6430933B1 (en) * 1998-09-10 2002-08-13 Alstom Oscillation attenuation in combustors
US20040211186A1 (en) 2003-04-28 2004-10-28 Stuttaford Peter J. Flamesheet combustor
US20060260316A1 (en) * 2005-05-23 2006-11-23 Power Systems Mfg., Llc Flashback Suppression System for a Gas Turbine Combustor
US20140090389A1 (en) * 2012-10-01 2014-04-03 Peter John Stuttaford Variable length combustor dome extension for improved operability

Also Published As

Publication number Publication date
CN105627366A (zh) 2016-06-01
US20160146464A1 (en) 2016-05-26

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