EP1413831A1 - Chambre de combustion annulaire pour turbine à gaz et turbine à gaz - Google Patents

Chambre de combustion annulaire pour turbine à gaz et turbine à gaz Download PDF

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Publication number
EP1413831A1
EP1413831A1 EP02023471A EP02023471A EP1413831A1 EP 1413831 A1 EP1413831 A1 EP 1413831A1 EP 02023471 A EP02023471 A EP 02023471A EP 02023471 A EP02023471 A EP 02023471A EP 1413831 A1 EP1413831 A1 EP 1413831A1
Authority
EP
European Patent Office
Prior art keywords
combustion chamber
lining element
hooking means
annular combustion
axial direction
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.)
Withdrawn
Application number
EP02023471A
Other languages
German (de)
English (en)
Inventor
Peter Tiemann
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.)
Siemens AG
Original Assignee
Siemens 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 Siemens AG filed Critical Siemens AG
Priority to EP02023471A priority Critical patent/EP1413831A1/fr
Priority to US10/672,510 priority patent/US6938424B2/en
Priority to JP2003357231A priority patent/JP4347657B2/ja
Priority to CNB2003101024611A priority patent/CN100532947C/zh
Publication of EP1413831A1 publication Critical patent/EP1413831A1/fr
Withdrawn legal-status Critical Current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
    • F23R3/60—Support structures; Attaching or mounting means
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23M—CASINGS, LININGS, WALLS OR DOORS SPECIALLY ADAPTED FOR COMBUSTION CHAMBERS, e.g. FIREBRIDGES; DEVICES FOR DEFLECTING AIR, FLAMES OR COMBUSTION PRODUCTS IN COMBUSTION CHAMBERS; SAFETY ARRANGEMENTS SPECIALLY ADAPTED FOR COMBUSTION APPARATUS; DETAILS OF COMBUSTION CHAMBERS, NOT OTHERWISE PROVIDED FOR
    • F23M5/00—Casings; Linings; Walls
    • F23M5/02—Casings; Linings; Walls characterised by the shape of the bricks or blocks used
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23M—CASINGS, LININGS, WALLS OR DOORS SPECIALLY ADAPTED FOR COMBUSTION CHAMBERS, e.g. FIREBRIDGES; DEVICES FOR DEFLECTING AIR, FLAMES OR COMBUSTION PRODUCTS IN COMBUSTION CHAMBERS; SAFETY ARRANGEMENTS SPECIALLY ADAPTED FOR COMBUSTION APPARATUS; DETAILS OF COMBUSTION CHAMBERS, NOT OTHERWISE PROVIDED FOR
    • F23M5/00—Casings; Linings; Walls
    • F23M5/04—Supports for linings
    • 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
    • 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 invention relates to an annular combustion chamber for a gas turbine, the annular combustion chamber extending along an axial direction, including a combustion chamber and having on its inner side facing the combustion chamber a supporting structure on which a lining element attached to this lines the annular combustion chamber.
  • gas turbines are widely used to convert fossil energies into electrical energy in connection with a generator.
  • the fuel is mixed with compressed air and fed to a combustion chamber in which it is burned.
  • the resulting working medium flows along a hot gas duct past several turbine stages.
  • Each turbine stage consists of a large number of guide and rotor blades arranged separately in two rings.
  • the guide vanes are attached to a fixed stator and the rotor blades to a rotor that drives the generator.
  • the combustion chamber is located in a combustion chamber, which is lined with heat-resistant lining elements.
  • Lining elements of a combustion chamber in the sense of the invention are liners and further components which delimit the combustion chamber and which are arranged in a combustion chamber and are exposed to the hot gas.
  • the combustion chamber is lined with a plurality of lining elements which are adjacent to one another in the axial direction and in the circumferential direction of the turbine shaft.
  • a liner is known from US Pat. No. 4,614,082.
  • a combustion chamber is shown in FIG. 2, which has a plurality of liners.
  • the adjacent liners overlap in such a way that the flow in the direction of the Working medium seen front liner with its end covering the beginning of the following liner. This also applies to the liners following in the direction of flow of the working medium, which thus form a sequence of overlapping liners.
  • the required stiffness of a liner in relation to the conditions prevailing in the combustion chamber is produced by a side wall running in the circumferential direction, which extends over the entire width of the liner. This side wall of the liner is arranged on the rear side facing away from the hot gas. It extends away from it and bends further in the axial direction in order to reach behind adjacent liners.
  • closed, cooled ring combustion chambers are lined with liners which are provided on their rear side facing away from the hot gas with side walls extending in the axial direction.
  • the liners themselves are very stiff due to their side walls, which is necessary due to the conditions prevailing in the combustion chamber.
  • the rails carrying the liners, which are arranged within the annular combustion chamber, can therefore be made softer.
  • the liner also has its own rigidity due to the side wall. This rigidity, in conjunction with the temperature fluctuations caused by the start-up of the gas turbine, the operation and the shutdown, cause tension between the support structure and the liner, which make it more difficult to remove the lining element from the annular combustion chamber. It is also important to note that the lining elements must withstand the static and dynamic pressures prevailing in the combustion chamber.
