Patent Publication Number: US-6708495-B2

Title: Fastening a CMC combustion chamber in a turbomachine using brazed tabs

Description:
FIELD OF THE INVENTION 
     The present invention relates to the specific field of turbomachines and it relates more particularly to the problem posed by assembling a combustion chamber made of a composite material of the ceramic matrix composite (CMC) type in the metal chamber casings of a turbomachine. 
     PRIOR ART 
     Conventionally, in a turbojet or a turboprop, the high pressure turbine, in particular its inlet nozzle (HPT nozzle), the combustion chamber, and the inner and outer shells (or casings) of said chamber are all made out of the same material, generally a metal. Nevertheless, under certain particular conditions of use implementing particularly high combustion temperatures, a metal chamber turns out to be completely unsuitable from a thermal point of view and it is necessary to make use of a chamber that is based on high temperature composite materials of the CMC type. However, difficulties of implementation and materials costs mean that such materials are generally restricted to being used for the composite chamber itself, with the high pressure turbine inlet nozzle and the inner and outer shells of the chamber then still being made more conventionally out of metal materials. Unfortunately, metals and composites have coefficients of thermal expansion that are very different. This gives rise to particularly awkward problems of connection with the inner and outer shells of the combustion chamber and of interface at the nozzle at the inlet to the high pressure turbine. 
     OBJECT AND BRIEF SUMMARY OF THE INVENTION 
     The present invention mitigates those drawbacks by proposing a mounting for the combustion chamber in the casings with the ability to absorb the displacements induced by the various coefficients of expansion of those parts. An object of the invention is thus to propose a mounting which makes the best use of the existing characteristics of the combustion chamber. 
     These objects are achieved by a turbomachine comprising inner and outer annular shells of metal material containing in a gas flow direction F: a fuel injector assembly, an annular combustion chamber of composite material and having a longitudinal axis, and an annular nozzle of metal material and forming the fixed-blade inlet stage of a high pressure turbine, wherein said composite material combustion chamber is held in position between said inner and outer metal annular shells by a plurality of flexible tongues, first ends of said tongues being interconnected by a metal ring fixed securely to each of said inner and outer metal annular shells by first fixing means, and second ends being fixed by second fixing means to a ring of composite material fixed securely to said composite material combustion chamber, the flexibility of said fixing tongues allowing expansion to take place freely in a radial direction at high temperatures between said composite material combustion chamber and said metal annular shells. 
     With this particular structure for the fixed connection, the various kinds of wear due to contact corrosion in prior art systems can be avoided. The use of a ring made of composite material to provide sealing of the stream also makes it possible to keep the initial structure of the chamber intact. In addition, the presence of flexible metal tongues replacing the traditional flanges gives rise to a saving in mass that is particularly appreciable. In addition to being flexible, these tongues make it easy to accommodate the expansion difference that appears at high temperatures between metal parts and composite parts (by accommodating the displacements due to expansion) while still ensuring that the combustion chamber is properly held and well centered in the annular shell. 
     The first and second fixing means are preferably constituted by a plurality of bolts. 
     In an advantageous embodiment in which each of said metal annular shells is made up of two portions, said metal ring interconnecting said first ends of said metal fixing tongues is mounted between connecting flanges of said two portions. In an alternative embodiment, said metal ring can be fixed directly to said annular shell by fixing means. 
     Depending on the intended embodiment, said first ends of the fixing tongues can either be fixed by brazing to said metal ring, or else they can be formed integrally with said metal ring. 
     In a preferred embodiment, said composite ring is brazed onto a downstream end of the combustion chamber. In an alternative embodiment, the composite ring is sewn onto the downstream end. In another embodiment, the composite ring is implanted on the downstream end. 
     Said composite ring includes a determined portion forming a bearing plane for a sealing gasket (advantageously of the circular “spring blade” gasket type) ensuring that the stream of gas between said combustion chamber and said nozzle is sealed. Said determined portion is preferably an end portion of said composite ring. 
