Rotor element for a turbine engine with composite blades linked to a metal disk

Abstract

The invention relates to a rotor element (48) for a turbine engine, comprising a metal disk (62) and a plurality of blades (49) mounted on the disk (62). At least one of said blades (49) consists of a hybrid blade (79) made of composite material and comprising a metal root (80) linked to the disk (62).

Description

[0001]DESCRIPTION TITLE: ROTOR ELEMENT FOR TURBOMACHINE WITH COMPOSITE BLADES LINKED TO A METAL DISC FIELD OF THE INVENTION The present invention relates to the field of turbomachines and more particularly to that of rotors for turbomachines of the type comprising blades made of composite material. TECHNOLOGICAL BACKGROUND Conventionally, turbomachine rotors, in particular fan and compressor rotors, are formed of blades mounted around a hub, commonly called a "disc", generally metallic. Most often, the blades are also metallic and are fixed to the disc by means of a broached attachment, forming a root of the blade, housed in a cell formed on the periphery of the disc. With the aim of lightening turbomachine rotors, so as to improve consumption and contribute to healthy shaft dynamics, various solutions have been proposed. For compressor rotors, for example,It has been proposed to use single-piece bladed discs (better known by the acronym "DAM" or the English term "blisk") in which the blades are integral with the disc. This solution makes it possible to lighten the rotor by reducing the number of parts and by eliminating the sealing systems required when using broached fasteners. For this purpose, the blades are typically machined from the same block of material as the disc, or welded onto the disc. For fan rotors, it has been proposed to use blades composed at least in part of a composite material structure comprising a fibrous reinforcement densified by a polymer matrix, the blades always being fixed to the disc by means of a broached fastener. This solution makes it possible to lighten the rotor by reducing the mass of the blades, the blades with a composite material structure being, with equivalent propulsive characteristics,lighter than metal blades. However, these solutions are not entirely satisfactory. Indeed, single-piece bladed discs remain relatively dense compared to solutions using composite materials, which have a complex architecture due to the necessary use of sealing systems. DISCLOSURE OF THE INVENTION One objective of the invention is to further lighten turbomachine rotors. Other objectives are to simplify their architecture and to ensure good mechanical strength. To this end, the invention relates, according to a first aspect, to a rotor element for a turbomachine comprising a metal disc and a plurality of blades mounted on the disc, in which at least one of said blades is constituted by a hybrid blade composed of composite material and comprising a metal heel bonded to the disc. According to a particular embodiment of the invention,the rotor element also has one or more of the following characteristics, taken in isolation or in any technically possible combination(s): - each of the blades is constituted by a hybrid blade; - at least one other of the blades is constituted by a metal blade; - at least one other of the blades is constituted by a composite blade; - the or each other blade is constituted by a metal blade; - the or each other blade is constituted by a composite blade; - the or each hybrid blade is composed mainly of composite material; - the heel is bonded to the disc by welding or brazing; - the disc has an edge and, for the or each hybrid blade, a bonding surface, arranged on said edge, to which the heel of said hybrid blade is bonded; - the bonding surface is substantially flat; - the bonding surface is substantially perpendicular to a radial direction of the disc; - the disc comprises,for the or each hybrid blade, a base projecting radially outwards from the edge, said base having, opposite the edge, a distal end constituting the connecting surface of said hybrid blade; - the base has a peripheral surface bordering the distal end and the hybrid blade is positioned relative to the base so that its intrados, its extrados, its leading edge and its trailing edge are each flush with the peripheral surface of the base; - the edge has a shape of revolution; - the edge partly delimits an air stream of the turbomachine; - the rotor element constitutes a rotor stage of a compressor or fan; - the heel has an aerodynamic profile; and - the heel is made of metal and the hybrid blade comprises a distal portion, remote from the disk, composed at least in part of a structure made of composite material,said distal portion being in the extension of the heel to the intrados and to the extrados. The