EP1265034A1 - Befestigung einer Turbinenbrennkammer aus keramischem Matrix-Verbundswerkstoff mit gelöteten Befestigungsfüssen - Google Patents
Befestigung einer Turbinenbrennkammer aus keramischem Matrix-Verbundswerkstoff mit gelöteten Befestigungsfüssen Download PDFInfo
- 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
Links
Images
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)
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
ID=8863987
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)
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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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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 |
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 |
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 |
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 |
EP1960636B1 (de) * | 2005-12-14 | 2016-01-27 | 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 |
US20090067917A1 (en) * | 2007-09-07 | 2009-03-12 | The Boeing Company | Bipod Flexure Ring |
US8726675B2 (en) * | 2007-09-07 | 2014-05-20 | The Boeing Company | Scalloped 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 |
GB2460634B (en) * | 2008-06-02 | 2010-07-07 | Rolls Royce Plc | Combustion apparatus |
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 |
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US10823410B2 (en) | 2016-10-26 | 2020-11-03 | Raytheon Technologies Corporation | Cast combustor liner panel radius for gas turbine engine combustor |
US10670269B2 (en) | 2016-10-26 | 2020-06-02 | Raytheon Technologies Corporation | Cast combustor liner panel gating feature for a gas turbine engine combustor |
US10669939B2 (en) | 2016-10-26 | 2020-06-02 | Raytheon Technologies Corporation | Combustor seal for a gas turbine engine combustor |
US10830448B2 (en) | 2016-10-26 | 2020-11-10 | Raytheon Technologies Corporation | Combustor liner panel with a multiple of heat transfer augmentors for a gas turbine engine combustor |
US10935243B2 (en) | 2016-11-30 | 2021-03-02 | Raytheon Technologies Corporation | Regulated combustor liner panel for a gas turbine engine combustor |
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US10385731B2 (en) * | 2017-06-12 | 2019-08-20 | General Electric Company | CTE matching hanger support for CMC structures |
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US11377970B2 (en) | 2018-11-02 | 2022-07-05 | Chromalloy Gas Turbine Llc | System and method for providing compressed air to a gas turbine combustor |
US11248797B2 (en) * | 2018-11-02 | 2022-02-15 | Chromalloy Gas Turbine Llc | Axial stop configuration for a combustion liner |
FR3111964B1 (fr) | 2020-06-26 | 2023-03-17 | Safran Helicopter Engines | Assemblage d’une pièce de chambre de combustion par recouvrement par une autre pièce |
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US2509503A (en) * | 1946-02-12 | 1950-05-30 | Lucas Ltd Joseph | Combustion chamber for prime movers |
GB2035474A (en) * | 1978-11-09 | 1980-06-18 | Sulzer Ag | Seals |
GB1570875A (en) * | 1977-03-16 | 1980-07-09 | Lucas Industries Ltd | Combustion equipment |
EP0316233A1 (de) * | 1987-11-12 | 1989-05-17 | Societe Nationale D'etude Et De Construction De Moteurs D'aviation "Snecma" | Methode zum Zusammenfügen von zwei Teilen mit unterschiedlichem Wärmeausdehnungskoeffizienten |
US6131384A (en) * | 1997-10-16 | 2000-10-17 | Rolls-Royce Deutschland Gmbh | Suspension device for annular gas turbine combustion chambers |
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FR316233A (de) | ||||
US2509593A (en) | 1947-05-21 | 1950-05-30 | Rca Corp | Humidity compensated oscillator |
JPS52158202U (de) * | 1976-05-27 | 1977-12-01 | ||
JP2597800B2 (ja) * | 1992-06-12 | 1997-04-09 | ゼネラル・エレクトリック・カンパニイ | ガスタービンエンジン用燃焼器 |
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 |
-
2001
- 2001-06-06 FR FR0107363A patent/FR2825783B1/fr not_active Expired - Fee Related
-
2002
- 2002-06-03 JP JP2002161064A patent/JP3907529B2/ja not_active Expired - Fee Related
- 2002-06-04 EP EP02291363A patent/EP1265034B1/de not_active Expired - Lifetime
- 2002-06-04 DE DE60229465T patent/DE60229465D1/de not_active Expired - Lifetime
- 2002-06-05 US US10/162,385 patent/US6708495B2/en not_active Expired - Lifetime
Patent Citations (5)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US2509503A (en) * | 1946-02-12 | 1950-05-30 | Lucas Ltd Joseph | Combustion chamber for prime movers |
GB1570875A (en) * | 1977-03-16 | 1980-07-09 | Lucas Industries Ltd | Combustion equipment |
GB2035474A (en) * | 1978-11-09 | 1980-06-18 | Sulzer Ag | Seals |
EP0316233A1 (de) * | 1987-11-12 | 1989-05-17 | Societe Nationale D'etude Et De Construction De Moteurs D'aviation "Snecma" | Methode zum Zusammenfügen von zwei Teilen mit unterschiedlichem Wärmeausdehnungskoeffizienten |
US6131384A (en) * | 1997-10-16 | 2000-10-17 | Rolls-Royce Deutschland Gmbh | Suspension device for annular gas turbine combustion chambers |
Cited By (5)
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 |
---|---|
FR2825783A1 (fr) | 2002-12-13 |
EP1265034B1 (de) | 2008-10-22 |
DE60229465D1 (de) | 2008-12-04 |
FR2825783B1 (fr) | 2003-11-07 |
US20020184892A1 (en) | 2002-12-12 |
JP3907529B2 (ja) | 2007-04-18 |
US6708495B2 (en) | 2004-03-23 |
JP2003014234A (ja) | 2003-01-15 |
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