EP1265034B1 - 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
EP1265034B1
EP1265034B1 EP02291363A EP02291363A EP1265034B1 EP 1265034 B1 EP1265034 B1 EP 1265034B1 EP 02291363 A EP02291363 A EP 02291363A EP 02291363 A EP02291363 A EP 02291363A EP 1265034 B1 EP1265034 B1 EP 1265034B1
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EP
European Patent Office
Prior art keywords
metal
combustion chamber
ring
turbomachine according
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.)
Expired - Lifetime
Application number
EP02291363A
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English (en)
French (fr)
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EP1265034A1 (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 SAS
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Filing date
Publication date
Application filed by SNECMA SAS filed Critical SNECMA SAS
Publication of EP1265034A1 publication Critical patent/EP1265034A1/de
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Anticipated expiration legal-status Critical
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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 turbomachines and is more particularly concerned with the problem of mounting a combustion chamber made of CMC (ceramic matrix composite) type composite material in the metal chamber housings of a turbine engine.
  • CMC ceramic matrix composite
  • Such a turbomachine is known from GB 1,570,875 .
  • the high pressure turbine including its inlet nozzle (HPT nozzle), the combustion chamber and the envelopes (or housings) internal and external of this chamber are made of the same material, usually metallic.
  • HPT nozzle inlet nozzle
  • the combustion chamber and the envelopes (or housings) internal and external of this chamber are made of the same material, usually metallic.
  • metallic metallic
  • the use of a metal chamber is from a thermal point of view totally inadequate and it must be resorted to a chamber based on CMC type high temperature composite materials.
  • the difficulties of implementation and the cost of these materials mean that their use is most often limited to the combustion chamber itself, the inlet valve of the high pressure turbine and the inner and outer shells of the chamber remaining then more typically made of metal materials.
  • metal materials and composite materials have very different coefficients of thermal expansion. This results in particularly acute problems of connection with the inner and outer casings of the combustion chamber and interface at the distributor, at the inlet of the high pressure turbine.
  • the present invention overcomes these disadvantages by proposing a mounting of the combustion chamber in the casings having the capacity to absorb the displacements induced by the differences in the expansion coefficients of these parts.
  • An object of the invention is also to provide an assembly that best takes advantage of the existing characteristics of the combustion chamber.
  • a turbomachine comprising, in inner and outer annular envelopes made of metallic material and in a direction F of gas flow, a fuel injection assembly, an annular combustion chamber made of composite material having a longitudinal axis, and an annular distributor of metallic material forming the fixed blade inlet stage of a high pressure turbine, said combustion chamber of composite material being held in position between said inner and outer metal annular envelopes by a plurality of flexible tongues, first ends of said tongues being interconnected by a metal ring fixed integrally to one of said inner and outer metal annular envelopes by first fixing means, characterized in that second ends of said tongues are fixed by second means of fastening to a crown in co material mposite fixed integrally to said combustion chamber of composite material, the flexibility of said fastening tabs allowing high temperatures radial free expansion of said composite material combustion chamber relative to said metal annular envelopes.
  • the first and second fastening means are preferably constituted by a plurality of bolts.
  • each of said annular metal envelopes is formed in two parts
  • said crown metal interconnecting said first ends of said metal fastening tabs is mounted between connecting flanges of these two parts.
  • said metal ring may be fixed directly to said annular envelope by fixing means.
  • said first ends of the fixing tongues may either be soldered to said metal ring or form a single piece with this metal ring.
  • said composite ring is brazed on a downstream end of the combustion chamber.
  • this composite crown is sewn on this downstream end.
  • this composite ring is located on this downstream end.
  • Said composite ring comprises a specific part forming a support plane for a seal (advantageously of the type "circular seal") ensuring the tightness of the gas stream between said combustion chamber and said distributor.
