EP0489193A1 - Chambre de combustion pour turbine à gaz - Google Patents

Chambre de combustion pour turbine à gaz Download PDF

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Publication number
EP0489193A1
EP0489193A1 EP90123311A EP90123311A EP0489193A1 EP 0489193 A1 EP0489193 A1 EP 0489193A1 EP 90123311 A EP90123311 A EP 90123311A EP 90123311 A EP90123311 A EP 90123311A EP 0489193 A1 EP0489193 A1 EP 0489193A1
Authority
EP
European Patent Office
Prior art keywords
combustion chamber
gas turbine
wall parts
flame tube
cooling air
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
EP90123311A
Other languages
German (de)
English (en)
Other versions
EP0489193B1 (fr
Inventor
Pierre Meylan
Hans Schwarz
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.)
ABB Asea Brown Boveri Ltd
ABB AB
Original Assignee
ABB Asea Brown Boveri Ltd
Asea Brown Boveri AB
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 Asea Brown Boveri Ltd, Asea Brown Boveri AB filed Critical ABB Asea Brown Boveri Ltd
Priority to DE59010740T priority Critical patent/DE59010740D1/de
Priority to EP90123311A priority patent/EP0489193B1/fr
Priority to US07/799,316 priority patent/US5226278A/en
Priority to JP31888291A priority patent/JP3180830B2/ja
Publication of EP0489193A1 publication Critical patent/EP0489193A1/fr
Application granted granted Critical
Publication of EP0489193B1 publication Critical patent/EP0489193B1/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
    • 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
    • F23R3/002—Wall structures
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05B—INDEXING SCHEME RELATING TO WIND, SPRING, WEIGHT, INERTIA OR LIKE MOTORS, TO MACHINES OR ENGINES FOR LIQUIDS COVERED BY SUBCLASSES F03B, F03D AND F03G
    • F05B2260/00—Function
    • F05B2260/20—Heat transfer, e.g. cooling
    • F05B2260/202—Heat transfer, e.g. cooling by film cooling