invention also relates, according to a second aspect, to a turbomachine comprising a rotor element according to the first aspect. The invention also relates, according to a third aspect, to an aircraft comprising at least one turbomachine according to the second aspect. The invention finally relates, according to a fourth aspect, to a method of manufacturing a rotor element according to the first aspect, comprising the following steps: - providing a metal disk, - providing at least one hybrid blade composed of composite material and comprising a metal heel, and - connecting the heel of the or each hybrid blade to the disk. According to particular embodiments of the invention, the manufacturing method also has one or more of the following characteristics,taken in isolation or in any technically possible combination(s): - the heel of each hybrid blade is connected to the disc by welding or brazing; - the welding of the heel of each hybrid blade to the disc is carried out by friction, for example by linear or orbital friction; - the connection of the heel of each hybrid blade to the disc produces at least one weld bead, the manufacturing method comprising an additional step of machining the weld bead; and - the method comprises an additional step of static balancing of the rotor element, for example by machining a circumferential bead formed in the disc or adding weights fixed to the disc. BRIEF DESCRIPTION OF THE FIGURES Other characteristics and advantages of the invention will appear on reading the description which follows, given solely by way of example and with reference to the appended drawings,in which: - Figure 1 is a top view of an aircraft according to an exemplary embodiment of the invention, - Figure 2 is a longitudinal sectional view of an upper half of a turbomachine of the aircraft of Figure 1, - Figure 3 is a perspective view, three-quarter front, of a rotor stage of a compressor of the turbomachine of Figure 2, - Figure 4 is a perspective view, three-quarter front, of a blade of the rotor stage of Figure 3, - Figure 5 is a front view of a disc of the rotor stage of Figure 3, - Figure 6 is a view of a detail marked VI in Figure 5, - Figure 7 is a sectional view of the rotor stage of Figure 3 along a sectional plane marked VII-VII in Figure 3, - Figure 8 is a view of a detail marked VIII in Figure 7, - Figure 9 is a diagram illustrating a method of manufacturing a rotor comprising the rotor stage of Figure 3,- Figure 10 is a side view of the rotor stage of Figure 3 during manufacture, at a first stage of the process of Figure 9, - Figure 11 is a view of a detail marked XI of Figure 10, - Figure 12 is a side view of the rotor stage of Figure 3 during manufacture, at a second stage of the process of Figure 9, and - Figure 13 is a view of a detail marked XIII of Figure 12, - Figure 14 is a side view of the rotor stage of Figure 3 during manufacture, at a third stage of the process of Figure 9, - Figure 15 is a view of a detail marked XV of Figure 14, - Figure 16 is a sectional view of a portion of the blade of Figure 4 along a plane oriented parallel to an axis of elongation of said blade and including the chord of the latter, according to a first embodiment of the invention,- Figure 17 is a sectional view of a portion of the blade of Figure 4 along a plane orthogonal to the chord of said blade, according to the first embodiment of the invention, - Figures 18 and 19 are views similar respectively to those of Figures 16 and 17, according to a second embodiment of the invention, and - Figures 20 and 21 are views similar respectively to those of Figures 16 and 17, according to a third embodiment of the invention. DETAILED DESCRIPTION OF AN EXAMPLE EMBODIMENT The aircraft 10 shown in Figure 1 comprises turbomachines 12 to propel it. In the example shown, the aircraft 10 is an airplane. This comprises, in a conventional manner, a fuselage 14, a tailplane 16 and two wings 18. The turbomachines 12 are here two in number and are each housed under a respective wing 18. In a variant (not shown), the turbomachines 12 are arranged along the fuselage 14,for example near the empennage 16. As a further variant (also not shown), the aircraft 10 comprises a single turbomachine 12 or at least three turbomachines 12. One of the turbomachines 12 is shown in Figure 2. As visible in this Figure, the turbomachine 12 is elongated along a longitudinal axis X. It typically has angular symmetry around said longitudinal axis X, that is to say that there is at least one angle for which the turbomachine is invariant by rotation around the longitudinal axis X. Here and in the following, the terms “interior” and “exterior”, “internal” and “external”, as well as their variations, are understood with reference to the axis X,an element described