  • the underlying task for the invention is to specify an annular combustion chamber, the lining elements of which meet the mechanical requirements, such as rigidity and secure attachment, while at the same time being easy to maintain.
  • Another object of the invention is to provide a maintenance-friendly gas turbine.
  • annular combustion chamber with a lining element is specified according to the invention, wherein a plurality of hooking means are arranged on the lining element on two edge regions running in the axial direction on their rear side facing away from the combustion chamber, which have a hook width in the axial direction, and that the lining element is attached to the corresponding supporting structure in such a way that to remove the lining element from the supporting structure, the lining element is displaced in the axial direction by the hook width of the hooking means.
  • the selected arrangement, shape and position of the hooking means of the lining element makes it easy to mount an individual lining element.
  • the lining element itself has an axial softness due to the large number of interlocking elements spaced apart from one another. In the unassembled state, this is only determined by the wall thickness of the lining element.
  • the axial softness of the lining element contributes to the simple and safe assembly and disassembly in addition to the relatively short displacement path, which corresponds to the width of a hooking means.
  • the lining element mounted on the rigid and solid support structure takes on its rigidity. The rigidity of the required for the operation of the gas turbine Lining element is then given in the assembled state.
  • the axial softness of the lining element itself advantageously contributes to the fact that the tension between the supporting structure and the lining element that is usually present in the assembled state does not occur at all due to the thermal stress. Thus, only small forces are required to disassemble an inventive lining element.
  • a lining element can be assembled and disassembled independently of lining elements adjacent in the axial and circumferential directions to the turbine shaft.
  • a large number of further hooking means is arranged centrally as a central receptacle between two edge regions of the lining element that run in the axial direction.
  • a coolant flows between the combustion chamber and the back of the lining element facing away from the hot gas, e.g. Cooling air or cooling steam, which has a higher pressure than the working medium.
  • the higher pressure of the coolant on the rear side of the lining element applied to the working medium can possibly cause a deformation of the lining element towards the working medium. This deformation is reduced to an extent to be tolerated, in which the span to be bridged in the circumferential direction between the two edge regions is reduced by further hooking means arranged in the center.
  • the further hooking means arranged in the center can have identical or similar profiles to the hooking means of the edge regions, or they can also be significantly different.
  • each hooking means in the axial direction Lining element has a distance that is identical to or greater than the hook width of the hooking means, allows the removal of the assembled lining element after its displacement by this hook width.
  • each hooking means has an identical hook width.
  • two hooking means of the lining element which are directly adjacent in the axial direction are at a distance which is twice as large as the hook width of a hooking means.
  • two hooking means of the lining element which are directly adjacent in the axial direction are at a distance which is three times as large as the hook width of a hooking means.
  • Each distance between two hooking means of the lining element which are immediately adjacent in the axial direction is preferably identical.
  • a symmetrical and uniform design of frequently used elements such as hooking means simplify the manufacture of the lining element.
  • the lining element has stiffening ribs running in the circumferential direction of the annular combustion chamber on its rear side facing away from the combustion chamber. These increase the stiffness of the lining element that is already present in the circumferential direction. Accidental bending of the lining element in the radial direction can consequently be reduced and possibly avoided.
  • the stiffening ribs are preferably spaced apart from the hooking means. This results in between the ends of the stiffening ribs and the interlocking elements local bending points arranged.
  • the stiffening ribs ensure a rigidity of the lining element in the central region between the hooking means lying opposite one another in the circumferential direction, the local bending points in turn facilitating the installation and removal of the lining element.
  • the tensions between the support structure and the lining element that result from thermal loads do not have a negative effect on the dismantling of the lining element, ie greater expenditure of force is not required for the dismantling.
  • interlocking elements are preferably L-shaped and / or T-shaped.
  • Other forms of interlocking elements are also suitable for the lining elements.
  • spherical, or conical or frustoconical and similar interlocking elements such as a bayonet perform the same task.
  • the object related to the gas turbine is achieved by a gas turbine with an annular combustion chamber according to one of the above-mentioned embodiments.
