EP3026346A1 - Chemise de chambre de combustion - Google Patents

Chemise de chambre de combustion Download PDF

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Publication number
EP3026346A1
EP3026346A1 EP14194791.1A EP14194791A EP3026346A1 EP 3026346 A1 EP3026346 A1 EP 3026346A1 EP 14194791 A EP14194791 A EP 14194791A EP 3026346 A1 EP3026346 A1 EP 3026346A1
Authority
EP
European Patent Office
Prior art keywords
liner
section
combustor
combustion
flow
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Withdrawn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
EP14194791.1A
Other languages
German (de)
English (en)
Inventor
Adnan Eroglu
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 EP14194791.1A priority Critical patent/EP3026346A1/fr
Priority to US14/951,308 priority patent/US20160146467A1/en
Priority to CN201510831724.5A priority patent/CN105627363A/zh
Publication of EP3026346A1 publication Critical patent/EP3026346A1/fr
Withdrawn legal-status Critical Current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
    • F23R3/02—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
    • F23R3/42—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
    • F23C—METHODS OR APPARATUS FOR COMBUSTION USING FLUID FUEL OR SOLID FUEL SUSPENDED IN  A CARRIER GAS OR AIR 
    • F23C9/00—Combustion apparatus characterised by arrangements for returning combustion products or flue gases to the combustion chamber
    • F23C9/006—Combustion apparatus characterised by arrangements for returning combustion products or flue gases to the combustion chamber the recirculation taking place in the combustion chamber
    • 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
    • F23R3/44—Combustion chambers comprising a single tubular flame tube within a tubular casing
    • 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 to gas turbine engines, and more particularly relates to a combustor liner for improving combustion performance.
  • 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 emissions 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. Additionally, aerodynamics of the burner can lead to flame attachment in the mixing zone under certain circumstances, causing flashback and overheating risk. Furthermore, current fuel staging strategies could cause asymmetrical heat load on the combustor liner, which could lead to creep problems.
  • thermoacoustic instabilities which can enlarge the operation window available of the current combustor designs.
  • a combustor liner for a gas turbine having substantially cylindrical shape and comprising a first section and a second section wherein the first section is upstream of the second section with respect to the hot gas flow during operation, characterized in that the first section is ring shaped and comprises a rounded lip section and a trailing section, wherein an inner radius (R1) of the trailing section is increasing along a centerline of the liner in the direction of the hot gas flow during operation.
  • the radius (R1) of the trailing section is increasing monotonically along the centerline of the combustion liner.
  • the length in the axial direction of the first section is in the range from 20 percent to 80 percent of the total length in the axial direction of the liner.
  • an angle ( ⁇ ) between the trailing section and an outer surface of the combustion liner is in the range of 5 to 15 degrees.
  • a radius of outer surface of the combustion liner is substantially constant.
  • a radius (R2) of the second part is substantially constant along the centerline of the liner.
  • the first section of the combustor liner is substantially hollow.
  • An additional volume can be used for placement of at least one damper (preferably Helmholtz damper) and/or a means for a liquid fuel injection.
  • the present application also relates to a combustor comprising the liner described above and a combustion zone delimited by the combustion liner.
  • the combustor comprises a substantially cylindrical flow sleeve, wherein the combustion liner is located at least partially within the flow sleeve thereby forming a first passage between the flow sleeve and the combustion liner; a dome located forward of the flow sleeve and encompassing at least partially a first section of the combustion liner, the dome having a substantially rounded head end thereby forming a turning passage between the rounded lip section of the first section of combustion liner and the dome ; and at least one pilot channel comprising a means for supplying a pilot fuel and a first swirling device.
  • the turning passage can for example have a cross section shaped like half annulus. The turning passage extends from the first passage into combustion zone and guides cooling air leaving the first passage around the upstream end of the first section of combustion liner into the combustion zone of
  • the first passage and/or the turning passage comprise a fuel injection means and a second swirling device.
  • the first swirling device and/or the second swirling device are axial or radial swirlers.
  • 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 comprises: supplying a first flow of air into the pilot channel ;supplying a first stream of fuel into the pilot channel 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 first passage ;supplying a second stream of fuel into the first passage or second passage to mix with the second flow of air , and feeding the resulting second mixture into the combustion zone for providing a main flame; wherein the first mixture and second mixture are guided along the inner wall of the liner and form a central recirculation zone in the center of the combustion zone.
  • FIG. 1 An example of a premixing flamesheet combustor 100 for a gas turbine of the prior art is shown in Fig. 1 .
  • the combustion system 100 includes a flow sleeve 102 containing a combustion liner 104.
  • the combustion liner 104 has a constant radius along the centreline AA' of the combustor 100.
  • a fuel injector 106 is secured to a casing 108 with the casing 108 encapsulating a radial mixer 110.
  • Secured to the forward portion of the casing 108 are a cover 112 and pilot nozzle assembly 114.
  • the combustor 100 is a type of reverse flow premixing combustor
  • Fig.2 shows cross section of a central portion of a flamesheet combustor 100 during an operation.
  • the fuel is provided to the combustor 100 via fuel injection nozzles 106 (main fuel) and 114 (pilot fuel).
  • the air is mixed with pilot fuel and main fuel respectively.
  • the radial mixer 110 provides swirled air to the fuel-air mixture to improve flame stabilization. Use of the mixer 110 stabilizes the combustion process by developing a reverse flow inside the combustor 100. The reverse flow returns free radicals and heat upstream to the unburnt air-fuel mixture. In this way, two separate recirculation zones, a central recirculation zone 210 and an outer recirculation zone 220 are created as shown in Fig. 2 .
  • the flame is anchored in the central recirculation zone 210 at ignition and part-load conditions with the help of pilot fuel. At higher loads, the flame is transferred to outer recirculation zone 220 by increasing supply of main fuel.
  • FIG. 3a shows a cross section view of a combustion liner 300 for a gas turbine in accordance with an embodiment of the present invention.
  • the combustor liner 300 has substantially cylindrical shape and comprises a first section 310 and a second section 320 wherein the first section is upstream of the second section with respect to the hot gas flow during operation.
  • the first section 310 is ring shaped and comprises a rounded lip section 330 and a trailing section 340.