as “inner” or “internal” being oriented towards the X axis while an “outer” or “external” element is oriented opposite the X axis. The term “radial” and its variations are understood in reference to a direction orthogonal to the X axis. The turbomachine 12 comprises, in a conventional manner, a nacelle 20, an internal vein 22 for circulating an air flow through the nacelle 20, a combustion chamber 24 housed in the vein 22, a motor body 26 and a gas exhaust nozzle 28. In the following, the terms “upstream” and “downstream” are understood in reference to a direction of flow of an air flow through the vein 22. The motor body 26 comprises a compressor 30,a turbine 32 and a transmission shaft 34 coupling the turbine 32 to the compressor 30 for driving the compressor 30 by the turbine 32. The compressor 30 is arranged upstream of the combustion chamber 24 and supplies the combustion chamber 24 with compressed air. It comprises a stator 35 and a rotor 36. The stator 35 is secured to the nacelle 20. The rotor 36 is capable of being driven in rotation relative to the stator 35 by means of the transmission shaft 34; for this purpose, it is typically fixedly mounted on the transmission shaft 34. The stator 35 comprises at least one, here a plurality, of stator stage(s) 37 each formed of blades (not referenced) arranged substantially in the same radial plane. Similarly, the rotor 36 comprises at least one, here a plurality,of rotor stage(s) 38 each formed of blades (not referenced) arranged substantially in the same radial plane. The stator stage(s) 37 and rotor stage(s) 38 are the same number and alternate with each other. The turbine 32 is arranged downstream of the combustion chamber 24 and receives the exhaust gases leaving the combustion chamber 24. The transmission shaft 34 has the longitudinal axis X as its axis of rotation. The transmission shaft 34 is guided in rotation relative to the nacelle 20 by means of bearings (not shown). In the example shown, the turbomachine 12 is a multi-spool turbomachine, in particular a double-spool turbomachine, comprising a low-pressure spool 40 in addition to the engine spool 26. The engine spool 26 then constitutes a high-pressure spool, the compressor 30 being a high-pressure compressor,the turbine 32 being a high pressure turbine and the transmission shaft 34 being a high pressure shaft. The low pressure body 40 comprises a low pressure compressor 41, a low pressure turbine 42 and a low pressure shaft 43 coupling the low pressure turbine 42 to the low pressure compressor 41 for driving the low pressure compressor 41 by the low pressure turbine 42. The low pressure compressor 41 is arranged upstream of the high pressure compressor 30 and supplies the latter with compressed air. It comprises a stator 45 and a rotor 46. The stator 45 is integral with the nacelle 20. The rotor 46 is capable of being driven in rotation relative to the stator 45 by means of the low pressure shaft 43; for this purpose, it is typically mounted fixedly on the low pressure shaft 43. The stator 45 comprises at least one, here a plurality, of stator stage(s) 47 each formed of blades (not referenced) arranged substantially in the same radial plane. Similarly,the rotor 46 comprises at least one, here a plurality, of rotor stage(s) 48 each formed of blades 49 (Figure 3) arranged substantially in the same radial plane. The stator stage(s) 47 and rotor stage(s) 48 are the same number and alternate with each other. The low pressure turbine 42 is arranged downstream of the high pressure turbine 32 and receives the exhaust gases leaving the latter. The low pressure shaft 43 is guided in rotation relative to the nacelle 20 by means of bearings (not shown). The low pressure shaft 43 is coaxial with the high pressure shaft 34. It therefore also has the longitudinal axis X as its axis of rotation. In particular, the low pressure shaft 43 extends inside the high pressure shaft 34. Here, the turbomachine 12 also comprises a fan 50 for driving the air flow in an external circulation vein 52 surrounding the nacelle 20. A primary air flow A (hot) is thus distinguished,consisting of the portion of the air flow entrained in the internal circulation vein 22, and a secondary air flow B (cold), consisting of the portion of the air flow entrained in the external circulation vein 52. The fan 50 comprises a fan rotor 54. This fan rotor 54 is rotatably mounted relative to the nacelle 20 around the longitudinal axis X. It comprises a disk 55 provided with fan blades 56 extending substantially radially outwards from the disk 55. These blades 56, when they are rotated, drive the air flow in the external circulation vein 52. The fan rotor 54 is rotated by the low-pressure turbine 42, via the low-pressure shaft 43. In the example shown, this