EP1016830B1 - Chambre de combustion pour une turbine à gaz - Google Patents

Chambre de combustion pour une turbine à gaz Download PDF

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Publication number
EP1016830B1
EP1016830B1 EP99811182A EP99811182A EP1016830B1 EP 1016830 B1 EP1016830 B1 EP 1016830B1 EP 99811182 A EP99811182 A EP 99811182A EP 99811182 A EP99811182 A EP 99811182A EP 1016830 B1 EP1016830 B1 EP 1016830B1
Authority
EP
European Patent Office
Prior art keywords
combustion chamber
chamber
torus
annular
diffuser
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
EP99811182A
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German (de)
English (en)
Other versions
EP1016830A3 (fr
EP1016830A2 (fr
Inventor
Jakob J. 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.)
GE Vernova GmbH
Original Assignee
Alstom Technology AG
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Filing date
Publication date
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Publication of EP1016830A2 publication Critical patent/EP1016830A2/fr
Publication of EP1016830A3 publication Critical patent/EP1016830A3/fr
Application granted granted Critical
Publication of EP1016830B1 publication Critical patent/EP1016830B1/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
    • 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

Definitions

  • the present invention relates to the field of gas turbines. It affects a combustion chamber for a gas turbine, in which combustion chamber with fuel mixed with an air stream entering the combustion chamber and then is burned, and the resulting combustion air flow downstream of the combustion chamber is fed to a turbine.
  • a combustion chamber for a gas turbine in which combustion chamber with fuel mixed with an air stream entering the combustion chamber and then is burned, and the resulting combustion air flow downstream of the combustion chamber is fed to a turbine.
  • gas turbines are often designed in such a way that the one passed through them Air flow through two combustion chambers and, accordingly, twice over turbines is directed.
  • the air drawn in is initially passed through a compressor group led, and then passed into a primary combustion chamber, where supplied Fuel-air mixture is ignited and burned.
  • Fuel-air mixture is ignited and burned.
  • From the primary combustion chamber the hot combustion air flows over a first turbine and becomes downstream fed to the first turbine of a secondary combustion chamber, where fuel and if necessary additional air is added and the mixture ignites. Since the from the gases escaping from the first turbine are often very hot, i.e. above the auto-ignition temperature of the fuels can be in the secondary combustion chamber mostly do without active ignition. Downstream of the secondary combustion chamber there is then a second turbine, over which the hot combustion gases flow out of the secondary combustion chamber.
  • Such gas turbines the individual components usually in series along a major axis of the Gas turbine arranged.
  • a gas turbine group is from, for example EP 0 620 362 A1 known.
  • the individual channels for the air flows and the combustion chambers are mostly all in the form of hollow cylinders, which reach around the axis of the gas turbine.
  • Secondary combustion chambers for such gas turbines are generally proportionate simply designed because they do not need burners, but the fuel after a suitable swirling of the hot ones emerging from the first turbine Air can simply be injected into the air flow via nozzles, and itself the mixture ignites itself after a characteristic time.
  • the hollow-cylindrical secondary combustion chamber is, for example, out known from EP 0 669 500 A1.