Definitions

  • the invention relates to a gas turbine combustion chamber with an annular flame tube, which delimits a combustion chamber and is exposed on its side facing away from the combustion chamber to an air flow supplied by the compressor of the gas turbine, and which is composed essentially of overlapping wall parts, the wall parts on their side facing away from the combustion chamber each have a plurality of inlet openings distributed over the circumference, through which the cooling air is introduced into a distribution space arranged in the flame tube and communicating with the combustion space.
  • the invention has for its object to minimize the cooling air consumption in a gas turbine combustion chamber of the type mentioned in order to reduce the emission of NO x .
  • the wall parts are curved elements in the turbine axial direction, which overlap in the circumferential direction and are provided with means for guiding the cooling air from the distribution space arranged at the inlet-side end of the wall part, at least approximately in the circumferential direction to the outlet-side end of the wall part.
  • the cooling air flowing out of the overlap gaps between two adjacent wall parts is deflected in a grille before entering the combustion chamber.
  • the angle of attack of the grille can be increasingly changed from the flame tube inlet to its outlet in order to match the swirling flow of the combustion gases near the wall.
  • the turbine 1, of which the first axially flowed stages in the form of three guide rows 2 'and run rows 2''is shown in FIG. 1, essentially consists of the bladed turbine rotor 3 and the blade carrier 4 equipped with guide blades.
  • the blade carrier is in the Turbine housing 5 suspended.
  • the turbine housing 5 also includes the collecting space 6 for the compressed combustion air. From this collecting space, the combustion air reaches the annular combustion chamber 7, which in turn opens into the turbine inlet, ie upstream of the first guide row 2 '.
  • the compressed air arrives in the collecting space from the diffuser 8 of the compressor 9. Of the latter, only the last three stages are shown in the form of three guide rows 10 'and three rows 10''.
  • the rotor blades of the compressor and the turbine sit on a common shaft 11, the central axis of which represents the longitudinal axis 12 of the gas turbine unit.
  • the compressed combustion air enters the collecting chamber 6 in the direction of the arrow in the burner 13, which is only shown by way of example, of which 36 pieces are evenly distributed around the circumference.
  • the fuel is injected into the combustion chamber 15 via a fuel nozzle 14.
  • the fuel nozzle In the plane of the primary air inlet, the fuel nozzle is surrounded by a swirl body 16 in the form of vortex blades.
  • the air reaches the primary zone of the combustion chamber 15 through the vortex blades, in which the combustion process takes place.
  • the vortex blades create a swirl flow with an air core directed against the burner, which anchors the flame to the burner so that it does not tear off despite the high air speed.
  • the turbulent flow ensures rapid combustion.
  • the annular combustion chamber 15 extends downstream of the burner orifices up to the turbine inlet. It is delimited both inside and outside by the flame tube 17.
  • this flame tube is designed as a self-supporting structure. It consists of one on both its inner ring and its outer ring Number of longitudinally arranged wall parts 18 with tangential overlap gaps 22 (FIGS. 2 and 6). These wall parts, which can be cast parts, are bent in the axial direction of the turbine in accordance with the course of the combustion chamber through which flow and extend over the entire axial length of the flame tube.
  • the number of wall parts is determined from the requirement that the cooling air flowing into the combustion chamber from the gaps should be used as efficiently as possible as film cooling. It follows that the distance between two cooling air gaps and thus the tangential extent of a wall part is approximately as large as the effective length of the cooling air film. And from this one can see the manufacturing advantage, among other things, that only as many columns or wall parts have to be provided as are actually necessary. Furthermore, this design allows the implementation of ring-shaped flame tubes of any dimensions and geometries. This design is easy to maintain, if only in the event of damage, only the damaged wall parts have to be replaced.
  • the flame tube on its side facing away from the combustion chamber is exposed to the air flow in the collecting chamber 6 supplied by the compressor 9.
  • the wall parts On their side facing the collecting space 6, the wall parts each have a plurality of inlet openings (19 in FIGS. 5 and 6) distributed over the circumference, via which the cooling air flows into a distribution space (20 in FIG. 5) arranged in the wall part and communicating with the combustion chamber and 6) is initiated.
  • the cooling air duct on the wall parts 18 is shown schematically in FIG.
  • the cooling air is guided as far as possible in the circumferential direction along the surfaces of the wall parts facing the collecting space 6 using the means described below.
  • the cooling air When flowing into the combustion chamber 15, the cooling air must of course not be directed against the swirl flow of the combustion gases indicated by arrows. This means that the inflow openings and the outflow gaps in the wall parts of the inner tube of the flame tube are arranged exactly opposite to those of the outer tube of the flame tube.
  • the cooling air Seen against the direction of flow of the combustion gases, which in this view have a swirl in the counterclockwise direction, the cooling air also flows through the outer ring in a counterclockwise direction, while brushing the wall parts of the inner ring in a clockwise direction.
  • the requirement also applies that the cooling air is introduced into the combustion chamber 15 in order to maintain cooling film in such a way that it not only coincides in the same direction, but also in its direction as closely as possible in the direction of flow of the combustion gases near the wall of the flame tube.
  • FIG. 3 in which the flow conditions in the combustion chamber are shown on the basis of the partial development of a cylindrical section.
  • this 3 denotes the vertical B the plane of the burner mouth, the vertical T the plane of the turbine inlet.
  • the flow in the combustion chamber will be explained on the basis of numerical data, which, however, can only have an exemplary character due to numerous other decisive parameters.
  • the combustion air leaves the swirl body at an angle of approx. 75 °.
  • X there is an acceleration of the working medium due to the combustion, which leads to a slight deflection in the axial direction.
  • the fuel gases now flow at an angle of approx. 55 °.
  • zone Y the gas flow is accelerated in the axial direction; the flow through the channel becomes increasingly steeper (Fig. 1). This constriction before the turbine inlet causes the gases in zone Z to be deflected to approximately 20 ° and thus act on the guide vanes 2 ′ of the first turbine stage.
  • FIG. 4 and 5 show a top view of the structure of a wall part 18, specifically the side facing the collecting space.
  • 6 shows a wall part of the inner flame tube ring in cross section.
  • the wall parts are not bent in the circumferential direction. Rather, it is almost flat plates that are bent according to the longitudinal direction of the turbine according to the course of the combustion chamber. These plates are on one side on the side facing the collecting space provided at the first end with a holding device in the form of a gripper 21. With this gripper 21, the adjacent plate in the circumferential direction is held, as shown by the broken line at the left end of the plate. This creates a simple assembly means which also allows the overlap gap 22 to be kept within narrow limits in every operating state.
  • an offset 23 is provided which can be used for fastening purposes of the flame tube.
  • the flame tube structure is self-supporting; it is understood that this is only possible up to a certain order of magnitude.
  • these offsets 23 can be connected to the actual supporting structures on the wall parts. These are to be designed in any case so that a free expansion of the wall parts is not hindered during operation.
  • the wall parts are equipped with longitudinal ribs 24, which extend from the inlet-side distribution chamber 20 to the outlet-side passages 30. These passages can be designed as bores in a bead carrying the grippers 21.
  • the longitudinal ribs 24 divide the side of the wall part facing away from the combustion chamber 15 into channels 25, in which the cooling air is guided in the circumferential direction to the passages 30.
  • Both the distribution space 20 as well as the ribs 24 and the channels 25 are provided with a cover 26 against the collecting space 6. In this cover there are a number of inlet openings 19 for the cooling air in the plane of the distribution space 20. 5, although they are invisible in this view, since the cover 26 has been omitted in FIGS. 4 and 5 for reasons of clarity.
  • the distribution space 20 at the inlet-side end of the wall part is divided into a plurality of distribution segments 28 by means of partition walls 27.
  • the cooling air flowing out of the passages 30 into the overlap gap 22 is deflected in a grille 29 before entering the combustion chamber 15.
  • This is arranged on the inlet-side end of the overlapping adjacent wall part (FIG. 6) on its side facing the combustion chamber.
  • the angle of attack of the grille is increasingly changed from the flame tube inlet to its outlet in accordance with the swirling flow of the combustion gases near the wall.
  • the invention is not limited to the embodiment shown and described.
  • the long sides of the wall parts could just as well run helically, for example at 45 °.
  • this grating could just as well be designed as a separate structural unit.
  • the ribs will only be attached to a part of the walls instead of over their entire axial length. It is also conceivable for the surface of the wall part to be grooved instead of the longitudinal ribs, with or without turbulence grille.

