EP0781967A2 - Chambre de combustion annulaire pour turbine à gaz - Google Patents

Chambre de combustion annulaire pour turbine à gaz Download PDF

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Publication number
EP0781967A2
EP0781967A2 EP96810777A EP96810777A EP0781967A2 EP 0781967 A2 EP0781967 A2 EP 0781967A2 EP 96810777 A EP96810777 A EP 96810777A EP 96810777 A EP96810777 A EP 96810777A EP 0781967 A2 EP0781967 A2 EP 0781967A2
Authority
EP
European Patent Office
Prior art keywords
combustion chamber
gas turbine
air
compressor
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.)
Granted
Application number
EP96810777A
Other languages
German (de)
English (en)
Other versions
EP0781967B1 (fr
EP0781967A3 (fr
Inventor
Klaus Dr. Döbbeling
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.)
General Electric Switzerland GmbH
Original Assignee
ABB Research Ltd 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 ABB Research Ltd Switzerland filed Critical ABB Research Ltd Switzerland
Publication of EP0781967A2 publication Critical patent/EP0781967A2/fr
Publication of EP0781967A3 publication Critical patent/EP0781967A3/fr
Application granted granted Critical
Publication of EP0781967B1 publication Critical patent/EP0781967B1/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
    • F23R3/02—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
    • F23R3/02—Continuous combustion chambers using liquid or gaseous fuel characterised by the air-flow or gas-flow configuration
    • F23R3/04—Air inlet arrangements
    • F23R3/045—Air inlet arrangements using pipes
    • 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
    • F23R3/42—Continuous combustion chambers using liquid or gaseous fuel characterised by the arrangement or form of the flame tubes or combustion chambers
    • F23R3/50—Combustion chambers comprising an annular flame tube within an annular 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
    • F23R2900/03041—Effusion cooled combustion chamber walls or domes