  • the task is thereby in a combustion chamber of the type mentioned solved that the combustion chamber has an annular diffuser, in which the air flow enters that downstream of and in connection with the diffuser standing at least one substantially annular toroidal chamber arranged is that downstream of the annular toroidal chamber and over its circumference distributed mixing tubes branch, and that downstream of the mixing tubes an annular Combustion chamber is arranged, in which the mixing tubes open.
  • the core the invention consists in the combination of diffuser, toroidal ring Chamber and mixing tubes have a premixing structure available too place in which air flowing through in an optimal manner, i.e. fast and can be mixed efficiently with fuel. Another advantage comes from that the proposed configuration is reduced to thermoacoustic Oscillations tend.
  • a first preferred embodiment of the combustion chamber according to the invention is characterized in that the combustion chamber as a secondary combustion chamber is designed, and that the gas turbine has a primary combustion chamber, a downstream the primary combustion chamber acting first turbine, a downstream of the first turbine acting secondary combustion chamber, and one downstream of the secondary combustion chamber has acting second turbine.
  • the use of the combustion chamber as a secondary combustion chamber is advantageous, especially with such use at high Mach numbers short mixing times are required.
  • the ignition as described in a further embodiment, in the secondary combustion chamber is done by self-ignition, the quick and backflow-free mixing is in the proposed arrangement advantageous, and so it can controlled combustion in the area of the outlet of the mixing tubes in the combustion chamber or in Combustion chamber can be guaranteed.
  • the diffuser is designed such that the parallel air flow flowing to the gas turbine axis and entering the combustion chamber is initially deflected in a substantially radial direction, and that the Diffuser acts in a tangential manner on the toroidal chamber, so that the air flow entering the toroidal chamber curls up in the torus and swirls around the toroidal minor axis of the torus.
  • the mixing tubes on the one essentially opposite the diffuser Side attached in a manner essentially parallel to the axis of the gas turbine In this way, two swirls of different screw rotation meet in front of the mixing tubes on top of each other, and then flow with mutual mixing and Annihilation of the vertebrae through the mixing tubes.
  • the proposed design of the combustion chamber is based, among other things, on the spectacular merging and mixing behavior of colliding, with opposing Rotating, rotating, subcritical vortices. This phenomenon occurred on Flow behavior in front of and in radial outlet pipes of steam turbines discovered. It is shown there that only when radial outlet pipes are attached the loss of free flow of the rotating air is possible, while losses occur due to swirling in simple openings. A detailed look at the behavior of the air flow before and in one Such a radial outlet tube shows that two in front of the tubes with opposite directions Rotating, rotating subcritical vortices collide, and the rotation of the two vertebrae within a distance of less than one Completely cancel each other out.
  • FIG. 1 shows a longitudinal section along the axis 21 of a gas turbine.
  • the entire three-dimensional combustion chamber structure is essentially obtained by rotating the section about axis 21, i.e. it deals the channeling components with the exception of the mixing tubes 22 Axially symmetrical parts about the axis 21 of the gas turbine.
  • the hot one from the first Combustion chamber, the primary combustion chamber, escaping airflow flows at the shown gas turbine first via a first turbine 11, which in one Bearing 31 is stored.
  • There is a short one downstream of the first turbine 11 hollow cylindrical discharge line 12 through which the air flow from the first Turbine 10 flows parallel to the axis 21 of the gas turbine.
  • the outflow pipe 12 is preferably just long enough that the axial hollow cylindrical Flow profile in the airflow 10 can recover.