drive is carried out via a reduction gear 57 allowing the fan rotor 54 to rotate at a speed lower than that of the low pressure shaft 43. As a variant (not shown),this drive is direct, that is to say that the fan rotor 54 is integral in rotation with the low pressure shaft 43. The fan rotor 54 is in particular arranged upstream of the internal circulation vein 22 and also drives the air flow in the latter. In the example shown, the fan 50 also comprises a fan stator 58, also called a rectifier, comprising fixed blades 59 arranged at the periphery of the nacelle 20, in the external circulation vein 52, along a plane orthogonal to the longitudinal axis X. This fan stator 58 is here arranged downstream of the fan rotor 54. The external circulation vein 52 is here defined between the nacelle 20 and a fan casing 60 surrounding the fan 50. The turbomachine 12 is typically constituted by a turbojet with a high dilution ratio (in English “bypass ratio”),the dilution ratio being defined as the ratio of the flow rate of the secondary flow B (cold) to the flow rate of the primary flow A (hot). Alternatively (not shown), the turbomachine 12 is devoid of a fan casing 60, that is to say that the external circulation vein 52 has no peripheral delimitation. The turbomachine 12 is then constituted by a turbojet with an unducted fan or by a turboprop. A rotor stage 48 of the low-pressure compressor 41 is shown in more detail in Figure 3. As visible in this Figure, the rotor stage 48 comprises, in addition to the blades 49, a metal disc 62 for supporting the blades 49. With reference to Figure 4, each blade 49 is elongated in a direction of elongation Y that is substantially radial, that is to say perpendicular to the longitudinal axis X. In the following,“height” will be used to designate a distance along the elongation axis Y. Each blade 49 has an aerodynamic profile shaped so as to generate lift when it is moved in an air flow. Thus, each blade 49 comprises, as visible in Figure 4, a lower surface 70, an upper surface 72, a leading edge 74 constituting an upstream edge of the blade 49, a trailing edge 76 constituting a downstream edge of the blade 49, and a chord (not shown), orthogonal to the elongation axis Y, connecting the leading edge 74 to the trailing edge 76. In the example shown, each blade 49 is twisted around its elongation axis Y, that is to say that its chord pivots around said elongation axis Y when moving along the elongation axis Y. Each blade 49 also has, still with reference to Figure 4, a proximal end 77 for connection to the disk 62 and an opposite, free distal end 78. At least one of the blades 49, here several blades 49,is constituted by a hybrid blade 79 comprising a proximal portion 80 forming a heel, close to the disc 62, made of metal, and a distal portion 82, distant from the disc 62, composed at least in part of a structure made of composite material (not shown). The proximal portion 80 and the distal portion 82 each have an aerodynamic profile. The heel 80 delimits the proximal end 77 of the blade 79, which is substantially planar. It typically extends over a height of between 10 and 20% of the height of the blade 79. Alternatively, the heel 80 extends over a height less than 10% of the height of the blade or over a height greater than 20% of the height of the blade. The metal constituting the heel 80 is for example titanium or a titanium alloy. In a first embodiment, illustrated in Figures 16 and 17, the heel 80 is fixed to the composite material structure by means of at least one pin 81,here two pins 81. Each pin 81 is embedded in the composite material structure and shrunk into the heel 80. Advantageously, the pin(s) 81 has(have) an optimized geometry so as to ensure good retention of the distal portion 82 to the heel 80 despite the centrifugal forces to which the blade 79 is subjected. In a second embodiment, illustrated in Figures 18 and 19, the heel 80 is fixed to the composite material structure by means of a metal spar 83 integral with the heel 80 and extending into the composite material. This metal spar 83 is here elongated along the elongation direction Y. It is typically made of the same material as the heel 80. Advantageously, the metal spar 83 has an optimized geometry so as to ensure good retention of the distal portion 82 at the heel 80 despite the centrifugal forces to which the blade 79 is subjected. In a third embodiment, illustrated in Figures 20 and 21,the heel 80 is fixed to the composite material structure by a bathtub-type fixing, that is to say by means of ribs 85 integral with the heel 80 between which the composite material structure is embedded. Each rib 85 projects in the direction of elongation Y from the face 87 of the heel 80 oriented towards the distal portion 