Definitions

  • the invention relates to the field of combustion technology. It relates to a gas turbine ring combustion chamber, which is operated with premix burners, and to a method for operating this device.
  • Gas turbines essentially consist of the components compressor, combustion chamber and turbine. For reasons of environmental protection, instead of diffusion combustion, low-pollutant premix combustion is increasingly used.
  • the air emerging from the compressor has a very high speed (approx. 200 m / s) and, in order to recover the kinetic energy contained in it, is decelerated with as little loss as possible in a deflection diffuser.
  • the speed in the combustion chamber is again greatly reduced, at least locally downstream of the burner. Usually a local recirculation zone with negative velocities is created.
  • the speed in the combustion chamber is then about 50 m / s in order to obtain a sufficient dwell time and to keep the heat transfer between the hot gas and the combustion chamber wall low.
  • At the outlet of the combustion chamber there is an acceleration so that gas velocities close to the speed of sound are reached at the turbine inlet.
  • the invention tries to avoid all these disadvantages. It is based on the task of developing a gas turbine ring combustion chamber which is equipped with special premix burners and which is distinguished by a small size and is simplified compared to the known prior art, with improved premixing of fuel and air with a lower total pressure drop.
  • a gas turbine ring combustion chamber which is arranged downstream of a compressor and is equipped on its front plate with at least one ring-shaped premix burner row, in that a burner air duct designed as a diffuser leads to each burner directly downstream of the compressor outlet from the guide vanes of the last compressor row.
  • a burner air duct designed as a diffuser leads to each burner directly downstream of the compressor outlet from the guide vanes of the last compressor row.
  • the combustion air is divided into individual air streams for the burners and for cooling the combustion chamber and turbine immediately after leaving the compressor, after which the speed of the air for the burners is reduced to approximately that Half the value of the compressor outlet speed is delayed, then at least one longitudinal vortex is generated in the air for each combustion air duct, fuel being added during or downstream of the generation of the longitudinal vortex, the mixture now flows along in a mixing duct and flows with a total swirl into the combustion chamber and finally burns there.
  • the advantages of the invention include the fact that the combustion chamber has smaller dimensions compared to the prior art and the area to be cooled in the combustion chamber is reduced. The pressure loss between the compressor outlet and the turbine inlet is smaller. In addition, there is a very good and robust uniform distribution of the air to the burners and the premixing of fuel and combustion air is improved.
  • the ratio of the number of blades in the last compressor row to the number of premix burners is an integer, in particular 1 or 2, because then a combustion air channel can be coupled directly to one or two blade channels in the last compressor row.
  • the mixing channel has an approximately round cross section, because then a good mixing of air and fuel is achieved. Mixing channels with a rectangular cross section are also conceivable. Likewise, if there is only one burner row, the mixing channel can be designed as a segmented annular gap.
  • combustion air ducts are arranged spirally around the axis of the gas turbine. In this way, axial length can be saved.
  • the axes of the mixing channels are advantageous Mixture), arranged so that they form an angle, preferably an angle of 45 °, with the axis of the gas turbine. This further improves mixing and flame stabilization.
  • the system does not show, for example, the exhaust gas casing of the gas turbine with exhaust pipe and chimney, as well as the inlet parts of the compressor part and the low-pressure compressor stages.
  • the direction of flow of the work equipment is indicated by arrows.
  • Fig. 1 shows a partial longitudinal section of a gas turbine system with an annular combustion chamber according to the prior art.
  • An annular combustion chamber 4 which is equipped with premix burners 5 of the double-cone type, is arranged between a compressor 1 and a turbine 2, of which only one guide vane 3 of the first row of guide vanes is shown.
  • the fuel 6 is supplied to each premix burner 5 via fuel lances 7.
  • the annular combustion chamber 4 is cooled convectively or by means of impingement cooling.
  • the compressor 1 consists essentially of the blade carrier 8, in which the guide blades 9 are suspended, and of the rotor 10, which receives the rotor blades 11. In Fig. 1 are each only the last compressor stages are shown.
  • a deflection diffuser 12 is arranged at the outlet of the compressor 1. It opens into a plenum 13 arranged between the compressor 1 and the annular combustion chamber 4.
  • the air 14 emerging from the compressor 1 has a very high speed. It is delayed in the deflection diffuser 12 in order to recover the kinetic energy contained therein, so that only very low air velocities prevail in the plenum 13 adjoining the deflection diffuser 12. As a result, a uniform distribution of the air 14 to the burners 5 can be achieved and cooling air for the combustion chamber 4 and the turbine 2 can be removed without problems.
  • the speed since the speed must be high in order to avoid flame flashback in order to design the premixing process of air 14 and fuel 6 at the mixing point of the fuel 6 in a reliable manner, the air 14 in the premixing zone must be accelerated again strongly before the burner 5 downstream again in the Combustion chamber 4, for reasons of flame stability, the speed is reduced.
  • the gas is then accelerated again at the downstream end of the combustion chamber 4, so that 2 speeds close to the speed of sound are reached at the inlet into the turbine.
  • the multiple accelerations and decelerations between the compressor outlet and the turbine inlet are fraught with losses and the required multiple deflections of the air mass flow lead to a rather large overall height.
  • the outer diameter in the region of the combustion chamber is approximately 4.5 m.
  • FIG. 2 shows an embodiment of the invention using a four-row gas turbine ring combustion chamber.
  • the air 14 is no longer delayed to plenum conditions, but the delay in the air 14 is only limited to that Speed level of the premixing section. This eliminates the multiple redirection of the total air mass flow and the size in the area of the combustion chamber can be significantly reduced.
  • a burner air distribution system is arranged directly downstream of the compressor outlet on the guide vanes 9 of the last compressor blade row, in which a burner air duct 15 designed as a diffuser leads to each burner 5 of the annular combustion chamber 4.
  • At least one longitudinal vortex generator 16 is located at the downstream end of the combustion air duct 15.
  • At least one fuel injector 17 is provided in or downstream of the longitudinal vortex generator 16, and downstream of the fuel injector 17 is a mixing duct 19 of constant height H and with a length L that ends in the combustion chamber 4 corresponds to twice the value of the hydraulic channel diameter D.
  • the deflection diffuser 12 and the plenum 13 are therefore omitted.
  • the air from the compressor 1 is divided into a plurality of individual channels immediately after it leaves the compressor 1, specifically into the combustion air channels 15 and into annular channels 20 for the cooling air 21 of the combustion chamber 4 and the turbine 2 arranged on the hub side or on the housing side which is provided here at a high pressure level.
  • air 22 can be removed from the channels 20 for flushing out the boundary layer which forms in the mixing channel 19. This is only shown as an example for the innermost mixing channel 19.
  • the combustion air ducts 15 are designed as diffusers and delay the air speed to approximately half the value of the compressor outlet speed, with a maximum of 75% of the dynamic energy being able to be converted into pressure gain.
  • one or more longitudinal vortices per combustion air channel 15 are generated on the longitudinal vortex generator 16.
  • fuel 6 which is supplied for example through fuel lances 7, is mixed with the air 14 by an integrated fuel injector 17.
  • the fuel injection 17 can also be arranged downstream of the longitudinal vortex generator 16.
  • the longitudinal vortices generated guarantee good mixing of fuel 6 and combustion air 14 in the subsequent mixing channels 19. These have a constant height H and are approximately twice as long as two hydraulic channel diameters D.
  • the mixing channels 19 have a circular cross section, are a mixing tube.
  • the mixing tube axes 24 are arranged parallel to the axis 25 of the gas turbine.
  • the mixing channels 19 can also have a right-angled or polygonal cross section or they can also be a segmented annular gap.
  • the longitudinal vortices caused by the longitudinal vortex generator 16 generate a total swirl in the mixing channel 19, which after the fuel / air mixture 23 has emerged into the combustion chamber 4 leads to a highly turbulent flame stabilization zone in which the vortex bursts and a zone on the axis is present very low or negative axial speed is generated.
  • a flashback in the mixing zone can be caused by a balanced axial speed profile with an increase on the axis and by a additional injection of air 22 into the boundary layer of the mixing channel 19 can be reliably prevented.
  • a burner air duct 15 can be coupled directly to, for example, one or two blade ducts of the last row of compressors.
  • FIGS. 1 and 2 the reduction in the area of the combustion chamber wall to be cooled can be clearly seen according to the invention.
  • a gas turbine from the 170 MWel class, eg GT13E2 should serve as an example. While according to the prior art (FIG. 1) the outer diameter in the area of the combustion chamber is approximately 4.5 m, this value is only 3.5 m when using the invention, so that the size is reduced by approx. 20% is reached. Due to the greatly reduced area to be cooled in the new combustion chamber and the extremely low NOx emissions that can be achieved with good premix burner technology at relatively high flame temperatures (theoretically approx. 5 ppm NOx at 15% O 2 and 1850 K flame temperature), the combustion chamber can be cooled via film or effusion cooling.
  • FIG. 3 and 4 show a further exemplary embodiment.
  • FIG. 3 shows a partial cross section of a two-row annular combustion chamber corresponding to a section in the plane III-III of the four-row combustion chamber shown in FIG. 2.
  • the annular combustion chamber 4 according to FIG. 3 is therefore equipped with two rows of premix burners 5.
  • the arrows in FIG. 3 are intended to illustrate an opposing angle of incidence of the burners 5 in the adjacent rows. This opposite angle of attack ensures that no overall swirl is generated in the combustion chamber 4.
  • the cross section of the mixing channels In this exemplary embodiment, 19 is not round, but rather elliptical.
  • the mixing tube axes 24 are positioned in relation to the shaft in the circumferential direction, i.e. the mixing tube axis 24 forms an angle of ⁇ about 45 ° with the machine axis 25. This improves the mixing and flame stabilization in the combustion chamber 4.
  • combustion air ducts 15 are arranged spirally around the axis 25 of the gas turbine in order to keep the axial length of the machine as small as possible.
  • the invention is particularly suitable for the use of MBtu as fuel, that is to say fuel with a medium calorific value, which arises, for example, from the gasification of heavy oil, coal and tar.
  • the fuel admixture can be moved very simply to a higher speed range (> 100 m / s) in order to reliably avoid backfiring to the fuel injector even with these fuels, which are characterized by a high flame speed.
  • the high-frequency (> 1000 Hz) pressure pulsations (wake of the blades) generated by the last row of compressor runs particularly support the fuel-air mixing process because there is only a short deceleration distance between the end of the compressor 1 and the fuel injector 17, i.e. a short one designed as a diffuser Burner air duct 15 is required

