EP0193029B1 - Chambre de combustion pour turbines à gaz - Google Patents

Chambre de combustion pour turbines à gaz Download PDF

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Publication number
EP0193029B1
EP0193029B1 EP86101787A EP86101787A EP0193029B1 EP 0193029 B1 EP0193029 B1 EP 0193029B1 EP 86101787 A EP86101787 A EP 86101787A EP 86101787 A EP86101787 A EP 86101787A EP 0193029 B1 EP0193029 B1 EP 0193029B1
Authority
EP
European Patent Office
Prior art keywords
combustion chamber
chamber
burner
burner elements
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.)
Expired
Application number
EP86101787A
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German (de)
English (en)
Other versions
EP0193029A1 (fr
Inventor
Jaan Dr. Hellat
Jakob Dr. Keller
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
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BBC Brown Boveri AG Switzerland
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Filing date
Publication date
Application filed by BBC Brown Boveri AG Switzerland filed Critical BBC Brown Boveri AG Switzerland
Publication of EP0193029A1 publication Critical patent/EP0193029A1/fr
Application granted granted Critical
Publication of EP0193029B1 publication Critical patent/EP0193029B1/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
    • F23R3/42—Continuous combustion chambers using liquid or gaseous fuel characterised by the arrangement or form of the flame tubes or 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
    • 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
    • F23R3/28—Continuous combustion chambers using liquid or gaseous fuel characterised by the fuel supply
    • F23R3/34—Feeding into different combustion zones

