EP1265034A1 - Befestigung einer Turbinenbrennkammer aus keramischem Matrix-Verbundswerkstoff mit gelöteten Befestigungsfüssen - Google Patents

Befestigung einer Turbinenbrennkammer aus keramischem Matrix-Verbundswerkstoff mit gelöteten Befestigungsfüssen Download PDF

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Publication number
EP1265034A1
EP1265034A1 EP02291363A EP02291363A EP1265034A1 EP 1265034 A1 EP1265034 A1 EP 1265034A1 EP 02291363 A EP02291363 A EP 02291363A EP 02291363 A EP02291363 A EP 02291363A EP 1265034 A1 EP1265034 A1 EP 1265034A1
Authority
EP
European Patent Office
Prior art keywords
combustion chamber
crown
turbomachine according
metal
composite
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
EP02291363A
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English (en)
French (fr)
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EP1265034B1 (de
Inventor
Gwénaelle Calvez
Didier Hernandez
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
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 filed Critical SNECMA Moteurs SA
Publication of EP1265034A1 publication Critical patent/EP1265034A1/de
Application granted granted Critical
Publication of EP1265034B1 publication Critical patent/EP1265034B1/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 housings of a chamber turbine engine.
  • CMC type composite material combustion chamber ceramic matrix composite
  • the turbine high pressure in particular its inlet distributor (HPT nozzle), the combustion as well as the inner and outer casings (or casings) of this chamber are made of the same material, generally metallic.
  • HPT nozzle inlet distributor
  • the combustion as well as the inner and outer casings (or casings) of this chamber are made of the same material, generally metallic.
  • the use of a chamber metal proves from a thermal point of view totally unsuitable and it must be used a chamber based on high temperature composite materials such as CMC.
  • the difficulties of implementation and the cost of these materials make that their use is most often limited to the combustion chamber it same, the inlet distributor of the high pressure turbine and the casings internal and external of the chamber then remaining produced more conventionally in metallic materials.
  • Gold, metallic materials and composite materials have very different coefficients of thermal expansion. This results in particularly acute problems of connection with the internal and external envelopes of the combustion and interface 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 crankcases having the capacity to absorb the displacements induced by the differences in the expansion coefficients of these rooms.
  • An object of the invention is also to propose an arrangement which benefits the better existing characteristics of the combustion chamber.
  • a turbomachine comprising, in internal and external annular envelopes of metallic material and in a direction F gas flow, a fuel injection assembly, a annular combustion of composite material having a longitudinal axis, and a annular distributor made of metallic material forming the paddle inlet stage stationary of a high pressure turbine, characterized in that said combustion of composite material is maintained in position between said internal and external metallic annular envelopes by a plurality of flexible tongues, the first ends of said tongues being connected between them by a metal crown fixed integrally to each of said internal and external metallic annular envelopes by first means of fixing and second ends being fixed by second means of attachment to a crown of composite material fixedly attached to said combustion chamber of composite material, the flexibility of said tabs of fixing allowing at high temperatures a free radial expansion of said combustion chamber of composite material with respect to said casings metallic annulars.
  • the first and second fixing means preferably consist by a plurality of bolts.
  • each of said metallic ring envelopes is formed in two parts
  • said crown metallic connecting together said first ends of said tabs of metal fixing is mounted between the connecting flanges of these two parts.
  • said metal crown can be fixed directly to said annular casing by fixing means.
  • said first ends of the fixing tabs can either be fixed by soldering to said crown either be a single piece with this metal crown.
  • said composite crown is brazed on a downstream end of the combustion chamber.
  • this composite crown is sewn on this downstream end.
  • this composite crown is installed on this downstream end.
  • Said composite crown comprises a determined part forming a support surface for a seal (advantageously of the circular seal type "Lamellae") sealing the gas stream between said chamber combustion and said distributor.
  • said determined part is a end part of said composite 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 between the annular casings internal and external by a plurality of flexible tabs 58, 60 regularly distributed around the combustion chamber.
  • These mounting tabs are mounted for a first part of them (see tab referenced 58) between the outer annular envelope 12a, 12b and the outer axial wall 26 of the combustion chamber and for a second part (like the tongue 60) between the internal annular envelope 14a, 14b and the internal axial wall 28 of the chamber combustion.
  • Each flexible fixing tab made of metallic material which can have a substantially triangular shape as illustrated in FIG. 1A, or consist of a simple blade (of constant width or not), is welded or brazed by a first end 62; 64 to a metal crown 66a, 66b fixed together by first fixing means 52; 68 to either (depending on its location) external metallic annular envelopes 12 or internal 14 and intended to facilitate both the maintenance of these tabs and sealing against the annular space 16.
  • these tongues and the metal crown together form a part single metal in one piece.
  • this tongue is fixed integrally by second fixing means 74, 76 to a ceramic composite crown 78a; 78b brazed on a downstream end 88; 90 of external axial 26 and internal 28 walls of the material combustion chamber ceramic composite.
  • This solder can be replaced or reinforced by a sewing.
  • the connection between the chamber walls and the crowns can also be performed entirely by implantation (type bond known as anglemia "Pin'sée").
  • the number of tabs can, for example, be in number equal to that of the injection nozzles or equal to a multiple of this number.
  • FIG. 1 illustrates a first embodiment of the invention in which the second ends of the tabs 70, 72 are fixed respectively on the composite ceramic outer 78a and inner 78b crowns with a simple bolting (but crimping as illustrated in the partial view of Figure 1B would also be possible).
  • the metal crown 66a, 66b connecting the first ends 62, 64 of the tongues is preferably taken between existing connection flanges between the upstream and downstream parts of the enclosures internal 14 and external 12 annulars and fixedly held by the first fixing means 52, 68 which preferably are also of the bolt type.
  • washers made of ceramic composite material 74a; 76a for allow to "drown" the conical heads of the bolts screws forming the second fixing means 74; 76.
  • the metal ring 66a connecting between them by welding (or brazing) the first ends 62 of the tabs of fixing 58 of the external axial wall of the combustion chamber 26 is no longer mounted between flanges but itself welded (or brazed) at a key 106 integral with the outer annular casing 12.
  • the metal crown 66b connecting together by welding (or brazing) the first ends 64 of fastening tabs 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 casing 14 by fixing means 108, for example of 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 a part of the crown ceramic composite 78a; 78b forming a support plane for this circular joint sealing.
  • This part can be an end part of the crown.
  • Gasket is kept in abutment against this end part of the composite crown or any other part by means of an elastic element, of the circular spring type with blades 92, 94, fixed on the distributor. This provision ensures a perfect tightness of the hot stream between the combustion chamber 24 and the distributor 42.
  • the flexibility of the tabs of fixing allows to support the thermal expansion gap appearing at high temperatures between the composite material combustion chamber and the metallic ring envelopes while ensuring the maintenance and positioning of the chamber.

