EP1384950A2 - Chambre de combustion annulaire pour une turbine à gaz - Google Patents

Chambre de combustion annulaire pour une turbine à gaz Download PDF

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Publication number
EP1384950A2
EP1384950A2 EP03405504A EP03405504A EP1384950A2 EP 1384950 A2 EP1384950 A2 EP 1384950A2 EP 03405504 A EP03405504 A EP 03405504A EP 03405504 A EP03405504 A EP 03405504A EP 1384950 A2 EP1384950 A2 EP 1384950A2
Authority
EP
European Patent Office
Prior art keywords
combustion chamber
lining segments
segments
chamber according
divided
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
EP03405504A
Other languages
German (de)
English (en)
Other versions
EP1384950A3 (fr
EP1384950B1 (fr
Inventor
Peter Graf
Stefan Tschirren
Helmar Wunderle
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 Technology GmbH
Original Assignee
Alstom Technology AG
Alstom Schweiz 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, Alstom Schweiz AG filed Critical Alstom Technology AG
Publication of EP1384950A2 publication Critical patent/EP1384950A2/fr
Publication of EP1384950A3 publication Critical patent/EP1384950A3/fr
Application granted granted Critical
Publication of EP1384950B1 publication Critical patent/EP1384950B1/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
    • F23R3/42—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

