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EP0924470B1 - Chambre de combustion à prémélange pour turbine à gaz - Google Patents

Chambre de combustion à prémélange pour turbine à gaz Download PDF

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Publication number
EP0924470B1
EP0924470B1 EP98123199A EP98123199A EP0924470B1 EP 0924470 B1 EP0924470 B1 EP 0924470B1 EP 98123199 A EP98123199 A EP 98123199A EP 98123199 A EP98123199 A EP 98123199A EP 0924470 B1 EP0924470 B1 EP 0924470B1
Authority
EP
European Patent Office
Prior art keywords
combustor
mixing
zone
pilot
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.)
Expired - Lifetime
Application number
EP98123199A
Other languages
German (de)
English (en)
Other versions
EP0924470A2 (fr
EP0924470A3 (fr
Inventor
Nikolaos Dr. Zarzalis
Thomas Ripplinger
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.)
MTU Aero Engines AG
Original Assignee
MTU Aero Engines GmbH
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
Priority claimed from DE1997156663 external-priority patent/DE19756663B4/de
Priority claimed from DE1998110648 external-priority patent/DE19810648A1/de
Application filed by MTU Aero Engines GmbH filed Critical MTU Aero Engines GmbH
Publication of EP0924470A2 publication Critical patent/EP0924470A2/fr
Publication of EP0924470A3 publication Critical patent/EP0924470A3/fr
Application granted granted Critical
Publication of EP0924470B1 publication Critical patent/EP0924470B1/fr
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
    • 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
    • F23R3/28—Continuous combustion chambers using liquid or gaseous fuel characterised by the fuel supply