  • said determined portion is an end portion of said composite ring.
  • the distributor is fixed on a downstream portion 14b of the inner annular envelope of the turbomachine by first removable fastening means preferably constituted by a plurality of bolts 50 while resting on support means 49 integral with the outer annular envelope of the turbomachine.
  • Through-holes 54, 56 formed in the outer metal 46 and inner 48 metal platforms of the distributor 42 are furthermore provided for cooling the vanes 44 of the distributor at the inlet of the rotor of the high-pressure turbine from the oxidant compressed available at the outlet of the diffusion duct 18 and flowing in two flows F1, F2 on either side of the combustion chamber 24.
  • the combustion chamber 24 which has a coefficient of thermal expansion very different from the other metal parts forming the turbomachine, is fixedly held in position between the inner and outer annular envelopes by a plurality of flexible tongues 58, 60 regularly distributed around the combustion chamber.
  • These fixing tabs are mounted for a first part of them (see the tab referenced 58) between the outer annular casing 12a, 12b and the outer axial wall 26 of the combustion chamber and for a second part (like the tongue 60) between the inner annular envelope 14a, 14b and the inner axial wall 28 of the combustion chamber.
  • Each flexible fastening tab made of metallic material which can have a substantially triangular shape as illustrated by FIG. Figure 1A , or consist of a single blade (of constant width or not), is welded or soldered by a first end 62; 64 to a metal ring 66a, 66b secured integrally by first fixing means 52; 68 to one or the other (depending on its location) outer metal annular envelopes 12 or internal 14 and intended to facilitate both the maintenance of these tabs and the sealing with respect to the annular space 16.
  • these tabs and the metal ring together form a single piece of metal in one piece.
  • this tongue is fixed integrally by second fixing means 74, 76 to a ceramic composite ring 78a; 78b soldered on a downstream end 88; 90 of the outer axial walls 26 and inner 28 of the combustion chamber made of ceramic composite material.
  • 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 (connection of known type under the anglicism "pin'sée").
  • the number of tongues may, for example, be in number equal to that of the injection nozzles or equal to a multiple of this number.
  • the figure 1 illustrates a first embodiment of the invention in which the second ends of the tabs 70, 72 are respectively fixed on the outer ceramic cylindrical 78a and inner 78b crowns by a simple bolting (but crimping as shown in the partial view of the Figure 1B would also be possible).
  • the metal ring 66a, 66b interconnecting the first ends 62, 64 of the tabs is in turn preferably taken between existing connecting flanges between the upstream and downstream parts of the inner annular casing 14 and external 12 and fixedly held by the first fixing means 52, 68 which preferably are also bolt type.
  • the metal ring 66a interconnecting by welding (or brazing) the first ends 62 of the fastening tabs 58 of the outer axial wall of the combustion chamber 26 is no longer mounted between flanges but itself welded (or brazed) to level of a polarizer 106 secured to the outer annular casing 12.
  • the metal ring 66b interconnecting by welding (or brazing) the first ends 64 of the fixing tongues 60 of the internal axial wall of the combustion chamber 28 is no longer mounted between flanges but simply fixed directly to the annular envelope internal 14 by fastening means 108, for example bolt type.
  • the tightness of the gas stream between the combustion chamber 24 and the distributor 42 is provided by a "flap" circular seal 80, 82 mounted in a groove 84, 86 of each of the outer 46 and inner 48 platforms. distributor and which bears directly on a portion of the ceramic composite ring 78a; 78b forming a support plane for this seal sealing. This part may be an end portion of the crown.
  • the seal is held in abutment against this end portion of the composite ring or any other portion by means of a resilient member, circular leaf spring type 92, 94, fixed on the distributor.
  • an omega-type circular seal 96 mounted in a circular groove 98 of a flange of the inner annular casing 14 in direct contact with the inner circular platform 48 of the distributor and secondly by another "flap" circular seal 100 mounted in a circular groove 102 of the outer circular platform of the distributor 46 and one end is in direct contact with a circular spoiler 104 of the downstream portion 12b of the outer annular envelope.
  • the flexibility of the fastening tabs can withstand the thermal expansion gap occurring at high temperatures between the composite material combustion chamber and the metal annular envelopes while ensuring the maintenance and positioning of the chamber.