Definitions

  • the present invention relates to the field of gas turbine technology. It relates to an annular combustion chamber for a gas turbine according to the Preamble of claim 1.
  • Such a combustion chamber as e.g. 3 is shown at Gas turbines in use for a long time.
  • the combustion chamber 26, which is part of a gas turbine, not shown and of which only the section above the turbine axis is reproduced extends in the longitudinal direction along the turbine axis in Flow direction (from right to left in Fig. 3).
  • On the entry page (right Page in Fig. 3) is a number on a circular ring concentric to the turbine axis distributed by burners 27, which in the present case are known as so-called Double cone burners are designed according to EP 0321809.
  • burners 27 which in the present case are known as so-called Double cone burners are designed according to EP 0321809.
  • the swirling fuel-air mixture emerging from the burners 27 burns with flame formation in the primary zone following the burner 27 30 and the resulting hot gases emerge from the combustion chamber 26 at one Combustion chamber outlet 31 out and into the subsequent turbine section, where it expand under work performance.
  • Protecting gases are on the inside of the combustion chamber walls 29 special lining segments ("liner segments") 28 arranged and attached.
  • the lining segments 28 are continuous in the axial direction and therefore as long as the interior of the combustion chamber 26. This has the advantage that the number of parts and the length of the leaky column is minimal.
  • the known configuration of the lining segments has the disadvantage that that the segments are comparatively long. This creates in terms of manufacturability and mechanical integrity problems. These problems will be even larger and possibly not solvable if for very large gas turbines correspondingly large combustion chambers with very long lining segments are needed.
  • the object is achieved by the entirety of the features of claim 1.
  • the essence of the invention is that in a combustion chamber the mentioned type, the lining segments in the axial direction in several successive arranged parts are divided. By dividing the individual Partial elements smaller, which simplifies their manufacture and the mechanical stability increased. At the same time, the assembly of the segments is simplified.
  • the lining segments according to a preferred embodiment of the invention in two parts are divided when the liner segments are divided where the flow rate the hot gases is low, or if the liner segments are divided such that the lengths of the individual segment parts in axial direction are approximately the same.
  • the assembly can be further simplified if according to another embodiment the invention the lining segments attached to segment carriers are, and the segment carrier also divided into several parts in the axial direction are.
  • the lining segments are preferably convection-cooled.
  • the divided lining segments can be separately convection-cooled be through the downstream portions of the liner segments flowing cooling medium into the hot gas flow of the combustion chamber becomes.
  • Connection channels are provided through which the convectively cooling cooling medium from one part of the lining segments to the other part of the lining segments can flow.
  • Fig. 1 is a section through a combustion chamber arranged in a gas turbine with lining segments subdivided in the axial direction according to a preferred Embodiment of the invention reproduced.
  • the gas turbine 10 from which is only shown a part lying above the turbine axis has one outer turbine housing 11 on which a filled with compressed air Surrounds plenum 12, in which the actual annular combustion chamber 13 is arranged is.
  • the flow course takes place from right to left in FIG. 1.
  • burners 14, 15 which in two
  • the fuel-air mixture enters the primary zone one above the other 32 blown into the combustion chamber 13 and burns there with the formation of flames.
  • the resulting hot gases pass through the combustion chamber outlet 33 out of the combustion chamber 13 and into the subsequent turbine.
  • the combustion chamber 13 is separated from the surrounding plenum by several segment carriers 18, .., 21 12 separated.
  • segment carriers 18, .., 21 On the inner walls of the segment carrier 18, .., 21 are in axial First and second lining segments 16 and 17 fastened in the direction one behind the other, with inner (in Fig. 1 lower) and outer (in Fig. 1 upper) lining segments are provided.
  • the split liner segments 16, 17 have about the same (axial) length and are separated where the corresponding ones Butt segment carriers 19, 20 and 18, 21 together.
  • the place where the butted divided liner segments 16, 17 (space 24 in Fig. 2), is where the flow rate of the hot gases is low.
  • the split liner segments 16, 17 are convection cooled in the same way, as is already the case with the undivided lining segments is.
  • the division of the segment carriers 18, .., 21 ensures that the assembly is simplified. This applies in particular to the inner (lower) lining. If the inner lining is assembled from two parts, the Separation gap screwed along the entire length. The dividing line the segment carrier 18, 21 for the second lining segments 17 are for Bolt accessible so that a wedge is no longer needed.
  • the division of the lining segments according to the invention makes it possible to to realize larger combustion chambers without having correspondingly large ones Segments must be constructed. This way you can look at already proven ones Use segment sizes.
  • the invention also makes it possible in different Gas turbines have the same burners 14, 15 and first liner segments 16 to use. Adapted to different turbine inlet geometries then only the combustion chamber outlet 33 with the second lining segments 17 and their segment carriers 18, 21.
  • the configuration of the liner segments 16, 17 is the same as for the EV and SEV combustion chambers of the known gas turbines of the applicant of the type GT24B and GT26B (see the article by D. K. Mukherjee "State-of-the-art gas turbines - a brief update ", ABB review 2/1997, pp. 4-14 (1997)).
  • a specialty is the provision of connecting channels 22, 23 (Fig. 1 and Fig. 2) between the second liner segments 17 and the first liner segments 16. Through these connecting channels 22, 23 for the convective cooling of the Lining segments 16, 17 used cooling air from the second lining segments 17 flow into the first lining segments 16 and there to Contribute cooling.
  • the cooling system of the second liner segments 17 is operated only with a part of the total cooling mass flow to the flow velocities to avoid pressure drops in the connection channels 22, 23 to keep small.
  • the transition area between the inner second and first liner segments 17 and 16 is in Fig. 2 shown enlarged.

Landscapes

  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
EP03405504A 2002-07-25 2003-07-07 Chambre de combustion annulaire pour une turbine à gaz Expired - Lifetime EP1384950B1 (fr)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
DE10233805 2002-07-25
DE10233805A DE10233805B4 (de) 2002-07-25 2002-07-25 Ringförmige Brennkammer für eine Gasturbine

Publications (3)

Publication Number Publication Date
EP1384950A2 true EP1384950A2 (fr) 2004-01-28
EP1384950A3 EP1384950A3 (fr) 2007-04-04
EP1384950B1 EP1384950B1 (fr) 2012-10-17

Family

ID=29796562

Family Applications (1)

Application Number Title Priority Date Filing Date
EP03405504A Expired - Lifetime EP1384950B1 (fr) 2002-07-25 2003-07-07 Chambre de combustion annulaire pour une turbine à gaz

Country Status (3)

