EP1265035A1 - Doppelbefestigung einer Turbinenbrennkammer aus keramischem Matrix-Verbundwerkstoff - Google Patents

Doppelbefestigung einer Turbinenbrennkammer aus keramischem Matrix-Verbundwerkstoff Download PDF

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Publication number
EP1265035A1
EP1265035A1 EP02291364A EP02291364A EP1265035A1 EP 1265035 A1 EP1265035 A1 EP 1265035A1 EP 02291364 A EP02291364 A EP 02291364A EP 02291364 A EP02291364 A EP 02291364A EP 1265035 A1 EP1265035 A1 EP 1265035A1
Authority
EP
European Patent Office
Prior art keywords
metal
combustion chamber
turbomachine according
fixing means
distributor
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
EP02291364A
Other languages
English (en)
French (fr)
Other versions
EP1265035B1 (de
Inventor
Didier Hernandez
Gwenaelle Calvez
Alexandre Forestier
Eric Conete
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.)
Safran Aircraft Engines SAS
Original Assignee
SNECMA Moteurs SA
SNECMA SAS
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 SNECMA Moteurs SA, SNECMA SAS filed Critical SNECMA Moteurs SA
Publication of EP1265035A1 publication Critical patent/EP1265035A1/de
Application granted granted Critical
Publication of EP1265035B1 publication Critical patent/EP1265035B1/de
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
    • F23R3/60—Support structures; Attaching or mounting means
    • 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
    • F23R3/007—Continuous combustion chambers using liquid or gaseous fuel constructed mainly of ceramic components
    • 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
    • F05B2230/00—Manufacture
    • F05B2230/60—Assembly methods
    • F05B2230/604—Assembly methods using positioning or alignment devices for aligning or centering, e.g. pins
    • F05B2230/606—Assembly methods using positioning or alignment devices for aligning or centering, e.g. pins using maintaining alignment while permitting differential dilatation