  • Downstream of the discharge line 12 is a diffuser 13 in which the air flow is controlled by the axial direction is deflected. The deflection occurs after in Figure 1 outside in an almost radial direction, but it is also conceivable in principle that Distraction inside.
  • the curvature of the diffuser 13 can with the help of the inverse Euler equations can be optimized. It is essentially in the diffuser 13 therefore, the average flow velocity of the air flow approximately cut in half.
  • the diffuser 13 abuts a toroidal ring in a tangential manner downstream Chamber 14 to.
  • the torus 14 is arranged perpendicular to the axis 21 of the generator, with a large torus radius 29 around axis 21, i.e. the major toric axis 27 and the axis 21 of the generator coincide.
  • the circular line of the large torus radius 29 forms the annular torus minor axis 26, and the Torus outer wall 30 is formed by a small torus radius 28 around the ring Secondary torus axis 26 formed.
  • a plurality are located downstream of the toroidal annular chamber 14 of mixing tubes 22 which are perpendicular to the toroidal secondary axis 26 of branch off the chamber 14 and are distributed over the circumference of the chamber 14. Through these mixing tubes 22, the air flow rolled up in the torus 14 flows out of the annular toroidal chamber 14 from.
  • the mixing tubes 22 are cylindrical or at least partially conical and preferably have a radius in the area of the small torus radius 28.
  • the actual combustion chamber 18, which again essentially as a hollow cylinder is formed around the axis 21, and downstream of this combustion chamber 18 is a second turbine arranged.
  • Figure 2 shows part of a conical section through the chamber 14 and Mixing tubes 22 along the plane X-X in Figure 1 in view from the outside in.
  • the behavior of the air flow in the chamber can be determined on the basis of this section 14 and the mixing tubes 22 illustrate.
  • Air flow which is tangent to the center between two mixing tubes 22 Diffuser 13 enters the chamber 14, this separates into two to the left and vertebrae 24 and 25 evading to the right, which are different Screw direction, 24 corresponds to a left-handed screw, 25 a clockwise.
  • Each of the partial vertebrae 24 and 25 now "screws" in Direction of the nearest mixing tube 22 to flow out of the chamber 14 there to be able to.
  • Nozzles 32 are arranged with which liquid or gaseous Fuel can be injected.
  • liquid Fuel can also be supplied through nozzles 20, which are located opposite the mixing tubes 22 Wall sides of the toroidal chamber 14 attached are injected into the air flow:
  • nozzles 20 which are located opposite the mixing tubes 22
  • Wall sides of the toroidal chamber 14 attached are injected into the air flow:
  • Auto-ignition characteristic of the injected fuel according to the temperature of the Airflow and flow velocity are formed by auto-ignition of the mixture due to the high air temperature a flame front, which either at the area of the outlet of the mixing tubes 22 or behind it in the combustion chamber 18 can come to rest.
  • a I u I A e u e .
  • a I and A E are the cross-sectional areas of the tangential entry of the diffuser 13 into the annular toroidal chamber 14 and the cross-sectional area 23 of the mixing tubes 22, and u I and u E are the corresponding flow velocities.
  • Table 1 gives the values for a secondary combustion chamber with 12 mixing tubes 22, each with an output radius of 300 mm.
  • Size unit Value 1 Value 2 Large radius of the torus m 0.0675 0.0675 Small radius of the torus m 0.0275 0.0275 Width of the entry slot m 0.0085 0.0135 Radius of the tangential input m 0.04425 0.04675 Area of the entrance slot per swirl arm, A I m ⁇ 2 0.01182 0.001983 Eccentricity radius , r I m 0.02325 0.02075 Cross-sectional area of the vertebral arm, A E m ⁇ 2 0.002376 0.002376 vertebrae 1.70 0.90