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)
EP90123311A 1990-12-05 1990-12-05 Chambre de combustion pour turbine à gaz Expired - Lifetime EP0489193B1 (fr)

Priority Applications (4)

Application Number Priority Date Filing Date Title
DE59010740T DE59010740D1 (de) 1990-12-05 1990-12-05 Gasturbinen-Brennkammer
EP90123311A EP0489193B1 (fr) 1990-12-05 1990-12-05 Chambre de combustion pour turbine à gaz
US07/799,316 US5226278A (en) 1990-12-05 1991-11-27 Gas turbine combustion chamber with improved air flow
JP31888291A JP3180830B2 (ja) 1990-12-05 1991-12-03 ガスタービンコンバスタ

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
EP90123311A EP0489193B1 (fr) 1990-12-05 1990-12-05 Chambre de combustion pour turbine à gaz

Publications (2)

Publication Number Publication Date
EP0489193A1 true EP0489193A1 (fr) 1992-06-10
EP0489193B1 EP0489193B1 (fr) 1997-07-23

Family

ID=8204799

Family Applications (1)

Application Number Title Priority Date Filing Date
EP90123311A Expired - Lifetime EP0489193B1 (fr) 1990-12-05 1990-12-05 Chambre de combustion pour turbine à gaz

Country Status (4)

Country Link
US (1) US5226278A (fr)
EP (1) EP0489193B1 (fr)
JP (1) JP3180830B2 (fr)
DE (1) DE59010740D1 (fr)

Cited By (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP0648979A1 (fr) * 1993-10-18 1995-04-19 ABB Management AG Méthode et moyens pour le refroidissement d'une chambre de combustion d'une turbine à gaz
US5735115A (en) * 1994-01-24 1998-04-07 Siemens Aktiengesellschaft Gas turbine combustor with means for removing swirl
US6003297A (en) * 1995-03-06 1999-12-21 Siemens Aktiengsellschaft Method and apparatus for operating a gas turbine, with fuel injected into its compressor
WO2024021252A1 (fr) * 2022-07-28 2024-02-01 哈电发电设备国家工程研究中心有限公司 Cylindre interne de carter de guidage d'air de turbine à gaz à double chambre de combustion

Families Citing this family (17)