Definitions

  • the present invention relates to a combustion chamber for gas turbines according to the first part of claim 1. It also relates to a method for operating such a combustion chamber.
  • combustion chambers with a number of burner elements distributed over the circumference of an essentially circular-cylindrical combustion chamber are known under the name “annular combustion chambers”.
  • ring combustion chambers Compared to single combustion chambers, ring combustion chambers have the advantage of enabling a more compact overall construction of the gas turbine.
  • the smaller dimensions result in general cost advantages in production.
  • the smaller surface area of an annular combustion chamber also means that the cooling problems can be better mastered.
  • the main disadvantages of this conventional design result from the necessity of distributing the power to individual burner elements, particularly when oil atomization and oil supply are problematic.
  • Another disadvantage is the difficulty of the burners in achieving a temperature distribution that is as uniform as possible within a short barrel length.
  • annular combustion chamber which is provided with a number of centrically symmetrical swirl bodies at its air inflow-side and front-side end. These are always arranged in pairs and it can be seen there that the swirl bodies are able to produce swirl flows with an opposite direction of rotation.
  • This publication also shows the interaction of the burner elements with the swirl bodies, wherein the burner element and swirl body can be integrated in a premixing tube. In the meantime, the swirl bodies are arranged in such a way that the individual swirl jets or swirl flows are only able to influence one another weakly.
  • the technology proposed here means that the desired vortex-free flow with a uniform total pressure cannot occur within the length of the combustion chamber: A uniform temperature distribution at the turbine inlet is therefore not guaranteed.
  • This disadvantage could be counteracted by an appropriate extension of the combustion chamber length. With this measure, however, other disadvantages would have to be accepted. So the constructional disadvantages caused by the extension of the combustion chamber length.
  • the impossibility of complying with the NO x emission tolerated by law is more serious here. The reason for this is that low NoX emission values - apart from the influence of an excessively high temperature - can only be maintained if the time of the gas particles in hot oxygen-free zones is as short as possible, namely not more than a few milliseconds.
  • the temperature in the reaction area must not fall below a certain limit. This requirement places a limit on small sizes.
  • the invention seeks to remedy this.
  • the invention is based on the object of minimizing the CO and NO x emissions in a combustion chamber of the type mentioned.
  • the combustion chamber should be characterized by a compact design with low pressure loss.
  • the object of the invention is to provide a uniform temperature distribution in the gas flow at the turbine inlet.
  • the objectives of the invention are achieved solely in that strongly swirled flows with opposite directions of rotation are brought to collision in a mirror-symmetrical arrangement in a small space, in such a way that the two swirl flows neutralize each other with regard to their swirl and that there is only one after the collision chamber short mixing chamber - with a length that corresponds approximately to the hydraulic diameter or the clear width of the mixing chamber - so that the desired homogeneous temperature distribution in the gas flow can occur before turbine entry.
  • Another advantage of the invention is that the permissible air ratio range of the individual burners can be maintained by the staged mode of operation of the individual pairs of burners. This regulation can further be supported by different mass flow effects on the individual mirror elements arranged in mirror symmetry. If this latter option becomes independent, the entire operating range of the combustion chamber can be covered by a few switching stages.
  • Fig. 1 shows a combustion chamber for gas turbines, which is housed in the GT ring housing 1. If the entire combustion chamber is embedded in a GT ring housing 1, it is connected to the compressor outlet 16 and turbine inlet 17. In this case, the GT ring housing wall bears the difference between the compressor end pressure and ambient pressure.
  • the geometrical shape of the combustion chamber is, as the axial section 18 symbolizes, circular-cylindrical and consists of two reaction chambers 8 arranged symmetrically on the end side, with respect to the central axis of the collision chamber 12, and the collision chamber 12 placed between them.
  • the reaction chambers 8 themselves are at their two ends a number of burner elements A, B arranged axially parallel, depending on the performance of the combustion chamber.
  • the two burner elements A, B which are each mirror-symmetrical to each other with respect to the central axis of the collision chamber 12, have the same structure except for the swirl body.
  • the swirl body 6 in the burner element A is oriented in the opposite direction of rotation to the mirror-symmetrically arranged swirl body 11 in the burner element B, as the indication of the swirl flows 13 and 14 is intended to symbolize.
  • the burner element A or B thus consists of a premixing tube 4, a fuel nozzle 5 - here a dual nozzle - and the swirl bodies 6 or 11 just mentioned.
  • Such a dual nozzle 5 has been described in detail in EP-A-0095788.
  • the dual nozzle 5 consists of a number of concentrically arranged ring cylinders: the compressor air 16 is enriched in the premixing tube 4 with gas from the dual nozzle 5 for the premix 3.
  • the pilot nozzle 7 is also operated with gas. Then follows the inside of the secondary air nozzle 9, which surrounds the central oil line opening into an atomizer nozzle.
  • the collision chamber 12 From the collision chamber 12 there is a radially inwardly directed annular mixing chamber 15, which then merges into the turbine inlet 17. Compared to the mixing chamber 15, the collision chamber 12 has a bulge 10 which prevents one-sided flow separations from taking place in the region of the entry into the collision chamber 12.
  • the figure shows that strongly swirled flows with opposite directions of rotation 13, 14 are brought into collision in a mirror-symmetrical arrangement of the burner elements A, B in a small space.
  • the swirl of the two swirl flows 13, 14 completely cancel each other out after a length which corresponds approximately to the inside width b of the mixing chamber 15.
  • the flows are completely mixed according to this length, which makes a homogeneous temperature distribution at the turbine inlet 17 possible.
  • the swirl body 6 shown in the burner element A is oriented not only in the opposite direction of rotation with respect to the mirror-symmetrically arranged swirl body 11 in the burner element B, but also with respect to the two swirl bodies adjacent on the end face.
  • Fig. 2 shows largely the same combustion chamber as it was already explained in Fig. 1.
  • a suitably strong swirl which can be achieved by an outflow angle of the swirl bodies 6, 11 of approximately 45 ° - paired with a nozzle-like narrowing of the premixing tube 4 after the swirl bodies 6, 11 - results in a stable backflow zone (vortex breakdown) in the Reaction chamber B is generated, which begins only slightly offset from the burner level 21 and which initiates the main reaction of the premixed air / fuel mixture.
  • An initial ignition which stabilizes the entire ignition process appropriately and extends the limits of the backfire and the take off, starts from the pilot nozzle 7, which in the case of a premix burner consumes approx. 10% of the fuel and acts as a diffusion burner.
  • the ratio of the dual nozzle diameter d to the premixing tube nozzle end D should preferably be in the interval 1/2 ⁇ d / D ⁇ 1/3.
  • the ratio of the cross-sectional area of the reaction chamber 8 to the free flow cross-sectional area - between the dual nozzle 5 and the premixing tube nozzle end D - of the burner elements A, B opening there should preferably be at least 3, but not more than 8.
  • the ratio of the mixing chamber cross-sectional area to the sum of the cross-sectional area of the reaction chambers 8 should be at least 1 but not more than 3.
  • the length L of the mixing chamber 15 should one or two times the clear width b.
  • the wall part at the transition from a reaction chamber 8 to the mixing chamber 15 should preferably have a radius of curvature R which is approximately one third of the inside width a of the reaction chamber 8.
  • a wall deflection with the same radius of curvature R is provided on the opposite side of the mixing chamber 15, which leads to a bulge 10 of the collision chamber 12 on the outer circumference.
  • This geometric fixation of the combustion chamber has the purpose of supporting the effects of the collision of the two swirl flows 13, 14.
  • the combustion chamber is preferably operated in a stepped mode of operation.
  • the staging sequence when using premix burners is selected as follows:
  • the mixing mechanism which is triggered by the frontal collision of the swirl flows 13, 14, is so strong that hot and cold flows (eg stage 2) can be mixed without any problems.
  • the aim is for the total amount of air supplied by the compressor 16 through the burner elements A; B to lead. If there is excess air that is not used for targeted film cooling of the combustion chamber walls, it can be introduced into the collision chamber 12 through nozzles 20. In this way, the excess air is mixed in optimally.
  • the mixing chamber 15 is open down here in the axis of the mixing chamber 15.
  • FIG. 3 is a three-dimensional simplified illustration of the combustion chamber according to FIG. 1.
  • the separation-free circular-cylindrical combustion chamber consisting of reaction chambers 8 and collision chambers 12, merges into a likewise separation-free annular mixing chamber 15.
  • the separation-free design shown here by a number of modular combustion chamber units. These units would then be arranged between the compressor and the turbine at regular intervals around the GT axis, the collision principle resulting from the arrangement of the burner elements A, B and the operating mode of the swirl bodies 6, 11 having to be maintained for each module.
  • the annular-cylindrical combustion chamber changes into a cylindrical one, the cross-sectional relationships listed above being maintained.
  • the individual mixing chambers 15 would of course then have to open into an annular collecting chamber before the turbine inlet 17.
  • Fig. 4 shows how the individual burner elements A, B are placed on the end face of the annular reaction chambers and regularly distributed over the circumference. So that the individual opposing pairs of burner elements A, B, which indeed produce swirl currents opposite to one another in the direction of rotation, do not substantially interfere with one another in certain combustion chamber sizes, the swirl bodies 6, 11 in the individual burner elements A, B can have an alternating direction of rotation in the circumferential direction.