Definitions

  • the present invention relates to the specific field of turbomachinery and it is more particularly interested in the problem posed by the assembly of a CMC type composite material combustion chamber (ceramic matrix composite) in the metal casing of a turbomachine.
  • the turbine high pressure in particular its inlet distributor (HPT nozzle), the combustion as well as the casing (or casing) of this chamber are produced in the same material, generally of metallic type.
  • HPT nozzle inlet distributor
  • the combustion as well as the casing (or casing) of this chamber are produced in the same material, generally of metallic type.
  • the use of a metal chamber proves to be totally unsuitable thermal point of view and it must be resorted to a room based on CMC type high temperature composite materials.
  • difficulties of implementation and the cost of these materials make their use is most often limited to the combustion chamber itself, the distributor inlet of the high pressure turbine and the remaining casing then produced more conventionally made of metallic materials.
  • metallic materials and composite materials have very different coefficients of thermal expansion. This results in particularly acute problems of connection with between the casing and the combustion and sealing chamber at the distributor, at the inlet of the high pressure turbine.
  • the present invention overcomes these drawbacks by proposing an assembly of the combustion chamber in the crankcase having the capacity to absorb the displacements induced by the differences in the expansion coefficients of these rooms.
  • a turbomachine comprising, in an envelope made of metallic material and in a direction F of gas flow, a system fuel injection, a combustion chamber made of composite material having a longitudinal axis, and a distributor of metallic material forming the input stage with fixed blades of a high pressure turbine, characterized in that said combustion chamber of composite material is held in position in said metallic envelope by a plurality of flexible metallic tabs having first and second ends, said first ends being interconnected by a metal crown forming a flange fixed to said metal casing by first fixing means and said seconds ends each fixed together by second fixing means on the one hand to said combustion chamber of composite material and on the other hand to one end of a wall of composite material, the other end of which forms a support surface for a sealing element integral with said distributor and ensuring the sealing of the gas stream between said combustion chamber and said distributor, the flexibility of said fixing lugs allowing temperatures a free radial expansion of said combustion chamber made of material composite with respect to said metallic envelope.
  • the first and second fixing means preferably consist by a plurality of bolts.
  • the second fixing means can also be constituted by crimping elements.
  • said sealing element is of the “lamellar” circular joint type. It can include a plurality of calibrated leak holes.
  • the envelope metal is formed in two parts, said metal ring connecting between them said first ends of said flexible metal tabs is mounted between connecting flanges of these two parts.
  • said metal crown can be fixed directly to said annular casing by conventional fixing means.
  • said first ends of the fixing lugs can either be fixed by soldering (or welding) to said metal crown forming flange either form a single piece with this metal crown.
  • the distributor is fixed on a downstream part 14b of the annular casing internal of the turbomachine by first removable fixing means preferably consisting of a plurality of bolts 50 while resting on support means 49 secured to the outer annular envelope of the turbomachine.
  • Passage orifices 54, 56 formed in the metal platforms external 46 and internal 48 of the distributor 42 are further provided to ensure cooling of the stationary vanes 44 of this distributor at the inlet of the rotor of the high pressure turbine from the compressed oxidizer available at the outlet of the diffusion duct 18 and flowing in two streams F1, F2 on either side of the combustion chamber 24.
  • the combustion chamber 24 which has a coefficient of thermal expansion very different from other metal parts forming the turbomachine, is fixedly held in position in its envelope by a plurality of flexible tabs 58, 60 regularly distributed around the chamber combustion between the inner and outer annular shells.
  • These legs fasteners are fitted for a first part of them (see tab referenced 58) between the outer annular envelope 12a, 12b and the axial wall external 26 of the combustion chamber and for a second part (like the tab 60) between the internal annular envelope 14a, 14b and the internal axial wall 28 of the combustion chamber.
  • the number of legs can, for example, be number equal to that of the injection nozzles or equal to a multiple of this number.
  • Each flexible fixing lug made of metallic material which may have a substantially triangular shape as illustrated in FIG. 1A, or to be consisting of a simple blade (not shown of constant width or not), is welded or brazed by a first end 62; 64 to a metal crown 66a, 66b forming a flange and fixed integrally by first fixing means 52; 68 to one or the other (depending on its location) of the annular envelopes metallic external or internal.
  • This flange fixing is intended to facilitate the retention of these legs on the metal envelopes.
  • these legs and the metal crown together form a unique metal piece in one piece.
  • this tab is fixed jointly by second fixing means 74, 76 on the one hand at a downstream end 88; 90 of external axial 26 and internal 28 walls of the material combustion chamber ceramic composite and on the other hand at one end of a composite wall ceramic 78a; 78b arranged in the extension of each of the axial walls external and internal, sort of forming a second part of the chamber, and the other end of which forms a support plane for a sealing element integral with the distributor and ensuring the tightness of the gas stream between the combustion chamber 24 and the distributor 42.
  • connection second ends of the legs 70, 72 with the downstream ends of the walls of the combustion chamber and the first ends of the composite walls ceramic forming the second part of the chamber is carried out by a simple bolting, preferably of captive nut type to facilitate possible assembly / disassembly and correspondingly limit the sizing of the legs.
  • the metal crown 66a, 66b connecting the first ends 62, 64 legs is preferably taken between existing connecting flanges between the upstream parts 12a, 14a and downstream 12b, 14b of the internal annular envelopes and external and fixedly held by the first fixing means 52, 68 which preferably are also of the bolt type.
  • the tightness of the gas stream between the combustion chamber 24 and the distributor 42 is provided by a “lamellar” circular joint 80, 82 mounted in a groove 84, 86 of each of the external 46 and internal 48 platforms of the distributor and which comes to rest directly on the second end part of the ceramic composite wall 78a; 78b forming a support plane for this joint sealing circular.
  • the seal is held in abutment against this second end of the composite wall by means of an elastic element, of the type leaf spring 92, 94, fixed on the distributor.
  • FIG. 1B illustrates a first variant of the embodiment previous in which the fixing of the legs (only the case of the leg 60 is illustrated) to the downstream end 90 of the combustion chamber 24 is effected by a connection crimped, the bolts 76 being replaced by crimping elements 76b.
  • the cooling being able to be carried out through elements crimping, it is no longer necessary to provide calibrated orifices at the lamellar seal 80, 82.
  • the metal crown 66a forming a flange connecting together the first ends by brazing (or welding) 62 of the fixing lugs 58 of the external axial wall of the combustion 26 is no longer mounted between flanges but itself brazed (or welded) at a centered key 106 and resting on the annular casing external 12.
  • the metal crown 66b forming a flange connecting together the first ends by brazing (or welding) 64 of the fixing lugs 60 of the internal axial wall of the combustion chamber 28 is no longer mounted between flanges but simply fixed directly to the internal annular envelope 14 by conventional fixing means 108, for example of the bolt type.
  • the flexibility of the fixing lugs supports the thermal expansion gap occurring at temperatures between the composite material combustion chamber and the casings metallic annulars while ensuring the maintenance and positioning of the bedroom.