82. Preferably, each rib 85 is positioned, as shown, at the periphery of said face 87 and is flush with an edge of said face 87. Here, the ribs 85 together delimit a closed contour in a plane orthogonal to the axis of elongation Y. The ribs 85 are typically made of the same material as the heel 80. Advantageously, the ribs 85 have an optimized geometry so as to ensure good retention of the distal portion 82 at the heel 80 despite the centrifugal forces to which the blade 79 is subjected. As a variant,the heel 80 is fixed to the composite material structure by conventional means known to those skilled in the art. It is for example screwed to the composite material structure, or introduced directly into a mold in which the fiber reinforcement of the composite material structure is deposited for co-injection during the injection step of the matrix of the composite material structure (typically in the case where the composite material structure is produced by resin transfer molding). Returning to Figure 4, the distal portion 82 extends over the majority of the height of the blade 79, advantageously over at least 80% of the height of the blade 79, for example at least 85%. It extends from the heel 80 to the distal end 78 of the blade 79. The distal portion 82 is in line with the heel 80 at the intrados 70 and at the extrados 72. In other words,the portions of the intrados 70 and the extrados 72 carried by the distal portion 82 are flush with the portions of the intrados 70 and the extrados 72 carried by the heel 80, respectively. Thus, there is no roughness or shoulder at the junction between the distal portion 82 and the heel 80, which avoids disturbing the aerodynamic flow and guarantees good mechanical strength. The composite material structure extends over the entire height of the distal portion 82. In other words, it extends from the heel 80 to the distal end 78 of the blade 79. The composite material structure comprises a fibrous reinforcement (not shown) and a matrix (not shown) in which the fibrous reinforcement is embedded. The fibrous reinforcement is, for example, obtained by three-dimensional weaving. For this purpose, it is typically formed from a single-piece fiber preform with varying thickness comprising warp strands and weft strands,these strands comprising for example carbon, glass, basalt, and / or aramid fibers. Said fiber preform is advantageously obtained by three-dimensional or multi-layer weaving, that is to say that the warp strands follow sinuous paths in order to link together weft strands belonging to different weft strand layers, it being noted that said three-dimensional weaving may include 2D surface weaves. Different three-dimensional weaving weaves may be used, such as interlock, multi-satin or multi-veil weaves, for example, as described in particular in document WO 2006 / 136755. The fiber reinforcement is then embedded in the matrix, for example by means of the technique known as resin transfer molding (better known by the acronym RTM, from the English “Resin Transfer Molding”). Alternatively,the composite material structure is obtained by draping different layers of prepreg each comprising the fiber reinforcement and the matrix. Preferably, the composite material structure constitutes the majority of the blade 79, so that the blade 79 is thus composed mainly of composite material. The expression “composed mainly” is understood here and hereinafter in mass proportions, that is to say that the majority component (here the composite material) constitutes more than 50% by weight of the object (here the blade 79). The composite material structure advantageously constitutes at least 70% by weight of the blade 79. Alternatively, the composite material structure is in the minority, the blade 79 then being composed mainly of metal. As a further alternative, the blade 79 is composed substantially half of metal and, for the other half, of composite material. Advantageously, the distal portion 82 is also composed of a metal part,here comprising the spar 83 or the ribs 85, extending the heel 80 inside and / or around the composite material structure. This metal part is preferably made of the same material as the heel 80. It extends over all or part of the height of the distal portion 82. Preferably, the distal portion 82 then comprises a transition section 84 in contact with the heel 80 in which the density of the metal part decreases as the distance from the heel 80 increases. This ensures good mechanical strength of the blade 79. This transition section 84 extends over all or part of the distal portion 82. Returning to Figure 3, in the example shown, at least one of the blades 49, here two of the blades 49, is constituted by a blade 86 which is not a hybrid blade 79. This blade 86 comprises a blade 88 intended