Abstract:
A turbomachine assembly in which a foil is configured to cover mainly one of bulbs of a disc and to be held, radially with respect to the disc, by the bulb of the disc and a pocket for a blade that can collaborate therewith, when these two are effectively collaborating, and the bulb of the disc includes at least one longitudinal cavity configured to form, with the foil, when the foil is covering the bulb of the disc, a secondary passage through which a secondary cooling air flow can pass.

Description:
BACKGROUND OF THE INVENTION 
     Field of the Invention 
     The present invention relates to the field of turbojet engines. 
     Description of the Related Art 
     A multi-flow (for example, dual flow) turbojet engine, for propelling transport aeroplanes, generally comprises an upstream fan delivering an annular air flow, this flow comprising a primary, central, annular portion that supplies the engine driving the fan, together with a secondary, outer, annular portion, intended to be exhausted into the atmosphere while providing a considerable fraction of the thrust. 
     In order to compress this annular air flow, the turbojet engine is fitted with a set of wheels provided with blades of large dimensions, which are fixed to discs by means of bulbs and recesses with which each of the discs and blades is provided, the bulbs of the disc having a complementary shape to that of the recesses of the blade and being capable of cooperating therewith, the recesses of the disc having a complementary shape to that of the bulbs of the blade capable of cooperating therewith. 
     Because the use of the turbojet engine is limited by temperature, the end of the disc, which constitutes the part of the disc nearest to the flow path, must be cooled. 
     Currently, this cooling is produced by a cooling air flow passing into the bottom of the recesses of the disc. To this end, each bulb of the blade is arranged relative to one of the recesses of the disc so as to contrive between them, when said bulb cooperates with said recess, a main channel through which the cooling air flow can pass, a foil also being positioned between said disc and said blade. This air flow, having crossed the main channel, and thus having cooled its boundary area, can then be exhausted into the annular air flow. 
     In this configuration, the cooling air flow mainly allows the recess and the bottom of the bulb of the disc to be cooled, to the detriment of the apex of the bulb. However, the apex of the bulb constitutes the most thermally stressed zone of the disc and therefore requires the most cooling. Consequently, cooling of the disc is not currently effective. 
     BRIEF SUMMARY OF THE INVENTION 
     The subject-matter of the present invention is to improve the effectiveness of cooling of the disc of a turbojet engine as defined above. 
     To this end, according to the invention, the turbine engine assembly, comprising a disc and a blade each having a set of bulbs and recesses, the bulbs of the disc being capable of cooperating with the recesses of the blade, the recesses of the disc being capable of cooperating with the bulbs of the blade, at least one of said bulbs of the blade being arranged relative to one of said recesses of the disc so as to contrive between them, when the bulb cooperates with said recess, a main channel through which a cooling air flow can pass, the assembly further comprising a foil capable of being positioned at least partly between said disc and said blade, is remarkable in that:
         said foil is capable of covering mainly one of said bulbs of the disc and of being held, radially relative to said disc, by said bulb of the disc and the recess of the blade capable of cooperating therewith, when they are actually cooperating; and   said bulb of the disc has at least one longitudinal cavity capable of forming, together with said foil, when the foil is covering said bulb of the disc, a secondary channel through which said cooling air flow can pass.       

     In this way, thanks to the present invention, the cooling air flow is routed directly to the apex of each bulb of the disc, which allows this particularly sensitive zone of the disc to be cooled more effectively, and therefore means that less of the air flow needed for propelling the aeroplane has to be taken off, hence improving the overall performance of the turbojet engine. 
     Where the bulb of the disc intended to be covered by the foil is positioned between two recesses of said disc, the foil advantageously has, on one side, a long end intended to cover at least partly one of said recesses and, on the other side, a short end. This asymmetry of the foil has the advantage of making it easier to fit. 
     Another advantage of this asymmetry is that the long end of the foil can be extended by means for axially locking said foil with respect to the recess covered by said long end. 