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  • 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)

Claims (12)

  1. Turbomaschine, die - in einem ringförmigen Innen- und einem ringförmigen Außengehäuse aus metallischem Material (12, 14) und entlang einer Strömungsrichtung F der Gase - eine Einheit zum Einspritzen eines Treibstoffs (20, 22), eine ringförmige Brennkammer aus Verbundwerkstoff (24) mit einer Längsachse (10) sowie ein ringförmiges Leitrad aus metallischem Material (42), welches die Eintrittsstufe mit Leitschaufeln (44) einer Hochdruckturbine bildet, umfaßt, wobei die Brennkammer aus Verbundwerkstoff zwischen dem ringförmigen Metallinnengehäuse und dem ringförmigen Metallaußengehäuse durch eine Vielzahl von flexiblen Metallzungen (58, 60) in Position gehalten ist, wobei erste Enden (62, 64) der Zungen durch einen Metallkranz (66a, 66b), der mit einem der ringförmigen metallischen Innen- und Außengehäuse (12, 14) durch erste Befestigungsmittel (52; 68, 108) fest verbunden ist, untereinander verbunden sind, dadurch gekennzeichnet, daß zweite Enden (70, 72) der Zungen durch zweite Befestigungsmittel (74, 76) an einem Kranz aus Verbundwerkstoff (78a, 78b), der mit der Brennkammer aus Verbundwerkstoff (26, 28) fest verbunden ist, befestigt sind, wobei die Flexibilität der metallischen Befestigungszungen bei hohen Temperaturen eine freie radiale Ausdehnung der Brennkammer aus Verbundwerkstoff gegenüber den ringförmige Metallgehäusen ermöglicht.
  2. Turbomaschine nach Anspruch 1, dadurch gekennzeichnet, daß die ersten und zweiten Befestigungsmittel von einer Vielzahl von Bolzen gebildet sind.
  3. Turbomaschine nach Anspruch 1, dadurch gekennzeichnet, daß - da ein jedes der ringförmigen Metallgehäuse aus zwei Teilen (12a, 12b; 14a, 14b) besteht - der Metallkranz, welcher die ersten Enden der metallischen Befestigungszungen untereinander verbindet, zwischen Verbindungsflanschen dieser beiden Teile angebracht ist.
  4. Turbomaschine nach Anspruch 1, dadurch gekennzeichnet, daß der Metallkranz, welcher die ersten Enden der metallischen Befestigungszungen untereinander verbindet, durch Befestigungsmittel (108) direkt an dem ringförmige Gehäuse befestigt ist.
  5. Turbomaschine nach Anspruch 1, dadurch gekennzeichnet, daß die ersten Enden der metallischen Befestigungszungen an dem Metallkranz festgelötet oder festgeschweißt sind.
  6. Turbomaschine nach Anspruch 1, dadurch gekennzeichnet, daß die ersten Enden der metallischen Befestigungszungen mit dem Metallkranz ein einziges Teil bilden.
  7. Turbomaschine nach Anspruch 1, dadurch gekennzeichnet, daß der Verbundkranz an einem stromabwärtigen Ende (88, 90) der Brennkammer angelötet ist.
  8. Turbomaschine nach Anspruch 1, dadurch gekennzeichnet, daß der Verbundkranz an einem stromabwärtigen Ende (88, 90) der Brennkammer angenäht ist.
  9. Turbomaschine nach Anspruch 1, dadurch gekennzeichnet, daß der Verbundkranz an einem stromabwärtigen Ende (88, 90) der Brennkammer eingesetzt ist.
  10. Turbomaschine nach Anspruch 1, dadurch gekennzeichnet, daß der Verbundkranz einen bestimmten Teil aufweist, der eine Auflageebene für eine Dichtung (80, 82) bildet, welche die Dichtigkeit des Gaskanals zwischen der Brennkammer und dem Leitrad sicherstellt.
  11. Turbomaschine nach Anspruch 10, dadurch gekenntzeichnet, daß der bestimmte Teil ein Endteil des Verbundkranzes ist.
  12. Turbomaschine nach Anspruch 10, dadurch gekennzeichnet, daß die Dichtung vom Typ "Lamellen"-Runddichtung (80, 82) ist.
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)

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

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

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JP2003014234A (ja) 2003-01-15
US6708495B2 (en) 2004-03-23
FR2825783A1 (fr) 2002-12-13
EP1265034A1 (de) 2002-12-11
US20020184892A1 (en) 2002-12-12
JP3907529B2 (ja) 2007-04-18
DE60229465D1 (de) 2008-12-04
FR2825783B1 (fr) 2003-11-07

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