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  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Engine Equipment That Uses Special Cycles (AREA)

Claims (12)

1. Chambre de combustion pour turbines à gaz, formée essentiellement par un espace de combustion cylindro-annulaire, qui, du côté d'admission de l'air, comporte des éléments de brûleurs (A, B) régulièrement répartis dans le sens circonférentiel et constitués chacun d'un ajutage de combustible (5), d'un tube de prémélange (4) et d'ouïes de vrillage (6 ou 11), deux éléments de brûleurs voisins étant chacun pourvus d'ouïes de vrillage (6, 11) s'enroulant en sens opposés, et les conduites d'alimentation de combustible individuelles (19) étant raccordées aux éléments de brûleurs par une conduite annulaire de combustible (2), caractérisée en ce que l'espace de combustion cylindro-annulaire de la chambre de combustion est formé de deux chambres de réaction d'about (8) et d'une chambre de collision intermédiaire (12), les chambres de réaction (8) étant pourvues, à leurs extrémités d'about, d'un certain nombre d'éléments de brûleurs (A, B) disposés parallèlement dans le sens axial, qui, par rapport à l'axe central de la chambre de collision (12), de laquelle part une chambre de mélange annulaire (15), sont respectivement disposés de manière énantiomorphe, chaque élément de brûleur (A, B) étant pourvu d'une ouïe de vrillage (6, 11) qui, par rapport à l'ouïe de vrillage énantiomorphe de contrepartie, présente une orientation à sens d'enroulement opposé.
2. Chambre de combustion suivant la revendication 1, caractérisée en ce que la longueur (1) de la chambre de réaction (8) vaut de préférence 1 à 2 fois sa largeur intérieure (a).
3. Chambre de combustion suivant la revendication 1, caractérisée en ce que l'angle de sortie des ouïes de vrillage (6, 11) est de préférence de 45°.
4. Chambre de combustion suivant la revendication 1, caractérisée en ce que les tubes de prémélange (4) présentent un rétrécissement en forme d'ajutage en aval des ouïes de vrillage (6, 11).
5. Chambre de combustion suivant l'une ou l'autre des revendications 1 et 4, caractérisée en ce que le rapport entre le diamètre (d) de l'ajutage de combustible et l'extrémité d'ajutage (D) du tube de prémélange est de préférence situé dans l'intervalle ½<dl/D<½.
6. Chambre de combustion suivant l'une ou l'autre des revendications 2 et 5, caractérisée en ce que le rapport de l'aire de la section de la chambre de réaction (8) aux aires de sections d'écoulement libre, entre le diamètre (d) de l'ajutage de combustible et l'extrémité d'ajutage (D) du tube de prémélange, des éléments de brûleurs (A, B) qui y débouchent est de préférence au minimum de trois et au maximum de huit.
7. Chambre de combustion suivant la revendication 1, caractérisée en ce que le rapport de l'aire de la sectionde la chambre de mélange (15) à la somme des aires des sections des chambres de réaction (8) est de préférence au minimum de 1 et au maximum de 3.
8. Chambre de combustion suivant la revendication 1, caractérisée en ce que la longueur (L) de la chambre de mélange (15) vaut de préférence 1 à 2 fois son diamètre (d).
9. Chambre de combustion suivant la revendication 1, caractérisée en ce que le rayon de courbure (R) à la transition entre la chambre de réaction (8) et la chambre de mélange (15) vaut, de préférence, % de la largeur intérieure (a) de la chambre de réaction (8).
10. Chambre de combustion suivant la revendication 1, caractérisée en ce que la chambre de mélange (15) est centrée part rapport à l'axe de symétrie de la chambre de collision (12).
11. Procédé de conduite de la chambre de combustion suivant la revendication 1, caractérisé en ce que le flux massique passant par les éléments de brûleur (B) vaut le double de celui passant par les éléments de brûleur (A) énantiomorphes.
12. Chambre de combustion suivant la revendication 1, pour la réalisation du procédé suivant la revendication 11, caractérisée en ce que l'axe de symétrie de la chambre de mélange (15) est de préférence incliné de 30° vers le brûleur (A).
EP86101787A 1985-02-26 1986-02-13 Chambre de combustion pour turbines à gaz Expired EP0193029B1 (fr)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
CH861/85 1985-02-26
CH86185 1985-02-26

