EP1265034B1 - Mounting of a turbine ceramic matrix composite combustion chamber with brazed mounting lugs - Google Patents

Mounting of a turbine ceramic matrix composite combustion chamber with brazed mounting lugs Download PDF

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Publication number
EP1265034B1
EP1265034B1 EP20020291363 EP02291363A EP1265034B1 EP 1265034 B1 EP1265034 B1 EP 1265034B1 EP 20020291363 EP20020291363 EP 20020291363 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
characterized
ring
turbomachine according
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.)
Active
Application number
EP20020291363
Other languages
German (de)
French (fr)
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EP1265034A1 (en
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
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Safran Aircraft Engines 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
Priority to FR0107363 priority Critical
Priority to FR0107363A priority patent/FR2825783B1/en
Application filed by Safran Aircraft Engines SAS filed Critical Safran Aircraft Engines SAS
Publication of EP1265034A1 publication Critical patent/EP1265034A1/en
Application granted granted Critical
Publication of EP1265034B1 publication Critical patent/EP1265034B1/en
Application status is Active legal-status Critical
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    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23RGENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R3/00Continuous combustion chambers using liquid or gaseous fuel
    • F23R3/42Continuous combustion chambers using liquid or gaseous fuel characterised by the arrangement or form of the flame tubes or combustion chambers
    • F23R3/60Support structures; Attaching or mounting means
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23RGENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R3/00Continuous combustion chambers using liquid or gaseous fuel
    • F23R3/007Continuous combustion chambers using liquid or gaseous fuel constructed mainly of ceramic components
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05BINDEXING SCHEME RELATING TO MACHINES OR ENGINES OTHER THAN NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, TO WIND MOTORS, TO NON-POSITIVE DISPLACEMENT PUMPS, AND TO GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY
    • F05B2230/00Manufacture
    • F05B2230/60Assembly methods
    • F05B2230/604Assembly methods using positioning or alignment devices for aligning or centering, e.g. pins
    • F05B2230/606Assembly methods using positioning or alignment devices for aligning or centering, e.g. pins using maintaining alignment while permitting differential dilatation

Description

    Field of the invention
  • 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. Such a turbomachine is known from GB 1,570,875 .
  • Prior art
  • Conventionally, in a turbojet or a turboprop, 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. However, under certain particular conditions of use involving significantly high combustion temperatures, 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. However, 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. However, 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.
  • Object and definition of the invention
  • 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.
  • These objects are achieved by 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.
  • With this particular structure of fixed connection, the different wear due to contact corrosion of systems of the prior art can be avoided. The use of a crown made of composite material to seal the vein also allows to keep intact the initial structure of the chamber. In addition, the presence of flexible metal tongues instead of traditional flanges allows a particularly appreciable mass gain. These tongues, because of their flexibility, can easily withstand the expansion gap occurring at high temperatures between metal parts and composites (taking again the displacement due to expansion) while ensuring a perfect fit and a good centering of the chamber. combustion in the annular envelope.
  • The first and second fastening means are preferably constituted by a plurality of bolts.
  • According to an advantageous embodiment in which 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. In an alternative embodiment, said metal ring may be fixed directly to said annular envelope by fixing means.
  • According to the embodiment envisaged, said first ends of the fixing tongues may either be soldered to said metal ring or form a single piece with this metal ring.
  • According to a preferred embodiment, said composite ring is brazed on a downstream end of the combustion chamber. In an alternative embodiment, this composite crown is sewn on this downstream end. In another embodiment, 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. Preferably, said determined portion is an end portion of said composite ring.
  • Brief description of the drawings
  • The features and advantages of the present invention will emerge more clearly from the following description, given by way of non-limiting indication, with reference to the appended drawings in which:
    • the figure 1 is a schematic axial half-sectional view of a central zone of a turbomachine in a first embodiment of the invention,
    • the figure 2 is an enlarged view of some of the figure 1 in an alternative link configuration, and
    • the figure 3 is an enlarged view of another part of the figure 1 in an alternative link configuration.
    Detailed description of a preferred embodiment
  • The figure 1 axial half-section shows a central portion of a turbojet engine or a turboprop engine (called a turbomachine in the following description) comprising:
    • . an outer annular casing (or outer casing) in two parts 12a, 12b made of metallic material, of longitudinal axis 10,
    • . an inner annular casing (or inner casing) coaxial in two parts 14a, 14b also made of metallic material,
    • . an annular space 16 between the two envelopes 12a, 12b and 14a, 14b receiving the compressed oxidant, generally air, coming upstream of a compressor (not shown) of the turbomachine, through an annular duct of diffusion 18 defining a general flow F of gas flow,
      this space 16 comprising, in the gas flow direction, firstly an injection assembly formed of a plurality of injection systems 20 regularly distributed around the conduit 18 and each having a fuel injection nozzle 22 fixed on an upstream portion 12a of the outer annular envelope (for the sake of simplification of the drawings the mixer and the deflector associated with each injection nozzle have not been shown), then a combustion chamber 24 of composite material high temperature, for example CMC or other type (carbon for example), formed of an outer axial wall 26 and an inner axial wall 28, both coaxial axis 10, and a transverse wall 30 which constitutes the bottom of this combustion chamber and which comprises flaps 32, 34 fixed by any suitable means, for example by metal bolts or refractory screws with conical head, on upstream ends 36, 38 of the axial walls 26, 2 8, this bottom chamber 30 being provided with through holes 40 to allow the injection of fuel and a portion of the oxidant into the combustion chamber 24, and finally an annular distributor 42 of metal material forming a stage of inlet of a high pressure turbine (not shown) and conventionally comprising a plurality of stationary vanes 44 mounted between an outer circular platform 46 and an inner circular platform 48.
  • 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.
  • According to the invention, 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. In a preferred embodiment, these tabs and the metal ring together form a single piece of metal in one piece. At a second end 70, 72, 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. Note the presence of washers ceramic composite material 74a; 76a to allow to "drown" the conical heads of the screws of the bolts forming the second attachment means 74; 76.
  • In a variant illustrated in figure 2 , 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.
  • In another variant illustrated in figure 3 , 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. By this arrangement, it is ensured a perfect seal of the hot vein between the combustion chamber 24 and the distributor 42.
  • As regards the sealing of the gas flow flows between the combustion chamber and the turbine, it is formed firstly by 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.
  • In all the aforementioned configurations, 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.