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Publication number Priority date Publication date Assignee Title
GB9220937D0 (en) * 1992-10-06 1992-11-18 Rolls Royce Plc Gas turbine engine combustor
US5460002A (en) * 1993-05-21 1995-10-24 General Electric Company Catalytically-and aerodynamically-assisted liner for gas turbine combustors
US5906093A (en) * 1997-02-21 1999-05-25 Siemens Westinghouse Power Corporation Gas turbine combustor transition
US20050044857A1 (en) * 2003-08-26 2005-03-03 Boris Glezer Combustor of a gas turbine engine
US7010921B2 (en) * 2004-06-01 2006-03-14 General Electric Company Method and apparatus for cooling combustor liner and transition piece of a gas turbine
US7574865B2 (en) * 2004-11-18 2009-08-18 Siemens Energy, Inc. Combustor flow sleeve with optimized cooling and airflow distribution
US7934382B2 (en) * 2005-12-22 2011-05-03 United Technologies Corporation Combustor turbine interface
AU2009216831B2 (en) * 2008-02-20 2014-11-20 General Electric Technology Gmbh Gas turbine
US9080464B2 (en) * 2008-02-27 2015-07-14 Mitsubishi Hitachi Power Systems, Ltd. Gas turbine and method of opening chamber of gas turbine
CH699309A1 (de) * 2008-08-14 2010-02-15 Alstom Technology Ltd Thermische maschine mit luftgekühlter, ringförmiger brennkammer.
US20120027578A1 (en) * 2010-07-30 2012-02-02 General Electric Company Systems and apparatus relating to diffusers in combustion turbine engines
US9267687B2 (en) 2011-11-04 2016-02-23 General Electric Company Combustion system having a venturi for reducing wakes in an airflow
US9739201B2 (en) 2013-05-08 2017-08-22 General Electric Company Wake reducing structure for a turbine system and method of reducing wake
US9322553B2 (en) 2013-05-08 2016-04-26 General Electric Company Wake manipulating structure for a turbine system
US9435221B2 (en) 2013-08-09 2016-09-06 General Electric Company Turbomachine airfoil positioning
DE102014206018A1 (de) * 2014-03-31 2015-10-01 Siemens Aktiengesellschaft Gasturbinenanlage
US11073285B2 (en) 2019-06-21 2021-07-27 Raytheon Technologies Corporation Combustor panel configuration with skewed side walls

Citations (5)

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Publication number Priority date Publication date Assignee Title
GB642257A (en) * 1947-12-04 1950-08-30 Shell Refining & Marketing Co Improvements in and relating to combustion chambers
GB1099374A (en) * 1965-03-23 1968-01-17 Prvni Brnenska Strojirna Zd Y Improvements in or relating to cooled walls of gas-turbine combustion chambers
US3422620A (en) * 1967-05-04 1969-01-21 Westinghouse Electric Corp Combustion apparatus
GB2102558A (en) * 1981-06-12 1983-02-02 Westinghouse Electric Corp Combustor or combustion turbine
US4773227A (en) * 1982-04-07 1988-09-27 United Technologies Corporation Combustion chamber with improved liner construction

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US2647369A (en) * 1946-09-06 1953-08-04 Leduc Rene Combustion chamber for fluid fuel burning in an air stream of high velocity
DE1059719B (de) * 1955-06-16 1959-06-18 Jan Jerie Dr Ing Gekuehlte Wand einer Verbrennungskammer, insbesondere fuer Gasturbinen
US3420058A (en) * 1967-01-03 1969-01-07 Gen Electric Combustor liners
US3978662A (en) * 1975-04-28 1976-09-07 General Electric Company Cooling ring construction for combustion chambers
US4077205A (en) * 1975-12-05 1978-03-07 United Technologies Corporation Louver construction for liner of gas turbine engine combustor
US4996838A (en) * 1988-10-27 1991-03-05 Sol-3 Resources, Inc. Annular vortex slinger combustor

Patent Citations (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB642257A (en) * 1947-12-04 1950-08-30 Shell Refining & Marketing Co Improvements in and relating to combustion chambers
GB1099374A (en) * 1965-03-23 1968-01-17 Prvni Brnenska Strojirna Zd Y Improvements in or relating to cooled walls of gas-turbine combustion chambers
US3422620A (en) * 1967-05-04 1969-01-21 Westinghouse Electric Corp Combustion apparatus
GB2102558A (en) * 1981-06-12 1983-02-02 Westinghouse Electric Corp Combustor or combustion turbine
US4773227A (en) * 1982-04-07 1988-09-27 United Technologies Corporation Combustion chamber with improved liner construction

Cited By (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP0648979A1 (fr) * 1993-10-18 1995-04-19 ABB Management AG Méthode et moyens pour le refroidissement d'une chambre de combustion d'une turbine à gaz
US5651253A (en) * 1993-10-18 1997-07-29 Abb Management Ag Apparatus for cooling a gas turbine combustion chamber
US5735115A (en) * 1994-01-24 1998-04-07 Siemens Aktiengesellschaft Gas turbine combustor with means for removing swirl
US6003297A (en) * 1995-03-06 1999-12-21 Siemens Aktiengsellschaft Method and apparatus for operating a gas turbine, with fuel injected into its compressor
WO2024021252A1 (fr) * 2022-07-28 2024-02-01 哈电发电设备国家工程研究中心有限公司 Cylindre interne de carter de guidage d'air de turbine à gaz à double chambre de combustion

Also Published As

Publication number Publication date
JPH04273913A (ja) 1992-09-30
JP3180830B2 (ja) 2001-06-25
DE59010740D1 (de) 1997-09-04
EP0489193B1 (fr) 1997-07-23
US5226278A (en) 1993-07-13

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