Definitions

  • the invention relates to a premixing combustion chamber for a gas turbine, comprising a main stage with at least one premixing chamber and at least one for Part of the combustion chamber, which is designed to be rotationally symmetrical with respect to its longitudinal axis, with an i Main combustion zone and a downstream post-combustion zone, the at least one premixing chamber generating tangential swirl into the combustion chamber opens in the area of the main combustion zone; and a pilot stage with a pilot injector.
  • Premix combustion chambers are low-pollution gas turbine combustion chambers.
  • gas turbines can be stationary, e.g. as generator drives in power plants, as well as in Aircraft engines are used.
  • nitrogen oxide emissions from stationary gas turbines There are maximum limits in numerous industrialized countries for nitrogen oxide emissions from stationary gas turbines.
  • There Reduction also comes with corresponding recommendations for aircraft engines the stickodix formation in the combustion chambers as part of the lowering of the Pollutant emissions are of great importance.
  • fat-lean burn is currently used in the combustion with a first rich level and a second lean level with excess air he follows.
  • the hot gases from the pilot zone become the lean Main zone mixed in, the stabilizing effect strongly of the existing n-den Flow field depends and larger in different operating conditions Can be subject to fluctuations.
  • the flow from the main in the post-combustion zone is redirected by 90 °, which leads to an increased pressure loss leads.
  • the object of the invention is to provide a premixing combustor described genus to create, in which the stabilizing effect of pilot combustion is improved.
  • the solution to this problem is characterized in that the Main combustion zone in the combustion chamber essentially coaxial or parallel runs or is arranged to the afterburning zone, i.e. the flow path in runs essentially straight and without significant deflection, and the pilot stage the end of the combustion chamber remote from the afterburning zone is arranged.
  • premix combustion chamber The advantage of this premix combustion chamber is that the flow is within the combustion chamber from the main combustion zone to the afterburning zone is deflected by 90 ° and the associated pressure loss is eliminated.
  • the pilot stage arranged directly on the combustion chamber has a direct one Connection to the main combustion or recirculation zone, creating the stabilizing Effect of pilot combustion is significantly improved.
  • Premix combustors can be used in stationary gas turbines as well as in aircraft engines deploy.
  • the main combustion zone widens forming area of the combustion chamber in the flow direction, from the Main combustion zone runs towards the post-combustion zone, conical on. Through the opening angle of the cone, the recirculation zone and thus control flame stability. While there is an additional one at smaller opening angles Pre-evaporation results in the case of larger opening angles Promotes combustion stability.
  • the pilot stage at the end of the combustion chamber with a smaller radius is preferred arranged at the front and coaxial to it.
  • pilot stage may be one between the pilot injection device and the combustion chamber arranged pilot combustion chamber.
  • the premix combustion chamber 1 shows an exemplary embodiment of a premix combustion chamber, designated as a whole by 1 for a gas turbine.
  • the premix combustion chamber 1 essentially comprises a main stage 2 with a premixing chamber 6, a main combustion zone 3 and a post-combustion zone 5 and a pilot stage 4.
  • the premixing chamber 6 becomes the fuel together with part of the compressor air brought in.
  • the fuel is atomized and evaporated in the premixing chamber 6 and mixed with the air as homogeneously as possible.
  • the premixing chamber 6 is formed as a rectilinear rectangular channel, so that within the premixing chamber 6 a swirl-free flow with a relatively uniform velocity profile is produced.
  • the premixing chamber 6 can depending on the machine design also have other suitable cross-sectional shapes, such as e.g. oval or circular.
  • the cross-sectional shape does not necessarily have to be constant over the length of the premixing chamber 6.
  • the fuel-air mixture flows at an outlet end 8 of the premixing chamber 6 into the combustion chamber 9, which is designed as a truncated cone, in the region of the Main combustion zone 3 lying part and a cylindrical, in the area of Afterburning zone 5 includes lying part 12.
  • the flow is included the greatest possible eccentricity to a longitudinal or central axis M of the rotationally symmetrical Combustion chamber 9 introduced so that the flow of the Fuel / air mixture is impressed a peripheral speed.
  • Premixing chamber 6 is also designed with the lowest possible height H.
  • the combustion chamber 9 has a plurality of for cooling Air inlet openings.
  • pilot stage 4 At an end 10 of the combustion chamber 9 remote from the afterburning zone 5 the pilot stage 4 arranged.
  • pilot level is 4 thus at the front end 10 with the smallest radius as a truncated cone trained part of the combustion chamber 9 arranged.
  • the pilot stage 4 includes one Pilot injector 11, with the fuel in the main combustion zone 3 for Stabilization of the combustion can be introduced in particular in the partial load range can.
  • the hot gases from pilot stage 4 flow directly into the core of the recirculation zone the lean main level 2, resulting in improved stability of the Combustion leads. Both in the main and in the pilot level 2 or 4 gaseous and liquid fuels are used.
  • Fig. 2 shows another embodiment of the premix combustion chamber 1, the Modification is in the area of pilot level 4.
  • the pilot stage 4 additionally to the pilot injection device 11, a pilot combustion chamber 13 in which the Fuel is first mixed with air in a diffusion combustion and only then is then introduced into the front of the combustion chamber 9.
  • Fig. 3 shows an arrangement in which a plurality of premixing combustion chambers 1 with an annular combustion chamber 14 are combined.
  • the individual premixing combustion chambers comprise 1 a premixing chamber 6, which is eccentrically into a truncated cone trained part of the combustion chamber 9 opens a main stage 2, and an afterburning zone arranged essentially coaxially to the main stage 2 5, causing the flow between the main combustion zone 3 and the post-combustion zone 5 does not have to be deflected and therefore the combustion chamber pressure drop is reduced.
  • the combustion chamber 9 could also be cylindrical Have part 12 which is substantially coaxial to the longitudinal axis M of the combustion chamber 9 is arranged.
  • the annular combustion chamber 14 When the annular combustion chamber 14 is installed in a gas turbine, it becomes with its central axis M arranged coaxially with it and from an upstream one Air supplied to the compressor on the injection side.
  • the premix combustion chambers 1 are arranged equidistantly around the end circumference of the annular combustion chamber 14.
  • the wall of the combustion chamber 9 is for cooling with air inlet openings Mistake.
  • the main stage 2 and the pilot stage can 4 Depending on the load or flight phase, they can be operated separately or simultaneously.