    
    
     BRIEF DESCRIPTION OF THE DRAWINGS 
     The characteristics and advantages of the present invention appear better from the following description made by way of non-limiting indication and with reference to the accompanying drawings, in which: 
     FIG. 1 is a diagrammatic axial half-section of a central zone of a turbomachine in a first embodiment of the invention; 
     FIG. 1A is a fragmentary view of a flexible fixing tongue of a first embodiment of the invention; 
     FIG. 1B is a fragmentary cross sectional view of a portion of FIG. 1 in an alternative crimping connection configuration; 
     FIG. 2 is a view on a larger scale showing a portion of FIG. 1 in an alternative connection configuration; and 
     FIG. 3 is an enlarged view of another portion of FIG. 1 in an alternative connection configuration. 
    
    
     DETAILED DESCRIPTION OF A PREFERRED EMBODIMENT 
     FIG. 1 is an axial half-section view of a central portion of a turbojet or a turboprop (with the term “turbomachine” being used generically in the description below) and comprising: 
     an outer annular shell (or outer casing) made up of two portions  12   a  and  12   b  of metal material, having a longitudinal axis  10 ; 
     an inner annular shell (or inner casing) that is coaxial therewith and likewise comprises two portions  14   a  and  14   b , also made of metal material; and 
     an annular space  16  extending between the two shells  12   a ,  12   b  and  14   a ,  14   b  for receiving compressed oxidizer, generally air, coming from an upstream compressor (not shown) of the turbomachine via an annular diffuser duct  18  defining a general flow F of gas. 
     In the gas flow direction, this space  16  comprises firstly an injection assembly formed by a plurality of injection systems  20  that are regularly distributed around the duct  18 , each comprising a fuel injection nozzle  22  fixed to an upstream portion  12   a  of the outer annular shell  12  (in order to simplify the drawings, the mixer and the deflector associated with each injection nozzle are omitted), followed by a combustion chamber  24  of high temperature composite material, e.g. of the CMC type or of some other type (e.g. carbon), formed by an outer axially-extending side wall  26  and an inner axially-extending side wall  28 , both disposed coaxially about the axis  10 , and a transversely-extending end wall  30  of said combustion chamber and which has margins  32 ,  34  fixed by any suitable means, e.g. metal or refractory bolts with flat head screws, to the upstream ends  36 ,  38  of said side walls  26 ,  28 , this chamber end wall  30  being provided with through orifices  40  to enable fuel to be injected together with a fraction of the oxidizer into the combustion chamber  24 , and finally an annular nozzle  42  of metal material forming an inlet stage of a high pressure turbine (not shown) and conventionally comprising a plurality of fixed blades  44  mounted between an outer circular platform  46  and an inner circular platform  48 . 
     The nozzle is fixed to the downstream portion  14   b  of the inner annular shell of the turbomachine by first removable fixing means preferably constituted by a plurality of bolts  50 , while resting on support means  49  secured to the outer annular shell of the turbomachine. 
     Through orifices  54 ,  56  formed in the outer and inner metal platforms  46  and  48  of the nozzle  42  are also provided to cool the fixed blades  46  of this nozzle at the inlet to the rotor of the high pressure turbine using compressed oxidizer available at the outlet from the diffusion duct  18  and flowing in two flows F 1  and F 2  on either side of the combustion chamber  24 . 
     The combustion chamber  24  has a coefficient of thermal expansion that is very different from that of the other parts forming the turbomachine, since they are made of metal. In accordance with the invention, the combustion chamber  24  is held securely in position between the inner and outer annular shells by a plurality of flexible tongues  58 ,  60  regularly distributed around the combustion chamber. A first fraction of these fixing tongues (see the tongues referenced  58 ) is mounted between the outer annular shell  12   a ,  12   b  and the outer side wall  26  of the combustion chamber, while a second fraction (like the tongues  60 ) is mounted between the inner annular shell  14   a ,  14   b  and the inner side wall  28  of the combustion chamber. 