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  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
EP96810777A 1995-12-29 1996-11-12 Chambre de combustion annulaire pour turbine à gaz Expired - Lifetime EP0781967B1 (fr)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
DE19549143A DE19549143A1 (de) 1995-12-29 1995-12-29 Gasturbinenringbrennkammer
DE19549143 1995-12-29

Publications (3)

Publication Number Publication Date
EP0781967A2 true EP0781967A2 (fr) 1997-07-02
EP0781967A3 EP0781967A3 (fr) 1999-04-07
EP0781967B1 EP0781967B1 (fr) 2003-04-02

Family

ID=7781645

Family Applications (1)

Application Number Title Priority Date Filing Date
EP96810777A Expired - Lifetime EP0781967B1 (fr) 1995-12-29 1996-11-12 Chambre de combustion annulaire pour turbine à gaz

Country Status (5)

Country Link
US (1) US5839283A (fr)
EP (1) EP0781967B1 (fr)
JP (1) JPH09196379A (fr)
CN (1) CN1088151C (fr)
DE (2) DE19549143A1 (fr)

Cited By (3)

* Cited by examiner, † Cited by third party
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WO2000011403A1 (fr) * 1998-08-18 2000-03-02 Alliedsignal Inc. Dispositif de tourbillonnement de chambre de combustion axial elliptique
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
CN111396926A (zh) * 2020-04-02 2020-07-10 西北工业大学 一种放气式扩压器与火焰筒一体式的燃烧室

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US5839283A (en) 1998-11-24
JPH09196379A (ja) 1997-07-29
DE59610298D1 (de) 2003-05-08
DE19549143A1 (de) 1997-07-03
CN1088151C (zh) 2002-07-24
CN1158383A (zh) 1997-09-03
EP0781967A3 (fr) 1999-04-07

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