Landscapes

  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Ceramic Engineering (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
  • Chimneys And Flues (AREA)
  • Gasket Seals (AREA)
EP02291364A 2001-06-06 2002-06-04 Doppelbefestigung einer Turbinenbrennkammer aus keramischem Matrix-Verbundwerkstoff Expired - Lifetime EP1265035B1 (de)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
FR0107372A FR2825785B1 (fr) 2001-06-06 2001-06-06 Liaison de chambre de combustion cmc de turbomachine en deux parties
FR0107372 2001-06-06

Publications (2)

Publication Number Publication Date
EP1265035A1 true EP1265035A1 (de) 2002-12-11
EP1265035B1 EP1265035B1 (de) 2008-02-13

Family

ID=8863994

Family Applications (1)

Application Number Title Priority Date Filing Date
EP02291364A Expired - Lifetime EP1265035B1 (de) 2001-06-06 2002-06-04 Doppelbefestigung einer Turbinenbrennkammer aus keramischem Matrix-Verbundwerkstoff

Country Status (5)

Country Link
US (1) US6675585B2 (de)
EP (1) EP1265035B1 (de)
JP (1) JP4097994B2 (de)
DE (1) DE60224956T2 (de)
FR (1) FR2825785B1 (de)

Cited By (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP2888452B1 (de) * 2012-06-28 2019-11-06 Safran Aircraft Engines Gasturbinentriebwerk mit einem teil aus faserverbundwerkstoff und einem metallischen teil, beide teile verbunden mit einer weichen befestigung

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* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN1246638C (zh) * 2001-04-27 2006-03-22 西门子公司 燃烧室、尤其是燃气轮机的燃烧室
FR2840974B1 (fr) * 2002-06-13 2005-12-30 Snecma Propulsion Solide Anneau d'etancheite pour cahmbre de combustion et chambre de combustion comportant un tel anneau
US6895761B2 (en) * 2002-12-20 2005-05-24 General Electric Company Mounting assembly for the aft end of a ceramic matrix composite liner in a gas turbine engine combustor
FR2855249B1 (fr) * 2003-05-20 2005-07-08 Snecma Moteurs Chambre de combustion ayant une liaison souple entre un fond de chambre et une paroi de chambre
FR2860039B1 (fr) * 2003-09-19 2005-11-25 Snecma Moteurs Realisation de l'etancheite dans un turboreacteur pour le prelevement cabine par joints double sens a lamelles
FR2871846B1 (fr) * 2004-06-17 2006-09-29 Snecma Moteurs Sa Chambre de combustion en cmc de turbine a gaz supportee dans un carter metallique par des organes de liaison en cmc
FR2871845B1 (fr) * 2004-06-17 2009-06-26 Snecma Moteurs Sa Montage de chambre de combustion de turbine a gaz avec distributeur integre de turbine haute pression
FR2871847B1 (fr) * 2004-06-17 2006-09-29 Snecma Moteurs Sa Montage d'un distributeur de turbine sur une chambre de combustion a parois en cmc dans une turbine a gaz
US7421842B2 (en) * 2005-07-18 2008-09-09 Siemens Power Generation, Inc. Turbine spring clip seal
FR2892181B1 (fr) * 2005-10-18 2008-02-01 Snecma Sa Fixation d'une chambre de combustion a l'interieur de son carter
US7775050B2 (en) * 2006-10-31 2010-08-17 General Electric Company Method and apparatus for reducing stresses induced to combustor assemblies
FR2930628B1 (fr) * 2008-04-24 2010-04-30 Snecma Chambre annulaire de combustion pour turbomachine
FR2935753B1 (fr) * 2008-09-08 2011-07-01 Snecma Propulsion Solide Liaisons souples a butee pour fixation de piece en cmc
US9234431B2 (en) * 2010-07-20 2016-01-12 Siemens Energy, Inc. Seal assembly for controlling fluid flow
US8322141B2 (en) * 2011-01-14 2012-12-04 General Electric Company Power generation system including afirst turbine stage structurally incorporating a combustor
US9335051B2 (en) * 2011-07-13 2016-05-10 United Technologies Corporation Ceramic matrix composite combustor vane ring assembly
FR2989426B1 (fr) * 2012-04-11 2014-03-28 Snecma Turbomachine, telle qu'un turboreacteur ou un turbopropulseur d'avion
JP6228685B2 (ja) 2013-09-11 2017-11-08 ゼネラル・エレクトリック・カンパニイ ばねで荷重されシールされるセラミックマトリックス複合材燃焼器ライナ
WO2015038293A1 (en) 2013-09-11 2015-03-19 United Technologies Corporation Combustor liner
US10281153B2 (en) * 2016-02-25 2019-05-07 General Electric Company Combustor assembly
US10378771B2 (en) 2016-02-25 2019-08-13 General Electric Company Combustor assembly
US10519811B2 (en) * 2016-10-04 2019-12-31 United Technologies Corporation Flange heat shield
US10550725B2 (en) * 2016-10-19 2020-02-04 United Technologies Corporation Engine cases and associated flange
ES2760550T3 (es) * 2017-04-07 2020-05-14 MTU Aero Engines AG Disposición de junta para una turbina de gas
FR3084731B1 (fr) * 2019-02-19 2020-07-03 Safran Aircraft Engines Chambre de combustion pour une turbomachine
CN110822482B (zh) * 2019-11-28 2020-10-27 中国航发沈阳黎明航空发动机有限责任公司 一种中低热值气体和液体双燃料喷嘴及燃料切换方法
CN114413285B (zh) * 2022-01-29 2023-03-21 中国航发湖南动力机械研究所 一种大弯管密封结构
CN115405370B (zh) * 2022-11-03 2023-03-10 中国航发沈阳发动机研究所 一种半弹性涡轮外环结构