  • FIG. 1 shows a gas turbine 1 with a housing 2, a compressor 3, an annular combustion chamber 4 and a plurality of turbine stages 5 connected downstream of the annular combustion chamber 4.
  • the air sucked in by the compressor 3 is compressed therein and then passed on to a burner 6.
  • the compressed air is mixed with a fuel and burned into a working medium M when it is injected into a combustion chamber 7 arranged in the annular combustion chamber 4.
  • the working medium M then flows through a hot gas duct 21 past the turbine stages 5, which are each formed by a plurality of guide vanes 22 and rotor blades 23 arranged separately in two rings.
  • the energy of the working medium M is converted into rotational energy by means of the rotor blades 23 arranged on a rotor 8 rotatably mounted about the axis of rotation 9.
  • annular combustion chamber 4 in cross section.
  • the annular combustion chamber 4 is open to the outflow end 24 facing the hot gas duct 21.
  • the burner 6 is arranged on the outflow end 24 facing the hot gas duct 21, opposite the injection end 25 of the annular combustion chamber 4. Between the injection end 25 and the outflow end 24 of the annular combustion chamber 4, the latter is lined with a plurality of lining elements 10 which are adjacent to one another and which are fastened to a support structure 26.
  • FIG. 2a shows a perspective view of an annular combustion chamber 4 which is partially open for the sake of better description.
  • the annular combustion chamber 4 is lined with a plurality of lining elements 10 which are arranged in the circumferential direction U in ring form 27.
  • FIG. 3 shows a lining element 10 which has a multiplicity of interlocking means 11 on the rear side 13 facing away from the hot gas.
  • These hooking means 11 are arranged in the two edge regions 15 of the lining element 10 extending in the axial direction A.
  • Each hooking means 11 has a width B.
  • the hooking means 11 are essentially L-shaped. They emboss from the rear side 13 of the lining element 10 and, as the course progresses, bend at right angles to the nearest side edge 16 of the lining element 11 which runs in the axial direction. The distances between two immediately adjacent hooking means 11 are designated by L.
  • the lining element 10 To fasten the lining element 10 to the corresponding support structure 26 of an annular combustion chamber 4, it is introduced into a recess in the support structure 26 receiving the hooking means 11 and displaced at least by the width B until the hooking means 11 are completely hooked to the support structure 26.
  • the hooking means 11 of the lining element 10 and the support structure 26 then firmly engage in one another.
  • FIG. 4 shows a lining element 10 which has stiffening ribs 12 on the rear side facing away from the hot gas.
  • the stiffening ribs 12 run in the circumferential direction U and are spaced apart from the hooking means 11.
  • the stiffening ribs 12 reduce the deflection of the lining wall 17 during operation of the gas turbine 1.
  • the ends 18 of the stiffening ribs 12 are spaced apart from the interlocking elements 11, so that local bending points 19 cause a low local softness there, which simplifies the removal and installation of the lining element 10 ,
  • a lining element 10 which has a so-called center receptacle 14 on the rear side 13 facing away from the hot gas is shown in Fig. 5.
  • the center receptacle 14 consists of further, individual interlocking elements 20 which, viewed in the circumferential direction U, are arranged centrally between two interlocking elements 11 arranged in different edge regions 15.
  • This center receptacle 14 reduces the deflection of the lining wall 17 during operation by reducing the span between the edge regions 15 and thus contributes to the rigidity.
  • the other hooking means 20 are essentially T-shaped. They form away from the back 13 and then bend tangentially to the circumferential direction U in two arms.
  • FIG. 6 shows a section through an annular combustion chamber 4, to which a lining element 10 is attached.
  • the support structure 26 is arranged on the side of the annular combustion chamber 4 facing the combustion chamber 7.
  • This has hooking means 28 which are designed to correspond to those of the lining elements 10.
  • the hooking means 11 of the lining element 10 engage in the corresponding hooking means 28 of the support structure 26.
  • the width B of an interlocking element 11 is smaller than the distance L between two adjacent interlocking elements 11.
  • the interlocking means 28 of the support structure 26 are spaced apart from one another at least as far as the interlocking elements 11 of the lining element 10 are wide.
  • stiffening ribs 12 extending in the circumferential direction U are arranged on the rear side 13 of the lining element 10 facing away from the combustion chamber 7, stiffening ribs 12 extending in the circumferential direction U are arranged.
  • the lining element 10 is detached from the support structure 26 by displacing the lining element 10 in or opposite to the axial direction A by at least the width B of a hooking means 11.
  • the fastening mechanism which is composed of the hooking elements 11 of the lining element 10 and the supporting structure 26 corresponding thereto, can have relatively large component tolerances. Oversizing the lining element 10 in relation to the corresponding support structure 26 does not pose a problem, since the axial softness in conjunction with the local bending points 19 arranged in the circumferential direction U compensate for the excess of the lining element 10.