  • An inner radius (R1) of the trailing section 340 is increasing along a centerline 350 of the liner 300 in the direction of the hot gas flow during operation.
  • the radius (R1) of the trailing section 340 is increasing monotonically along the centerline 350 of the liner 300. This means, for example, that the trailing section 340 can have at least one flat region with the constant radius (R1).
  • the length, in axial direction, of the first section 310 in respect to the total length, in axial direction, of the liner 300 can vary. In one preferred embodiment, the length of the first section 310 is in the range from 20 percent to 80 percent of the total length of the liner 300. As shown in Fig. 3a , there is an angle ( ⁇ ) between an outer surface 360 of the combustion liner 300 and the trailing section 340. The angle ( ⁇ ) can vary. In one preferred embodiment the angle ( ⁇ ) is in the range of 5 to 15 degrees.
  • the radius of the outer surface 360 of the combustion liner 300 is substantially constant along the centerline 350 of the liner 300. This means that the outer radius of the section 310 and the section 320 are substantially equal.
  • a radius (R2) of the second section 320 is substantially constant along the centerline 350 of the liner 300.
  • the radius (R1) and radius (R2) are equal at least at a point of connection between the first section 310 and the second section 320.
  • Fig 3b shows another embodiment of the combustor liner 300 according to the invention. Contrary from the first embodiment ( Fig. 3a ), where the trailing section radius (R1) is increasing smoothly towards second section, in this embodiment there is a sharp step-like increase in the radius (R1) of the trailing section 340. In one embodiment the step occurs after the radius (R1) already increased for at least 10 percent.
  • Combustion liners from the prior art (such as shown in Fig.1 ) have substantially cylindrical shape with constant radius along liner's centreline.
  • the combustor liners are made of thin metal sheets. Due to the low thickness of the walls, such liners have no possibilities to incorporate additional devices in the liner structures.
  • One of the features of the combustor liners according to the invention is that the first section 310 of combustion liner 300 is substantially hollow, while the second section 320 is made of thin material, normally of sheet metal.
  • the additional space inside the first section 310 can be advantageous comparing to liners from the prior art. In one embodiment according to the invention, this additional space inside the first section 310 can be used for placing a damper device.
  • Fig. 3c shows the combustor liner 300 wherein the first section 310 comprises a Helmholtz damper 370.
  • 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 370 comprises a damper volume, a neck 371 and a cooling channel 372.
  • the space inside the first section 310 is used to incorporate the means for liquid fuel injection 380.
  • the one example of such a means for liquid fuel injection is fuel nozzles.
  • the combustor liner 300 can be incorporated in a combustion system of a gas turbine.
  • Fig. 4 shows a combustor 400 for a gas turbine according to the invention, comprising the combustor liner 300 and a combustion zone 401 delimited by the combustion liner 300.
  • the combustor comprises a substantially cylindrical flow sleeve 410, wherein the combustion liner 300 is located at least partially within the flow sleeve 410.
  • the flow sleeve 410 and the combustion liner 300 form a first passage 420.
  • the combustor 400 further comprises a dome 425 located forward of the flow sleeve and encompassing at least partially a first section 310 of the combustion liner 300.
  • the dome 425 has a substantially round head 430 thereby forming a turning passage 440 between the rounded lip section 330 of the first section 310 of combustion liner and the dome 425.
  • the combustor comprises at least one pilot channel 455 comprising a means for supplying a pilot fuel 460 and a first swirling device 495.
  • the first passage 420 further comprises a main fuel injection means 450 and a second swirling device 490.
  • the combustor 400 comprises a bluff body 402 to stabilize a flame inside the combustion zone 401.
  • the bluff body 402 could contain additional fuel nozzles.
  • outer and central recirculation zones are created (as shown in Fig.2 ).
  • Fig.4 during the operation of the combustor 400, according to the invention, only the central recirculation zone 405 is created.
  • the outer recirculation zone is not present due to the design of the combustor liner 300 according to the invention. Elimination of outer recirculation zone removes the problems of bi-stable flame. Flame is stabilized through the central recirculation zone. There is neither competition nor transfer from one zone to another.
  • Fig. 5 shows the combustor 400 comprising the combustor liner 300 with the Helmholtz damper 370. This embodiment offers additional acoustic damping possibilities for the combustor 400.
  • Fig. 6 shows the combustor liner 400 comprising the means for liquid fuel injection 380. This embodiment offers additional liquid fuel supply possibilities for the combustor 400.
  • Fig. 7 shows another embodiment of the combustor according to the invention.
  • a radial staging means 710 are positioned in the turning passage 440, preferably downstream of the dome 425.
  • the radial staging means 710 comprises at least one and preferably two separated parts, an inner part and an outer part.
  • Inner part comprises an inner main swirler 712 while the outer part comprises an outer swirler 711.
  • the swirlers 711 and 712 are supplied with fuel from a fuel injector 721, which is preferably positioned downstream the turning passage 440.
  • This staging configuration can prevent flame attachment problems completely and enable smooth loading of the combustor by increasing the fuel radially from inside to outside gradually.
  • the present invention also provides a method for operating the gas turbine combustor 400 according to the invention.
  • the method comprises the steps: supplying a first flow of air 480 into the pilot channel 455;supplying a first stream of fuel into the pilot channel 455 to mix with the first flow of air 480, and feeding the resulting first mixture into the combustion zone 401 for providing pilot flame; supplying a second flow of air 470 into the first passage 420;supplying a second stream of fuel into the first passage 420 or turning passage 440 to mix with the second flow of air 470, and feeding the resulting second mixture into the combustion zone 401 for providing a main flame; wherein the first mixture and second mixture are guided along the inner wall of the liner and form a central recirculation zone 405 in the center of the combustion zone 401.
  • the first flow of air 480 and the second flow of air 470 are normally supplied from a compressor plenum (not shown).
  • the main advantages of the present invention are improved stability due to single recirculation zone, thus elimination of competition between inner and outer recirculation zones and loading the combustor without any flame transfer from inside to outside.
  • the flame is always anchored in the centre as the fuel added to outer layers as increased load.
  • Additional advantages of the present application in addition to improved stability, are: reduced heat load to liner at part load due to cooler outer streams (liner loading is high only at peak loads);uniform heat load to liner, preventing creep and deformation; more uniform combustor exit temperature distribution; creation of additional volume for acoustic damping and dual-fuel injection(liquid fuel); elimination of flame-holding and flashback risk by moving the main premix injection downstream of bend.