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)
EP02291363A 2001-06-06 2002-06-04 Befestigung einer Turbinenbrennkammer aus keramischem Matrix-Verbundswerkstoff mit gelöteten Befestigungsfüssen Expired - Lifetime EP1265034B1 (de)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
FR0107363A FR2825783B1 (fr) 2001-06-06 2001-06-06 Accrochage de chambre de combustion cmc de turbomachine par pattes brasees
FR0107363 2001-06-06

Publications (2)

Publication Number Publication Date
EP1265034A1 true EP1265034A1 (de) 2002-12-11
EP1265034B1 EP1265034B1 (de) 2008-10-22

Family

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Family Applications (1)

Application Number Title Priority Date Filing Date
EP02291363A Expired - Lifetime EP1265034B1 (de) 2001-06-06 2002-06-04 Befestigung einer Turbinenbrennkammer aus keramischem Matrix-Verbundswerkstoff mit gelöteten Befestigungsfüssen

Country Status (5)

Country Link
US (1) US6708495B2 (de)
EP (1) EP1265034B1 (de)
JP (1) JP3907529B2 (de)
DE (1) DE60229465D1 (de)
FR (1) FR2825783B1 (de)

Cited By (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
WO2010116051A3 (fr) * 2009-04-07 2011-05-19 Snecma Turbomachine a chambre annulaire de combustion
FR2976021A1 (fr) * 2011-05-30 2012-12-07 Snecma Turbomachine a chambre annulaire de combustion
FR3010774A1 (fr) * 2013-09-16 2015-03-20 Snecma Turbomachine a chambre de combustion maintenue par une couronne de fixation metallique
CN105298684A (zh) * 2015-09-18 2016-02-03 中国航空工业集团公司沈阳发动机设计研究所 一种航空发动机用尾椎连接结构