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  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Fluidized-Bed Combustion And Resonant Combustion (AREA)

Claims (12)

  1. Chambre de combustion (13, 14, 18, 22) pour une turbine à gaz, dans laquelle chambre de combustion (13, 14, 18, 22), du combustible est mélangé à un courant d'air (10) entrant dans la chambre de combustion (13, 14, 18, 22) et est ensuite brûlé, et le courant d'air de combustion se produisant (33) est acheminé en aval de la chambre de combustion (13, 14, 18, 22) à une turbine,
    caractérisée en ce que
    la chambre de combustion (13, 14, 18, 22) présente un diffuseur annulaire (13) dans lequel entre le courant d'air (10), en ce qu'en aval du diffuseur (13), et connectée à celui-ci, est disposée au moins une chambre (14) essentiellement annulaire toroïdale, en ce qu'en aval de la chambre annulaire toroïdale (14) partent des tubes de mélange (22) répartis sur sa circonférence et en ce qu'en aval des tubes de mélange (22) est disposé un espace de combustion annulaire (18) dans lequel débouchent les tubes de mélange (22).
  2. Chambre de combustion (13, 14, 18, 22) selon la revendication 1, caractérisée en ce que la chambre de combustion est réalisée sous forme de chambre de combustion secondaire (13, 14, 18, 22), et en ce que la turbine à gaz présente une chambre de combustion primaire, une première turbine (11) agissant en aval de la chambre de combustion primaire, une chambre de combustion secondaire (13, 14, 18, 22) agissant en aval de la première turbine (11), ainsi qu'une deuxième turbine agissant en aval de la chambre de combustion secondaire (13, 14, 18, 22).
  3. Chambre de combustion (13, 14, 18, 22) selon la revendication 2, caractérisée en ce que la chambre annulaire toroïdale (14) présente un axe imaginaire annulaire du tore (26) qui s'étend avec un gros rayon de tore (29) autour d'un axe principal du tore (27), et dans laquelle chambre annulaire toroïdale (14) une paroi extérieure du tore (30) est formée avec un petit rayon de tore (28) autour de l'axe imaginaire du tore (26) et en ce que l'axe principal du tore (27) est orienté essentiellement parallèlement à l'axe (21) de la turbine à gaz.
  4. Chambre de combustion (13, 14, 18, 22) selon la revendication 3, caractérisée en ce que les tubes de mélange (22) sont essentiellement coniques ou cylindriques, et en ce que les axes des tubes de mélange (22) sont disposés en dehors du plan de la chambre annulaire toroïdale (14) et essentiellement perpendiculairement à l'axe imaginaire annulaire du tore (26).
  5. Chambre de combustion (13, 14, 18, 22) selon la revendication 3, caractérisée en ce que le rayon des tubes de mélange (22) est dans la région du petit rayon du tore (28) ou notamment de préférence est égal à celui-ci.
  6. Chambre de combustion (13, 14, 18, 22) selon l'une quelconque des revendications 4 ou 5, caractérisée en ce que la chambre de combustion primaire est une chambre de combustion annulaire, en ce qu'entre la première turbine (11) et le diffuseur (13) est disposée une conduite de sortie de l'écoulement (12) réalisée sous forme d'un cylindre creux et en ce que l'axe de la conduite de sortie de l'écoulement (12) s'étend parallèlement à l'axe (21) de la turbine à gaz.
  7. Chambre de combustion (13, 14, 18, 22) selon la revendication 6, caractérisée en ce que le diffuseur annulaire (13) est réalisé de telle sorte que le courant d'air (10) s'écoulant à travers la conduite de sortie d'écoulement (12) parallèlement à l'axe (21) de la turbine à gaz est dévié par le diffuseur (13), et en ce que notamment de préférence cette déviation est effectuée dans la direction s'étendant essentiellement radialement par rapport à l'axe (21) de la turbine à gaz.
  8. Chambre de combustion (13, 14, 18, 22) selon l'une quelconque des revendications 6 ou 7, caractérisée en ce que le diffuseur (13) vient en prise de manière tangentielle avec la chambre annulaire toroïdale (14).
  9. Chambre de combustion (13, 14, 18, 22) selon la revendication 8, caractérisée en ce que le diffuseur (13) est réalisé de telle sorte que le courant d'air (10) soit dévié essentiellement radialement vers l'extérieur depuis le diffuseur (13) et en ce que le diffuseur (13) vient en prise de manière tangentielle de l'intérieur avec la chambre annulaire toroïdale (14).
  10. Chambre de combustion (13, 14, 18, 22) selon l'une quelconque des revendications 8 ou 9, caractérisée en ce que les tubes de mélange (16) sont disposés essentiellement en regard de l'ouverture d'entrée du diffuseur (13) dans la chambre annulaire toroïdale (14) et en ce qu'ils sont orientés essentiellement parallèlement à l'axe (21) de la turbine à gaz.
  11. Chambre de combustion (13, 14, 18, 22) selon la revendication 10, caractérisée en ce que des moyens (20, 32) sont disposés en aval de la chambre annulaire toroïdale (14) soit du côté arrière de la chambre annulaire toroïdale (14) opposé aux tubes de mélange (16) soit dans la région centrale des tubes de mélange (16), avec lesquels le combustible liquide peut être injecté.
  12. Chambre de combustion (13, 14, 18, 22) selon l'une quelconque des revendications 10 ou 11, caractérisée en ce que des moyens (32) sont disposés dans la région centrale des tubes de mélange (16) en aval de la chambre annulaire toroïdale (14), par le biais desquels du combustible gazeux peut être injecté.
EP99811182A 1998-12-29 1999-12-21 Chambre de combustion pour une turbine à gaz Expired - Lifetime EP1016830B1 (fr)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
DE19860583 1998-12-29
DE19860583A DE19860583A1 (de) 1998-12-29 1998-12-29 Brennkammer für eine Gasturbine