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US2509503A (en) * 1946-02-12 1950-05-30 Lucas Ltd Joseph Combustion chamber for prime movers
US4030875A (en) * 1975-12-22 1977-06-21 General Electric Company Integrated ceramic-metal combustor
GB2035474A (en) * 1978-11-09 1980-06-18 Sulzer Ag Seals
GB1570875A (en) * 1977-03-16 1980-07-09 Lucas Industries Ltd Combustion equipment
US5524430A (en) * 1992-01-28 1996-06-11 Societe National D'etude Et De Construction De Moteurs D'aviation S.N.E.C.M.A. Gas-turbine engine with detachable combustion chamber
EP1035377A2 (de) * 1999-03-08 2000-09-13 Mitsubishi Heavy Industries, Ltd. Abdichtung für das Endstück einer Gasturbinenbrennkammer
US6131384A (en) * 1997-10-16 2000-10-17 Rolls-Royce Deutschland Gmbh Suspension device for annular gas turbine combustion chambers

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US5181377A (en) * 1991-04-16 1993-01-26 General Electric Company Damped combustor cowl structure
CA2070511C (en) * 1991-07-22 2001-08-21 Steven Milo Toborg Turbine nozzle support
US5291733A (en) * 1993-02-08 1994-03-08 General Electric Company Liner mounting assembly
US5564271A (en) * 1994-06-24 1996-10-15 United Technologies Corporation Pressure vessel fuel nozzle support for an industrial gas turbine engine
US6397603B1 (en) * 2000-05-05 2002-06-04 The United States Of America As Represented By The Secretary Of The Air Force Conbustor having a ceramic matrix composite liner
US6334298B1 (en) * 2000-07-14 2002-01-01 General Electric Company Gas turbine combustor having dome-to-liner joint
US6497104B1 (en) * 2000-10-30 2002-12-24 General Electric Company Damped combustion cowl structure

Patent Citations (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2509503A (en) * 1946-02-12 1950-05-30 Lucas Ltd Joseph Combustion chamber for prime movers
US4030875A (en) * 1975-12-22 1977-06-21 General Electric Company Integrated ceramic-metal combustor
GB1570875A (en) * 1977-03-16 1980-07-09 Lucas Industries Ltd Combustion equipment
GB2035474A (en) * 1978-11-09 1980-06-18 Sulzer Ag Seals
US5524430A (en) * 1992-01-28 1996-06-11 Societe National D'etude Et De Construction De Moteurs D'aviation S.N.E.C.M.A. Gas-turbine engine with detachable combustion chamber
US6131384A (en) * 1997-10-16 2000-10-17 Rolls-Royce Deutschland Gmbh Suspension device for annular gas turbine combustion chambers
EP1035377A2 (de) * 1999-03-08 2000-09-13 Mitsubishi Heavy Industries, Ltd. Abdichtung für das Endstück einer Gasturbinenbrennkammer

Cited By (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP2888452B1 (de) * 2012-06-28 2019-11-06 Safran Aircraft Engines Gasturbinentriebwerk mit einem teil aus faserverbundwerkstoff und einem metallischen teil, beide teile verbunden mit einer weichen befestigung

Also Published As

Publication number Publication date
DE60224956D1 (de) 2008-03-27
US20020184888A1 (en) 2002-12-12
JP4097994B2 (ja) 2008-06-11
JP2003035418A (ja) 2003-02-07
EP1265035B1 (de) 2008-02-13
FR2825785B1 (fr) 2004-08-27
US6675585B2 (en) 2004-01-13
FR2825785A1 (fr) 2002-12-13
DE60224956T2 (de) 2009-02-05

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