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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)
EP02023471A 2002-10-21 2002-10-21 Chambre de combustion annulaire pour turbine à gaz et turbine à gaz Withdrawn EP1413831A1 (fr)

Priority Applications (4)

Application Number Priority Date Filing Date Title
EP02023471A EP1413831A1 (fr) 2002-10-21 2002-10-21 Chambre de combustion annulaire pour turbine à gaz et turbine à gaz
US10/672,510 US6938424B2 (en) 2002-10-21 2003-09-26 Annular combustion chambers for a gas turbine and gas turbine
JP2003357231A JP4347657B2 (ja) 2002-10-21 2003-10-17 ガスタービンと燃焼器
CNB2003101024611A CN100532947C (zh) 2002-10-21 2003-10-21 燃气透平的环形燃烧室及燃气透平

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
EP02023471A EP1413831A1 (fr) 2002-10-21 2002-10-21 Chambre de combustion annulaire pour turbine à gaz et turbine à gaz

Publications (1)

Publication Number Publication Date
EP1413831A1 true EP1413831A1 (fr) 2004-04-28

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EP02023471A Withdrawn EP1413831A1 (fr) 2002-10-21 2002-10-21 Chambre de combustion annulaire pour turbine à gaz et turbine à gaz

Country Status (4)

Country Link
US (1) US6938424B2 (fr)
EP (1) EP1413831A1 (fr)
JP (1) JP4347657B2 (fr)
CN (1) CN100532947C (fr)

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EP2463583A1 (fr) * 2010-12-06 2012-06-13 Alstom Technology Ltd Turbine à gaz et procédé de reconditionnement d'une telle turbine à gaz
EP2886962A1 (fr) * 2013-12-23 2015-06-24 Rolls-Royce plc Chambre de combustion
EP3009744A1 (fr) * 2014-10-13 2016-04-20 Rolls-Royce plc Élément de garniture pour une chambre de combustion et procédé associé
EP3098391A3 (fr) * 2015-05-07 2017-02-22 General Electric Company Brides anti-cordage de bande de turbine
DE102015225107A1 (de) 2015-12-14 2017-06-14 Rolls-Royce Deutschland Ltd & Co Kg Gasturbinenbrennkammer mit Schindelbefestigung mittels Rastelementen
DE102016217876A1 (de) 2016-09-19 2018-03-22 Rolls-Royce Deutschland Ltd & Co Kg Brennkammerwand einer Gasturbine mit Befestigung einer Brennkammerschindel
DE102018204453A1 (de) 2018-03-22 2019-09-26 Rolls-Royce Deutschland Ltd & Co Kg Brennkammerbaugruppe mit unterschiedlichen Krümmungen für eine Brennkammerwand und eine hieran fixierte Brennkammerschindel