Claims (12)

  1. A turbomachine comprising inner and outer annular shells (12, 14) of metal material containing in a gas flow direction F: a fuel injector assembly (20, 22), an annular combustion chamber (24) of composite maternal and having a longitudinal axis (10), and an annular nozzle (42) of metal material and forming the fixed-blade inset stage (44) of a high pressure turbine, said composite material combustion chamber being held in position between said inner and outer metal annular shells by a plurality of flexible metal tongues (58, 60), first ends (62, 64) of said tongues being interconnnected by a metal ring (66a, 66b) fixed securely to one of said inner and outer metal annular shells (12, 14) by first fixing means (52, 68, 108), the turbomachine being characterized in that second ends (70, 72) of said tongues are fixed by second fixing means (74, 76) on a ring (78a, 78b) of composite maternal fixed securely to said composite material combustion chamber (26, 28), the flexibility of said metal fixing tongues allowing expansion to take place freely in a radial direction at high temperature between said composite material combustion chamber and said annular shells.
  2. A turbomachine according to claim 1, characterized in that said first and second fixing means are constituted by a plurality of balts.
  3. A turbomachine according to claim 1, characterized in that each of said metal annular shells is made up of two portions (12a, 12b, 14a, 14b), and said metal ring interconnnecting said first ends of said metal fixing tongues is mounted between the connection flanges of said two portions.
  4. A turbomachine according to claim 1, characterized in that said metal ring interconnecting said first ends of said metal fixing tongues is fixed directly to said annular shell by fixing means (108).
  5. A turbomachine according to claim 1, characterized in that said first ends of the metal fixing tongues are fixed by brazing or welding to said metal ring.
  6. A turbomachine according two claim 1, characterized in that said first ends of the metal fixing tongues are integrally formed with said metal ring.
  7. A turbomachine according to claim 1, characterized in that said composite ring is brazed onto a downstream end (88, 90) of the combustion chamber.
  8. A turbomachine according to claim 1, characterized in that said composite ring is sewn onto a downstream end (88, 90) of the combustion chamber.
  9. A turbomachine according to claim 1, characterized in that said composite ring is implanted on a downstream end (88, 90) of the combustion chamber.
  10. A turbomachine according to claim 1, characterized in that said composite ring includes a determined portion forming a bearing plane for a sealing gasket (80, 82) ensuring that the stream of gas between said combustion chamber and said nozzle is sealed.
  11. A turbomachine according to claim 10, characterized in that said determined portion is an end portion of said composite ring.
  12. A turbomachine according to claim 10, characterized in that said sealing element is of the circular spring blade gasket type (80, 82).
EP20020291363 2001-06-06 2002-06-04 Mounting of a turbine ceramic matrix composite combustion chamber with brazed mounting lugs Active EP1265034B1 (en)

Priority Applications (2)

Application Number Priority Date Filing Date Title
FR0107363 2001-06-06
FR0107363A FR2825783B1 (en) 2001-06-06 2001-06-06 Hanging of cmc combustion chamber of turbomachine by brazed legs

Publications (2)

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

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EP20020291363 Active EP1265034B1 (en) 2001-06-06 2002-06-04 Mounting of a turbine ceramic matrix composite combustion chamber with brazed mounting lugs

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

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Also Published As

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

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