     Each flexible fixing tongue of metal material can be substantially triangular in shape as shown in FIG. 1A or it can be constituted by a single blade (of optionally constant width), and it is welded or brazed at a first end  62 ;  64  to a metal ring  66   a ,  66   b  fixed securely by first fixing means  52 ;  68  to one or the other of the inner and outer metal annular shells  12 ,  15  (depending on where it is located) and intended to make it easier both to hold these tongues and to seal the annular gap  16 . In a preferred embodiment, these tongues and the metal ring together form a single one-piece metal part. At a second end  70 ;  72 , each tongue is securely fixed via second fixing means  74 ,  76  to a ceramic composite ring  78   a ;  78   b  brazed onto a downstream end  88 ;  90  of the outer and inner side walls  26  and  28  of the ceramic composite material combustion chamber. This brazing can be replaced or even reinforced by stitching. The connection between the chamber walls and the rings can also be made entirely by implantation (connection of the type known by the term “pin&#39; sage”). By way of example, the number of tongues can be a number that is equal to the number of injection nozzles or to a multiple of said number. 
     FIG. 1 shows a first embodiment of the invention in which the second ends of the tongues  70 ,  72  are respectively fixed on the outer and inner ceramic composite rings  78   a  and  78   b  by simple bolting (but crimping could also be envisaged, as shown in fragmentary view in FIG.  1 B). The metal ring  66   a ,  66   b  interconnecting the first ends  62 ,  64  of the tongues is preferably clamped between the existing connection flanges between the upstream and downstream portions of the inner and outer annular shells  14 ,  12  and held securely by the first fixing means  52 ,  68  which are preferably likewise of the bolt type. It should be observed that ceramic composite material washers  74   a ;  76   a  are provided to enable the flat headed screws of the bolts forming the second fixing means  74 ;  76  to be “embedded”. 
     In the variant shown in FIG. 2, the metal ring  66   a  interconnecting the first ends  62  of the fixing tongues  58  of the outer side wall  26  of the combustion chamber by welding (or brazing) is no longer mounted between flanges but is itself welded (or brazed) to a centered keying element  106  secured to the outer annular shell  12 . 
     In another variant shown in FIG. 3, the metal ring  66   b  interconnecting the first ends  64  of the fixing tongues  60  of the inner side wall  28  of the combustion chamber by welding (or brazing) is no longer mounted between flanges but is merely fixed directly to the inner annular shell  14  by fixing means  108 , e.g. of the bolt type. 
     The stream of gas between the combustion chamber  24  and the nozzle  42  is sealed by a circular “spring blade” gasket  80 ,  82  mounted in a groove  84 ,  86  of each of the outer and inner platforms  46  and  48  of the nozzle and which bear directly against a portion of the ceramic composite ring  78   a ;  78   b  forming a bearing plane for said circular sealing gasket. The portion can be an end portion of the ring. The gasket is pressed against said end portion of the composite ring or any other portion by means of a resilient element  92 ,  94  fixed to the nozzle. By means of this disposition, perfect sealing is ensured for the hot stream between the combustion chamber  24  and the nozzle  42 . 
     The gas flows between the combustion chamber and the turbine are sealed firstly by an omega type circular sealing gasket  96  mounted in a circular groove  98  of a flange of the inner annular shell  14  in direct contact with the inner circular platform  48  of the nozzle, and secondly by another circular spring blade gasket  100  mounted in a circular groove  102  of the outer circular platform of the nozzle  46  and having one end in direct contact with a circular projection  104  on the downstream portion  12   b  of the outer annular shell. 
     In all of the above-described configurations, the flexibility of the fixing tongues makes it possible to accommodate the thermal expansion difference that appears at high temperatures between the composite material combustion chamber and the metal annular shells, while continuing to hold and position the combustion chamber.