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US20100095680A1 (en) * 2008-10-22 2010-04-22 Honeywell International Inc. Dual wall structure for use in a combustor of a gas turbine engine
US20100095679A1 (en) * 2008-10-22 2010-04-22 Honeywell International Inc. Dual wall structure for use in a combustor of a gas turbine engine
US8448416B2 (en) * 2009-03-30 2013-05-28 General Electric Company Combustor liner
US8695322B2 (en) * 2009-03-30 2014-04-15 General Electric Company Thermally decoupled can-annular transition piece
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US9897317B2 (en) * 2012-10-01 2018-02-20 Ansaldo Energia Ip Uk Limited Thermally free liner retention mechanism
EP2722495B1 (fr) * 2012-10-17 2015-03-11 ABB Turbo Systems AG Carter d'entrée de gaz et turbine associée pour gaz d'échappement
US10969103B2 (en) * 2013-08-15 2021-04-06 Raytheon Technologies Corporation Protective panel and frame therefor
US10808937B2 (en) * 2013-11-04 2020-10-20 Raytheon Technologies Corporation Gas turbine engine wall assembly with offset rail
GB201405496D0 (en) 2014-03-27 2014-05-14 Rolls Royce Plc Linear assembly
GB201406386D0 (en) 2014-04-09 2014-05-21 Rolls Royce Plc Gas turbine engine
US10041675B2 (en) 2014-06-04 2018-08-07 Pratt & Whitney Canada Corp. Multiple ventilated rails for sealing of combustor heat shields
EP2952812B1 (fr) * 2014-06-05 2018-08-08 General Electric Technology GmbH Chambre de combustion annulaire d'une turbine á gaz et segment de manchon
US9989255B2 (en) * 2014-07-25 2018-06-05 General Electric Company Liner assembly and method of turbulator fabrication
US9534785B2 (en) * 2014-08-26 2017-01-03 Pratt & Whitney Canada Corp. Heat shield labyrinth seal
CN107003000B (zh) * 2014-09-29 2019-09-24 西门子股份公司 用于燃烧室的隔热罩的隔热元件
DE102014226707A1 (de) * 2014-12-19 2016-06-23 Rolls-Royce Deutschland Ltd & Co Kg Gasturbinenbrennkammer mit veränderter Wandstärke
WO2018087878A1 (fr) 2016-11-11 2018-05-17 川崎重工業株式会社 Chemise de chambre de combustion
KR101867549B1 (ko) * 2017-04-21 2018-06-15 주식회사 전일특수금속 소각로용 에어노즐 라이너
KR102099307B1 (ko) * 2017-10-11 2020-04-09 두산중공업 주식회사 라이너 냉각을 촉진하는 난류 생성 구조 및 이를 포함하는 가스 터빈용 연소기
US11143108B2 (en) * 2019-03-07 2021-10-12 Pratt & Whitney Canada Corp. Annular heat shield assembly for combustor
KR102313106B1 (ko) * 2020-09-15 2021-10-15 정윤도 소각로용 에어 노즐 라이너
KR102310029B1 (ko) * 2020-12-15 2021-10-07 주식회사 마이크로원 일체형 사이드월 라이너
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CH704185A1 (de) * 2010-12-06 2012-06-15 Alstom Technology Ltd Gasturbine sowie verfahren zum rekonditionieren einer solchen gasturbine.
AU2011253595B2 (en) * 2010-12-06 2015-07-16 General Electric Technology Gmbh Gas turbine and method for reconditioning such a gas turbine
EP2463583A1 (fr) * 2010-12-06 2012-06-13 Alstom Technology Ltd Turbine à gaz et procédé de reconditionnement d'une telle turbine à gaz
US9903590B2 (en) 2013-12-23 2018-02-27 Rolls-Royce Plc Combustion chamber
EP2886962A1 (fr) * 2013-12-23 2015-06-24 Rolls-Royce plc Chambre de combustion
EP3009744A1 (fr) * 2014-10-13 2016-04-20 Rolls-Royce plc Élément de garniture pour une chambre de combustion et procédé associé
US10451277B2 (en) 2014-10-13 2019-10-22 Rolls-Royce Plc Liner element for a combustor, and a related method
EP3098391A3 (fr) * 2015-05-07 2017-02-22 General Electric Company Brides anti-cordage de bande de turbine
US10392950B2 (en) 2015-05-07 2019-08-27 General Electric Company Turbine band anti-chording flanges
DE102015225107A1 (de) 2015-12-14 2017-06-14 Rolls-Royce Deutschland Ltd & Co Kg Gasturbinenbrennkammer mit Schindelbefestigung mittels Rastelementen
EP3182006A1 (fr) 2015-12-14 2017-06-21 Rolls-Royce Deutschland Ltd & Co KG Chambre de combustion de turbine à gaz comprenant une fixation de bardeau au moyen d'éléments d'encliquetage
US10060626B2 (en) 2015-12-14 2018-08-28 Rolls-Royce Deutschland Ltd & Co Kg Gas turbine combustion chamber with a shingle attachment by means of catching elements
DE102016217876A1 (de) 2016-09-19 2018-03-22 Rolls-Royce Deutschland Ltd & Co Kg Brennkammerwand einer Gasturbine mit Befestigung einer Brennkammerschindel
DE102018204453A1 (de) 2018-03-22 2019-09-26 Rolls-Royce Deutschland Ltd & Co Kg Brennkammerbaugruppe mit unterschiedlichen Krümmungen für eine Brennkammerwand und eine hieran fixierte Brennkammerschindel
US11320144B2 (en) 2018-03-22 2022-05-03 Rolls-Royce Deutschland Ltd & Co Kg Combustion chamber assembly with different curvatures for a combustion chamber wall and a combustion chamber shingle fixed thereto
DE102018204453B4 (de) 2018-03-22 2024-01-18 Rolls-Royce Deutschland Ltd & Co Kg Brennkammerbaugruppe mit unterschiedlichen Krümmungen für eine Brennkammerwand und eine hieran fixierte Brennkammerschindel

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US20040074239A1 (en) 2004-04-22
US6938424B2 (en) 2005-09-06
CN100532947C (zh) 2009-08-26
JP4347657B2 (ja) 2009-10-21
CN1497218A (zh) 2004-05-19

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