     In an advantageous embodiment, the bulb of the disc has at least two longitudinal cavities separated by an apex of said bulb, these being capable, when the foil is covering said bulb, of forming two secondary channels through which the cooling air flow can pass. In this way, retaining an apex of the bulb of the disc, which separates two secondary channels, makes it possible to ensure that the blade does not tilt over the apex of the bulb of the disc, which has the advantage of preventing any damage to the foil. 
     In order to route the cooling air flow into the secondary channels, delimited by the foil and the bulbs of the disc, the turbine engine assembly according to the invention can further comprise a sealing lock ring joining the disc and the blade upstream of them in terms of the direction in which the main and secondary cooling air flows pass, said sealing lock ring having a set of radial cavities via which said secondary cooling air flow can be routed into the secondary channel. 
     To channel the cooling air flow upstream and route it in front of the secondary channels, the sealing lock ring also has a set of radial grooves through which the blades pass when said sealing lock ring joins the disc and the blade. This creates a passage for the secondary cooling air flow. 
     Advantageously, the radial cavities of the sealing lock ring are arranged so that said radial cavities are positioned facing the bulbs of the disc when the sealing lock ring joins the disc and the blade. Moreover, to ensure that each radial cavity is correctly positioned facing each bulb of the disc, the turbine engine assembly according to the invention further comprises a movable ring having a set of radial projections capable of being housed in the radial grooves of the sealing lock ring, to prevent any rotation of said sealing lock ring. 
     In a first embodiment, the movable ring is arranged so as to cause the axial stopping of the blades upstream, in terms of the direction in which the main and secondary cooling air flows pass. 
     In a second embodiment, the sealing lock ring and the movable ring have respectively the same number of radial grooves and radial projections as the number of blade bulbs intended to cooperate with the disc, and the movable ring is arranged so as to cause the axial stopping of the blades upstream, in terms of the direction in which the main and secondary cooling air flows pass. 
     The invention also relates to a foil for an assembly according to one of the embodiments described above, said foil being capable of being positioned at least partly between the disc and the blade, this foil being remarkable:
         in that it is capable of covering mainly one of the bulbs of the disc and of being held, radially relative to the disc, by said bulb of the disc and the recess of the blade capable of cooperating therewith, when they are actually cooperating; and   in that, when it covers said bulb of the disc, it forms, together with at least one longitudinal cavity of said bulb of the disc, a secondary channel through which the secondary cooling air flow can pass.       

     The invention also relates to a turbine engine comprising at least one assembly according to one of the embodiments described above. 
    
    
     
       BRIEF DESCRIPTION OF THE SEVERAL VIEWS OF THE DRAWINGS 
       The figures in the appended drawing will make it easier to understand how the invention can be implemented. In these figures, identical reference signs denote similar technical elements. 
         FIG. 1  is a diagram, in a cross-sectional plane, of a turbine engine according to the invention. 
         FIG. 2  is a diagram, in a cross-sectional plane, of a turbine engine assembly according to the invention, in which the disc and one blade are shown in part, in the area where they join. 
         FIG. 3  is a perspective view of the apexes of the disc and one blade, separated by a foil, when they are cooperating. 
         FIG. 4  is a perspective view of the apex of the disc in  FIG. 3 . 
         FIG. 5  is a perspective view of the foil in  FIG. 3 . 
         FIG. 6  is a perspective view of the apex of the disc in  FIG. 3 , when it is covered by the foil. 
         FIG. 7  is a perspective view of the turbine engine assembly according to the invention, also comprising a sealing lock ring and a movable ring. 
         FIG. 8  is a perspective view of the turbine engine assembly in  FIG. 7  without the movable ring. 
         FIG. 9  is a perspective view of one portion of the sealing lock ring, according to a first embodiment thereof. 
         FIG. 10  is a perspective view of the lock ring of the portion of the sealing lock ring in  FIG. 9 , when it is cooperating with the movable ring. 
         FIG. 11  is a perspective view of the turbine engine assembly fitted with the sealing lock ring and the movable ring in  FIGS. 9 and 10 . 
         FIG. 12  is a perspective view of one portion of the sealing lock ring, according to a second embodiment thereof. 