Publications (3)

Publication Number Publication Date
EP1016830A2 EP1016830A2 (fr) 2000-07-05
EP1016830A3 EP1016830A3 (fr) 2002-07-24
EP1016830B1 true EP1016830B1 (fr) 2004-05-26

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EP99811182A Expired - Lifetime EP1016830B1 (fr) 1998-12-29 1999-12-21 Chambre de combustion pour une turbine à gaz

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US (1) US6272864B1 (fr)
EP (1) EP1016830B1 (fr)
DE (2) DE19860583A1 (fr)

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* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US11175046B2 (en) 2019-05-09 2021-11-16 General Electric Company Combustor premixer assembly including inlet lips
CN116892736B (zh) * 2023-07-10 2025-10-17 中国人民解放军空军工程大学 一种旋转爆震燃烧室分布式喷油装置及方法

Family Cites Families (16)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE1011670B (de) * 1955-06-03 1957-07-04 H C Ernst Schmidt Dr Ing Dr Re Ringfoermige Misch- oder Brennkammer, insbesondere fuer Gasturbinen
US3309866A (en) * 1965-03-11 1967-03-21 Gen Electric Combustion process and apparatus
DE2232025A1 (de) * 1972-06-30 1974-01-17 Motoren Turbinen Union Gasturbinenanlage, insbesondere triebwerk mit gleichraumverbrennung
US4081957A (en) * 1976-05-03 1978-04-04 United Technologies Corporation Premixed combustor
US4073137A (en) * 1976-06-02 1978-02-14 United Technologies Corporation Convectively cooled flameholder for premixed burner
GB2098719B (en) * 1981-05-20 1984-11-21 Rolls Royce Gas turbine engine combustion apparatus
EP0193029B1 (fr) * 1985-02-26 1988-11-17 BBC Brown Boveri AG Chambre de combustion pour turbines à gaz
JPS62158927A (ja) 1986-01-07 1987-07-14 Nissan Motor Co Ltd 環型燃焼器
US4838029A (en) * 1986-09-10 1989-06-13 The United States Of America As Represented By The Secretary Of The Air Force Externally vaporizing system for turbine combustor
CH679799A5 (fr) * 1988-07-25 1992-04-15 Christian Reiter
DE3942042A1 (de) * 1989-12-20 1991-06-27 Bmw Rolls Royce Gmbh Brennkammer fuer eine gasturbine mit luftunterstuetzten brennstoffzerstaeuberduesen
DE4236071C2 (de) * 1992-10-26 2002-12-12 Alstom Verfahren für eine Mehrstufenverbrennung in Gasturbinen
US5575153A (en) * 1993-04-07 1996-11-19 Hitachi, Ltd. Stabilizer for gas turbine combustors and gas turbine combustor equipped with the stabilizer
CH687269A5 (de) 1993-04-08 1996-10-31 Abb Management Ag Gasturbogruppe.
CA2141066A1 (fr) * 1994-02-18 1995-08-19 Urs Benz Procede de refroidissement d'une chambre de combustion a auto-allumage
EP0870990B1 (fr) * 1997-03-20 2003-05-07 ALSTOM (Switzerland) Ltd Turbine à gaz avec chambre de combustion toroidale

Also Published As

Publication number Publication date
DE19860583A1 (de) 2000-07-06
US6272864B1 (en) 2001-08-14
EP1016830A3 (fr) 2002-07-24
DE59909580D1 (de) 2004-07-01
EP1016830A2 (fr) 2000-07-05

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