Publications (2)

Publication Number Publication Date
EP0193029A1 EP0193029A1 (fr) 1986-09-03
EP0193029B1 true EP0193029B1 (fr) 1988-11-17

Family

ID=4196892

Family Applications (1)

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EP86101787A Expired EP0193029B1 (fr) 1985-02-26 1986-02-13 Chambre de combustion pour turbines à gaz

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US (1) US4765146A (fr)
EP (1) EP0193029B1 (fr)
JP (1) JPS61202017A (fr)
DE (1) DE3661224D1 (fr)

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CH672366A5 (fr) * 1986-12-09 1989-11-15 Bbc Brown Boveri & Cie
US4991398A (en) * 1989-01-12 1991-02-12 United Technologies Corporation Combustor fuel nozzle arrangement
GB9023004D0 (en) * 1990-10-23 1990-12-05 Rolls Royce Plc A gas turbine engine combustion chamber and a method of operating a gas turbine engine combustion chamber
FR2672667B1 (fr) * 1991-02-13 1994-12-09 Snecma Chambre de combustion pour turboreacteur a faible niveau d'emissions polluantes.
DE4238323C2 (de) * 1992-11-13 2003-04-24 Alstom Mischer für Gase und/oder Flüssigkeiten
US5497613A (en) * 1993-12-03 1996-03-12 Westinghouse Electric Corporation Hot gas manifold system for a dual topping combustor gas turbine system
US6182451B1 (en) * 1994-09-14 2001-02-06 Alliedsignal Inc. Gas turbine combustor waving ceramic combustor cans and an annular metallic combustor
JPH08261011A (ja) * 1995-03-27 1996-10-08 Mitsubishi Heavy Ind Ltd ガスタービン装置
DE19860583A1 (de) * 1998-12-29 2000-07-06 Abb Alstom Power Ch Ag Brennkammer für eine Gasturbine
US6430919B1 (en) * 2000-03-02 2002-08-13 Direct Propulsion Devices, Inc. Shaped charged engine
DE10121768B4 (de) * 2001-05-04 2007-03-01 Robert Bosch Gmbh Durchmischungsvorrichtung für Gase in Brennstoffzellen
JP2003065537A (ja) * 2001-08-24 2003-03-05 Mitsubishi Heavy Ind Ltd ガスタービン燃焼器
AU2002347186A1 (en) * 2002-01-14 2003-07-24 Alstom Technology Ltd Burner arrangement for the annular combustion chamber of a gas turbine
TWI273642B (en) * 2002-04-19 2007-02-11 Ulvac Inc Film-forming apparatus and film-forming method
EP1847778A1 (fr) * 2006-04-21 2007-10-24 Siemens Aktiengesellschaft Système de combustion à prémélange d'une turbine à gaz et son procédé de fonctionnement
US7766006B1 (en) * 2007-03-09 2010-08-03 Coprecitec, S.L. Dual fuel vent free gas heater
US9546601B2 (en) * 2012-11-20 2017-01-17 General Electric Company Clocked combustor can array
CN115127119B (zh) * 2021-03-26 2023-11-24 中国航发商用航空发动机有限责任公司 抑制环形燃烧室振荡燃烧的方法

Citations (1)

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EP0095788A1 (fr) * 1982-05-28 1983-12-07 BBC Aktiengesellschaft Brown, Boveri & Cie. Chambre de combustion d'une turbine à gaz et sa méthode

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Publication number Priority date Publication date Assignee Title
EP0095788A1 (fr) * 1982-05-28 1983-12-07 BBC Aktiengesellschaft Brown, Boveri & Cie. Chambre de combustion d'une turbine à gaz et sa méthode

Also Published As

Publication number Publication date
US4765146A (en) 1988-08-23
JPS61202017A (ja) 1986-09-06
DE3661224D1 (en) 1988-12-22
EP0193029A1 (fr) 1986-09-03

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