         FIG. 13  is a perspective view of the lock ring of the portion of the sealing lock ring in  FIG. 12 , when it is cooperating with the movable ring. 
         FIG. 14  is a perspective view of the turbine engine assembly fitted with the sealing lock ring and the movable ring in  FIGS. 12 and 13 . 
     
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
     The turbine engine  1  in  FIG. 1  is of the dual-flow, two-spool type, having rotational symmetry about an axis X-X′. In a known manner, this turbine engine  1  comprises, within a nacelle  2  serving as an envelope for its various members, an air inlet through which an incoming air flow F can penetrate and then pass through an inlet fan  4 , positioned around an air inlet cone  3  that allows the total flow F to be guided aerodynamically and distributed about the axis X-X′. This air flow F is then separated into two flows, primary FP and secondary FS respectively, via an intermediate casing  5 , the end of which forms a separating slat. 
     In the rest of the description, the terms “upstream” and “downstream” relate to axial positions along the longitudinal axis X-X′ in the direction of flow of the air flow within the turbojet engine  1 . 
     The secondary flow FS passes through a straightening stage and is then exhausted downstream of the turbine engine. The primary flow FP passes successively through a low-pressure compression stage  5 A, a high-pressure compression stage  5 B, a combustion chamber  6 , a high-pressure turbine stage  7 A and a low-pressure turbine stage  7 B, finally being exhausted out of the turbine engine through a pipe (not given a reference sign). 
     The nacelle  2  of this turbojet engine is annular and positioned at least approximately coaxially about the longitudinal axis X-X′. This allows the gas flows generated by the turbine engine to be channelled while defining inner and outer aerodynamic flow lines for the gas flows. 
     As shown in  FIG. 2 , one of the discs  10  of the turbine engine  1  in  FIG. 1 , capable of rotating about the axis X-X′ of said turbine engine, has a set of connections by means of which a plurality of blades, including the blade  20 , are fixed thereto. To cool the head of the disc  10 , which constitutes the area closest to the main flow path, a portion  50  of the primary air flow FP is taken off, such that at least part thereof, denoted by the reference sign  51  in  FIG. 2  and referred to below as main cooling air flow, passes through the connections between the disc  10  and the blades  20 . 
     The connections between the disc  10  and the blades  20  are produced by a set of bulbs and recesses, with which each of them is provided. The disc  10  thus has, at its end, a set of bulbs  11  and recesses  12 , while the blade  20  has, at its end intended to cooperate with the disc  10 , a set of bulbs  21  and recesses  22 . The shape and dimensions of the bulbs  11  of the disc  10  are determined in such a way that they are capable of cooperating with the recesses  22  of the blade  20 . In the same way, the bulbs  21  of the blade  20  are determined in such a way that they are capable of cooperating with the recesses  12  of the disc  10 . Thus the blades  20  can easily be connected to the disc  10 . 
     A foil  30 , which will be described in greater detail below with reference to  FIG. 5 , is also positioned between the disc  10  and a blade  20 . 
     The passage of the main cooling air flow  51  close to the end of the disc  10 , for cooling said disc, is ensured when the disc  10  and the blade  20  are cooperating and when the foil  30  is positioned between said disc and said blade, by contriving a main channel  40  delimited on the one hand by the foil  30  and on the other hand by the bulb  21  of the blade  20 . In this way the recess  12  of the disc  10  can be cooled by the passage of the air flow  51 . 
     In accordance with the present invention, the bulb  11  of the disc  10 , shown in isolation in  FIG. 4  and having a lower portion  11 . 1  with a narrower cross-section and an upper portion  11 . 2  with a wider cross-section, has two longitudinal cavities  11 . 4  and  11 . 5 , separated by an apex  11 . 3 , in the area of the upper portion  11 . 2 . 