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  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)

Claims (10)

  1. Chambre de combustion à prémélange pour une turbine à gaz, comprenant un étage principal avec au moins une chambre de prémélange et une chambre de combustion qui est réalisée au moins en partie avec une symétrie de rotation autour de son axe longitudinal, et qui comporte une zone de combustion principale et une zone de postcombustion située en aval, ladite au moins une chambre de prémélange débouchant tangentiellement dans la chambre de combustion en engendrant un tourbillon giratoire, l'ensemble comprenant en outre un étage pilote avec un dispositif d'injection pilote,
    caractérisée en ce que la zone de combustion principale (3), dans la chambre de combustion (9), s'étend sensiblement de manière coaxiale à la zone de postcombustion (5), et l'étage pilote (4) est disposé à l'extrémité (10) de la chambre de combustion (9), qui est éloignée de la zone de postcombustion (5).
  2. Chambre de combustion à prémélange selon la revendication 1, caractérisée en ce que ladite au moins une chambre de prémélange (6) est réalisée sous forme de canal rectangulaire.
  3. Chambre de combustion à prémélange selon la revendication 1 ou 2, caractérisée en ce que la hauteur (H) de ladite au moins une chambre de prémélange (6) est faible par rapport à sa longueur et à sa largeur.
  4. Chambre de combustion à prémélange selon l'une ou plusieurs des revendications précédentes, caractérisée en ce qu'une extrémité de sortie (8) de ladite au moins une chambre de prémélange (6) est disposée de manière telle, par rapport à la chambre de combustion (9), que l'écoulement pénétrant dans la chambre de combustion (9) présente une excentricité maximale par rapport à l'axe longitudinal (M) de la chambre de combustion (9).
  5. Chambre de combustion à prémélange selon l'une ou plusieurs des revendications précédentes, caractérisée en ce que deux ou quatre chambres de prémélange (6) débouchent respectivement par paire dans la chambre de combustion (9), en des endroits au moins approximativement diamétralement opposés, en produisant une formation de tourbillon giratoire de même sens.
  6. Chambre de combustion à prémélange selon l'une ou plusieurs des revendications précédentes, caractérisée en ce que la zone de la chambre de combustion (9) entourant la zone de combustion principale (3), s'évase de manière conique dans le sens d'écoulement.
  7. Chambre de combustion à prémélange selon la revendication 6, caractérisée en ce que l'étage pilote (4) est disposé de manière frontale à l'extrémité (10) de rayon le plus faible de la chambre de combustion (9), et s'étend de manière coaxiale à celle-ci.
  8. Chambre de combustion à prémélange selon l'une ou plusieurs des revendications précédentes, caractérisée en ce que l'étage pilote (4) comporte une chambre de combustion pilote (13) disposée entre le dispositif d'injection pilote (11) et la chambre de combustion (9).
  9. Agencement de chambre de combustion à prémélange selon l'une ou plusieurs des revendications précédentes, caractérisée en ce que la partie de la chambre de combustion (9) formant la zone de postcombustion (5), est réalisée sous forme de chambre de combustion annulaire (14), sur la face frontale de laquelle sont raccordées, de manière équidistante, un grand nombre de chambres de combustion (9) englobant la zone de combustion principale (3), y compris les chambres de prémélange (6) et les étages pilotes (4).
  10. Agencement de chambre de combustion à prémélange selon la revendication 9, caractérisée en ce que chaque chambre de combustion (9) comprend une partie conique entourant sensiblement la zone de combustion principale (3), et en aval de celle-ci, une partie cylindrique (12) disposée coaxialement à son axe longitudinal (M) et débouchant dans la chambre de combustion annulaire (14).
EP98123199A 1997-12-19 1998-12-05 Chambre de combustion à prémélange pour turbine à gaz Expired - Lifetime EP0924470B1 (fr)

Applications Claiming Priority (4)

Application Number Priority Date Filing Date Title
DE19756663 1997-12-19
DE1997156663 DE19756663B4 (de) 1997-12-19 1997-12-19 Vormischbrennkammer für eine Gasturbine
DE1998110648 DE19810648A1 (de) 1998-03-12 1998-03-12 Vormischbrennkammer für eine Gasturbine
DE19810648 1998-03-12