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US10839321B2 (en) * 1997-01-06 2020-11-17 Jeffrey Eder Automated data storage system
JP3600912B2 (ja) * 2001-09-12 2004-12-15 川崎重工業株式会社 燃焼器ライナのシール構造
EP1312865A1 (de) * 2001-11-15 2003-05-21 Siemens Aktiengesellschaft Ringbrennkammer für eine Gasturbine
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
US7047722B2 (en) * 2002-10-02 2006-05-23 Claudio Filippone Small scale hybrid engine (SSHE) utilizing fossil fuels
US6775985B2 (en) * 2003-01-14 2004-08-17 General Electric Company Support assembly for 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
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
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
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
US7197877B2 (en) * 2004-08-04 2007-04-03 Siemens Power Generation, Inc. Support system for a pilot nozzle of a turbine engine
US7647779B2 (en) * 2005-04-27 2010-01-19 United Technologies Corporation Compliant metal support for ceramic combustor liner in a gas turbine engine
EP1731715A1 (de) * 2005-06-10 2006-12-13 Siemens Aktiengesellschaft Übergangsbereich zwischen einer Brennkammer und einer Turbineneinheit
FR2892181B1 (fr) * 2005-10-18 2008-02-01 Snecma Sa Fixation d'une chambre de combustion a l'interieur de son carter
WO2007068538A1 (de) * 2005-12-14 2007-06-21 Alstom Technology Ltd Strömungsmaschine
US7578134B2 (en) * 2006-01-11 2009-08-25 General Electric Company Methods and apparatus for assembling gas turbine engines
US8863528B2 (en) * 2006-07-27 2014-10-21 United Technologies Corporation Ceramic combustor can for a gas turbine engine
US8141370B2 (en) * 2006-08-08 2012-03-27 General Electric Company Methods and apparatus for radially compliant component mounting
US8556531B1 (en) * 2006-11-17 2013-10-15 United Technologies Corporation Simple CMC fastening system
US8726675B2 (en) * 2007-09-07 2014-05-20 The Boeing Company Scalloped flexure ring
US20090067917A1 (en) * 2007-09-07 2009-03-12 The Boeing Company Bipod Flexure Ring
GB2453946B (en) * 2007-10-23 2010-07-14 Rolls Royce Plc A Wall Element for use in Combustion Apparatus
GB0800294D0 (en) * 2008-01-09 2008-02-20 Rolls Royce Plc Gas heater
GB0801839D0 (en) * 2008-02-01 2008-03-05 Rolls Royce Plc combustion apparatus
GB2457281B (en) * 2008-02-11 2010-09-08 Rolls Royce Plc A Combustor Wall Arrangement with Parts Joined by Mechanical Fasteners
FR2929689B1 (fr) * 2008-04-03 2013-04-12 Snecma Propulsion Solide Chambre de combustion de turbine a gaz a parois interne et externe sectorisees
FR2929690B1 (fr) * 2008-04-03 2012-08-17 Snecma Propulsion Solide Chambre de combustion sectorisee en cmc pour turbine a gaz
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FR2935753B1 (fr) * 2008-09-08 2011-07-01 Snecma Propulsion Solide Liaisons souples a butee pour fixation de piece en cmc
US8322983B2 (en) * 2008-09-11 2012-12-04 Siemens Energy, Inc. Ceramic matrix composite structure
US8266914B2 (en) * 2008-10-22 2012-09-18 Pratt & Whitney Canada Corp. Heat shield sealing for gas turbine engine combustor
US8388307B2 (en) * 2009-07-21 2013-03-05 Honeywell International Inc. Turbine nozzle assembly including radially-compliant spring member for gas turbine engine
US8215115B2 (en) * 2009-09-28 2012-07-10 Hamilton Sundstrand Corporation Combustor interface sealing arrangement
US9335051B2 (en) * 2011-07-13 2016-05-10 United Technologies Corporation Ceramic matrix composite combustor vane ring assembly
JP5762558B2 (ja) * 2011-11-16 2015-08-12 三菱日立パワーシステムズ株式会社 ガスタービン燃焼器
RU2497251C1 (ru) * 2012-03-30 2013-10-27 Открытое акционерное общество "Уфимское научно-производственное предприятие "Молния" (ОАО УНПП "Молния") Свеча зажигания для камер сгорания энергетических и двигательных установок
EP2692995B1 (de) * 2012-07-30 2017-09-20 Ansaldo Energia IP UK Limited Stationäre Gasturbinenmotor und Verfahren zur Durchführung von Instandhaltungsarbeit
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US20170059159A1 (en) 2015-08-25 2017-03-02 Rolls-Royce Corporation Cmc combustor shell with integral chutes
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Cited By (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
WO2010116051A3 (fr) * 2009-04-07 2011-05-19 Snecma Turbomachine a chambre annulaire de combustion
FR2976021A1 (fr) * 2011-05-30 2012-12-07 Snecma Turbomachine a chambre annulaire de combustion
FR3010774A1 (fr) * 2013-09-16 2015-03-20 Snecma Turbomachine a chambre de combustion maintenue par une couronne de fixation metallique
CN105298684A (zh) * 2015-09-18 2016-02-03 中国航空工业集团公司沈阳发动机设计研究所 一种航空发动机用尾椎连接结构
CN105298684B (zh) * 2015-09-18 2017-11-03 中国航空工业集团公司沈阳发动机设计研究所 一种航空发动机用尾椎连接结构

Also Published As

Publication number Publication date
US20020184892A1 (en) 2002-12-12
FR2825783A1 (fr) 2002-12-13
EP1265034B1 (de) 2008-10-22
JP2003014234A (ja) 2003-01-15
US6708495B2 (en) 2004-03-23
FR2825783B1 (fr) 2003-11-07
JP3907529B2 (ja) 2007-04-18
DE60229465D1 (de) 2008-12-04

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