     Furthermore, the foil  30 , shown in isolation in  FIG. 5 , has a shape similar to that of the bulbs  11  and recesses  12  of the disc  10 . In particular, the foil  30  comprises two upper portions  35  and  36 , capable of covering the longitudinal cavities  11 . 4  and  11 . 5  of the disc  10 , together with two side portions  33  and  34  intended to cover the side walls of the bulb  11 . Given the (narrow and wide) cross-sections of the bulb  11 , the foil  30  is flexible, so that it can be inserted into the bulb  11 . Moreover, the foil  30  has a first, long end  31 , in the extension of the side portion  33 , and a second end  32 , shorter than the end  31 , in the extension of the side portion  34 . 
     The foil  30 , which is thin, is thus capable of covering mainly the bulb  11  and of being held, radially relative to the disc  10 , by said bulb  11  and the recess  22 , when they are actually cooperating. 
     When the foil  30  is positioned over the bulb  11  ( FIG. 6 ), the upper portions  35  and  36  respectively delimit, together with the longitudinal cavities  11 . 4  and  11 . 5 , two secondary channels  41  and  42  through which two secondary cooling air flows  52  and  53  can pass, for cooling the zone of the disc  10  that is most subject to thermal stresses, namely the apex of the bulb  11 . 
     It will be noted that, with the aim of axially locking the foil  30  with respect to the recess  12  covered by said long end  31 , the long end  31  of the foil  30  is extended by axial locking means  33 A,  33 B, positioned on either side of the end  31  so as to bear against the walls of the disc  10  when the foil  30  is inserted into the bulb  11 . 
     Referring now to  FIGS. 7 and 8 , a sealing lock ring  60  is positioned against the disc  10  and the blade  20  so that it can join said disc and said blade on the upstream side in terms of the direction in which the main cooling air flow  51  passes. This sealing lock ring  60 , shown in greater detail in  FIG. 9 , has a substantially crescent-shaped set of radial cavities  62 , positioned facing the bulbs  11  of the disc  10  when the sealing lock ring  60  joins the disc  10  and the blade  20 . The radial cavities  62  thus allow the secondary cooling air flows  52  and  53 , coming from the air flow  50 , to pass as far as the secondary channels  41  and  42 . 
     In this way the cooling air flow  50  is divided into a main air flow  51 , which passes through the main channel  40 , and two secondary air flows  52  and  53 , which pass through the secondary channels  41  and  42  respectively ( FIG. 8 ). 
     It will be noted that the invention can be implemented with a different number of secondary channels, in so far as the bulb  11  has at least one of these. It is, however, preferable to have at least two secondary channels, so that the bulb  11  has at least two longitudinal cavities (such as the cavities  11 . 4  and  11 . 5 ) each separated by an apex (such as the apex  11 . 3 ), which makes it possible to preserve the anti-tilt function of the blade  20  over the apex of the bulb  11  and thus not to damage the foil  30 . 
     The sealing lock ring  60  also has a set of radial grooves  61  through which the blades pass, at the time they are fitted, when the sealing lock ring  60  joins the disc  10  and the blade  20 . These radial grooves  61  also allow the sealing lock ring  60  to be fixed to a movable ring  70 , which has for the purpose a set of radial projections  71  with a complementary shape to the radial grooves  61  of the lock ring  60 , and thus to prevent any rotation of said sealing lock ring. 
     The movable ring  70  is arranged, relative to the blades  20 , so as to cause the axial stopping thereof downstream, with respect to the direction in which the main  51  and secondary  52 ,  53  cooling air flows pass. 
     In a variant of the invention, shown in  FIGS. 12 to 14 , the sealing lock ring  60  and the movable ring  70  are replaced by the lock ring  80  and the ring  90  respectively. The sealing lock ring  80  has a set of radial cavities  82  similar to the radial cavities  62 , together with equally spaced radial grooves  81 . As for the movable ring  90 , this has a set of equally spaced radial projections  91  with a complementary shape to the radial grooves  81 , so as to prevent the lock ring  80  from rotating with respect to the ring  90  ( FIG. 13 ). In this embodiment, the number of radial projections  91  of the ring  90  is equal to the number of hooks  23  of the blade  20 . 
     Moreover, the movable ring  90  is arranged, relative to the blades  20 , so as to cause the axial stopping thereof upstream in terms of the direction in which the main  51  and secondary  52 ,  53  cooling air flows pass.