Publications (3)

Publication Number Publication Date
EP0924470A2 EP0924470A2 (fr) 1999-06-23
EP0924470A3 EP0924470A3 (fr) 2001-03-14
EP0924470B1 true EP0924470B1 (fr) 2003-06-18

Family

ID=26042638

Family Applications (1)

Application Number Title Priority Date Filing Date
EP98123199A Expired - Lifetime EP0924470B1 (fr) 1997-12-19 1998-12-05 Chambre de combustion à prémélange pour turbine à gaz

Country Status (4)

Country Link
US (1) US6202420B1 (fr)
EP (1) EP0924470B1 (fr)
JP (1) JPH11248159A (fr)
DE (1) DE59808754D1 (fr)

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CN100443806C (zh) * 2006-05-16 2008-12-17 北京航空航天大学 切向驻涡燃烧室
EP2299178B1 (fr) 2009-09-17 2015-11-04 Alstom Technology Ltd Procédé et système de combustion de turbine à gaz pour mélanger sans danger des carburants riches en H2 avec de l'air
CN102032597B (zh) * 2010-11-29 2012-07-04 北京航空航天大学 一种离散管主燃级的预混预蒸发燃烧室
US8978388B2 (en) 2011-06-03 2015-03-17 General Electric Company Load member for transition duct in turbine system
US8650852B2 (en) 2011-07-05 2014-02-18 General Electric Company Support assembly for transition duct in turbine system
US8448450B2 (en) 2011-07-05 2013-05-28 General Electric Company Support assembly for transition duct in turbine system
US9328623B2 (en) * 2011-10-05 2016-05-03 General Electric Company Turbine system
US8701415B2 (en) 2011-11-09 2014-04-22 General Electric Company Flexible metallic seal for transition duct in turbine system
US8459041B2 (en) 2011-11-09 2013-06-11 General Electric Company Leaf seal for transition duct in turbine system
US8974179B2 (en) 2011-11-09 2015-03-10 General Electric Company Convolution seal for transition duct in turbine system
CN102393028B (zh) * 2011-12-09 2013-08-28 中国船舶重工集团公司第七�三研究所 天然气燃料燃气轮机干式低排放燃烧室
US9133722B2 (en) 2012-04-30 2015-09-15 General Electric Company Transition duct with late injection in turbine system
US9038394B2 (en) 2012-04-30 2015-05-26 General Electric Company Convolution seal for transition duct in turbine system
US8707673B1 (en) 2013-01-04 2014-04-29 General Electric Company Articulated transition duct in turbomachine
US9080447B2 (en) 2013-03-21 2015-07-14 General Electric Company Transition duct with divided upstream and downstream portions
CN103266922B (zh) * 2013-06-15 2014-11-12 厦门大学 一种带有级间燃烧室的涡轮静子叶片
US9458732B2 (en) 2013-10-25 2016-10-04 General Electric Company Transition duct assembly with modified trailing edge in turbine system
US10145251B2 (en) 2016-03-24 2018-12-04 General Electric Company Transition duct assembly
US10260752B2 (en) 2016-03-24 2019-04-16 General Electric Company Transition duct assembly with late injection features
US10227883B2 (en) 2016-03-24 2019-03-12 General Electric Company Transition duct assembly
US10260424B2 (en) 2016-03-24 2019-04-16 General Electric Company Transition duct assembly with late injection features
US10260360B2 (en) 2016-03-24 2019-04-16 General Electric Company Transition duct assembly
WO2019165385A1 (fr) * 2018-02-23 2019-08-29 Fulton Group N.A., Inc. Brûleur à combustion de combustible à prémélange à allumage vers l'intérieur
CN109113895B (zh) * 2018-09-11 2019-08-27 中国人民解放军国防科技大学 冲压发动机火焰稳定装置

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Also Published As

Publication number Publication date
DE59808754D1 (de) 2003-07-24
JPH11248159A (ja) 1999-09-14
EP0924470A2 (fr) 1999-06-23
US6202420B1 (en) 2001-03-20
EP0924470A3 (fr) 2001-03-14

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