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  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
EP14194791.1A 2014-11-25 2014-11-25 Chemise de chambre de combustion Withdrawn EP3026346A1 (fr)

Priority Applications (3)

Application Number Priority Date Filing Date Title
EP14194791.1A EP3026346A1 (fr) 2014-11-25 2014-11-25 Chemise de chambre de combustion
US14/951,308 US20160146467A1 (en) 2014-11-25 2015-11-24 Combustor liner
CN201510831724.5A CN105627363A (zh) 2014-11-25 2015-11-25 燃烧器衬套

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
EP14194791.1A EP3026346A1 (fr) 2014-11-25 2014-11-25 Chemise de chambre de combustion

Publications (1)

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EP3026346A1 true EP3026346A1 (fr) 2016-06-01

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EP14194791.1A Withdrawn EP3026346A1 (fr) 2014-11-25 2014-11-25 Chemise de chambre de combustion

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US (1) US20160146467A1 (fr)
EP (1) EP3026346A1 (fr)
CN (1) CN105627363A (fr)

Families Citing this family (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
JP6516996B2 (ja) * 2014-10-10 2019-05-22 川崎重工業株式会社 燃焼器及びガスタービンエンジン
EP3026347A1 (fr) * 2014-11-25 2016-06-01 Alstom Technology Ltd Brûleur a corps non profilé annulaire
CN114909675B (zh) * 2022-04-07 2024-03-01 中国联合重型燃气轮机技术有限公司 用于燃气涡轮机的燃烧室和燃气轮机
CN114877371B (zh) * 2022-05-06 2023-03-31 南京航空航天大学 一种具有双重稳定火焰机制的先进燃烧室及其燃烧方法

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EP2362147A1 (fr) * 2010-02-22 2011-08-31 Alstom Technology Ltd Dispositif de combustion pour turbine à gaz
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CN105627363A (zh) 2016-06-01
US20160146467A1 (en) 2016-05-26

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