Country Link
US (1) US7350360B2 (fr)
EP (1) EP1384950B1 (fr)
DE (1) DE10233805B4 (fr)

Families Citing this family (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP2282124A1 (fr) * 2009-08-03 2011-02-09 Alstom Technology Ltd Procédé de modification d'une chambre à combustion d'une turbine à gaz
US9255484B2 (en) * 2011-03-16 2016-02-09 General Electric Company Aft frame and method for cooling aft frame
US9612017B2 (en) 2014-06-05 2017-04-04 Rolls-Royce North American Technologies, Inc. Combustor with tiled liner
US10465907B2 (en) 2015-09-09 2019-11-05 General Electric Company System and method having annular flow path architecture
US10598380B2 (en) 2017-09-21 2020-03-24 General Electric Company Canted combustor for gas turbine engine
US11047575B2 (en) 2019-04-15 2021-06-29 Raytheon Technologies Corporation Combustor heat shield panel

Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4446693A (en) 1980-11-08 1984-05-08 Rolls-Royce Limited Wall structure for a combustion chamber
EP0321809A1 (fr) 1987-12-21 1989-06-28 BBC Brown Boveri AG Procédé pour la combustion de combustible liquide dans un brûleur
US5363643A (en) 1993-02-08 1994-11-15 General Electric Company Segmented combustor
EP1363075A2 (fr) 2002-05-16 2003-11-19 United Technologies Corporation Panneaux de protection thermique pour une chambre de combustion de turbine à gaz

Family Cites Families (10)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CH257279A (de) 1946-01-14 1948-09-30 Hunziker Reinhold Fangvorrichtung für Kleintiere wie Mäuse.
CH428324A (de) * 1964-05-21 1967-01-15 Prvni Brnenska Strojirna Brennkammer
US4567730A (en) * 1983-10-03 1986-02-04 General Electric Company Shielded combustor
US4628694A (en) * 1983-12-19 1986-12-16 General Electric Company Fabricated liner article and method
US5435127A (en) * 1993-11-15 1995-07-25 General Electric Company Method and apparatus for boosting ram airflow to an ejection nozzle
GB2298267B (en) * 1995-02-23 1999-01-13 Rolls Royce Plc An arrangement of heat resistant tiles for a gas turbine engine combustor
US6301877B1 (en) * 1995-11-13 2001-10-16 United Technologies Corporation Ejector extension cooling for exhaust nozzle
DE59706065D1 (de) * 1996-09-26 2002-02-21 Siemens Ag Hitzeschildkomponente mit kühlfluidrückführung und hitzeschildanordnung für eine heissgasführende komponente
DE19727407A1 (de) * 1997-06-27 1999-01-07 Siemens Ag Hitzeschild
US6606861B2 (en) * 2001-02-26 2003-08-19 United Technologies Corporation Low emissions combustor for a gas turbine engine

Patent Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4446693A (en) 1980-11-08 1984-05-08 Rolls-Royce Limited Wall structure for a combustion chamber
EP0321809A1 (fr) 1987-12-21 1989-06-28 BBC Brown Boveri AG Procédé pour la combustion de combustible liquide dans un brûleur
US5363643A (en) 1993-02-08 1994-11-15 General Electric Company Segmented combustor
EP1363075A2 (fr) 2002-05-16 2003-11-19 United Technologies Corporation Panneaux de protection thermique pour une chambre de combustion de turbine à gaz

Non-Patent Citations (1)

* Cited by examiner, † Cited by third party
Title
D. K. MUKHERJEE: "State-of-the-art gas turbines - a brief update", ABB REVIEW, vol. 2, 1997, pages 4 - 14

Also Published As

Publication number Publication date
EP1384950A3 (fr) 2007-04-04
DE10233805B4 (de) 2013-08-22
US20040154308A1 (en) 2004-08-12
DE10233805A1 (de) 2004-02-05
US7350360B2 (en) 2008-04-01
EP1384950B1 (fr) 2012-10-17

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