to extend into the air stream,said blade 88 delimiting the distal end 78 of the vane 86 and forming the part of the vane 86 which has an aerodynamic profile. It also comprises a root (not shown) forming the proximal end 77 of the vane 86. This root typically comprises a bulb and a stilt connecting the bulb to the blade 88, the bulb being connected to the stilt by a neck defining a local minimum of the section of the root. Said vane 86 is for example a metal vane, that is to say it is composed solely of metal, or a composite vane, that is to say the structure of the vane 86 is composed solely of composite material. For example, the vanes 49 comprise several vanes 86, said vanes 86 being all metal vanes, all composite vanes, or for some composite vanes and for others metal vanes. Alternatively (not shown), the blades 49 are all hybrid blades 79. With reference to Figure 5, the disc 62 comprises a large upstream face 90,a large downstream face 91 (Figure 7) and, at its outer periphery, a slice 92 connecting said large 90, 91 faces to each other. The disc 62 also has a central through orifice 94 opening into each of the large faces 90, 91, so that the disc 62 thus has an annular shape. The slice 92 delimits an outer peripheral edge of the disc 62. It has a shape of revolution, for example cylindrical of revolution or frustoconical. In particular, the edge 92 partly delimits the internal vein 22. The disc 62 also comprises a plurality of bases 96 each projecting radially outwards from the edge 92. Their number is equal to that of the hybrid blades 79. Thus, the disc 62 comprises, for each hybrid blade 79, a base 96. The bases 96 are preferably in one piece with the rest of the disc 62. They are typically machined from the same block of material as the rest of the disc 62. As seen in Figure 6,each base 96 has, opposite the edge 92, a distal end 100. This distal end 100 is substantially planar and substantially perpendicular to a radial direction of the disc 62, that is to say that there is a radial direction of the disc 62 passing through the distal end 100 and to which the latter is substantially orthogonal. Each base 96 also has a peripheral surface 102 bordering the distal end 100. This peripheral surface 102 is connected to the edge 92 by a fillet 104. This fillet 104 goes around the base 96. Thus, the transition between the edge 92 and the peripheral surface 102 of the base 96 is done progressively, which avoids aerodynamic disturbances. The distal end 100 constitutes a connecting surface 106 to which the heel 80, more specifically the proximal end 77, of a respective hybrid blade 79 is connected, preferably welded. Thus,the junction between the disc 62 and the hybrid blade 79 is at a distance from the edge 92, which guarantees good mechanical strength. Indeed, the main stresses are thus concentrated in the base 96 which, because it is in one piece with the rest of the disc 62, has better resistance. This connecting surface 106 is of a shape substantially identical to that of the proximal end 77 of the hybrid blade 79 and the hybrid blade 79 is positioned relative to the base 96 so that its intrados 70, its extrados 72, its leading edge 74 and its trailing edge 76 are each flush with the peripheral surface 102 of the base 96. Thus, the junction between the hybrid blade 79 and the disc 62 is free of roughness, which avoids disturbing the aerodynamic flow and guarantees good mechanical strength. The disc 62 is made of metal. This metal is advantageously of the same nature as the metal constituting the heel 80 of the hybrid blades 79,that is to say that the majority metallic element constituting each of said metals is identical to the majority metallic element constituting the other metals. Thus the metal constituting the disc 62 and that constituting the heel 80 of the hybrid blades 79 are for example identical metals, or different alloys of the same base metal. This ensures good mechanical strength of the hybrid blades 79 to the disc 62. In the example shown, the disc 62 also comprises cells 108 arranged in the edge 92. Their number is equal to that of the blades 86. Thus, the disc 62 comprises, for each blade 86, a cell 108. These cells 108 have a shape substantially complementary to that of the roots of the blades 86 and receive said roots when the blades 86 are fixed to the disc 62. The roots of the blades 86 cooperate with said cells 108 to keep the blades 86 attached to the disc 62. The bases 96 and the cells 108 are regularly distributed along the edge 92,that is to say that each base 96 or cell 108 is substantially equidistant from each adjacent base 96 or cell 108. In the variant where the blades 49 are all hybrid blades 79, the disk 62 does not include any cell 108. The bases 96 are then regularly distributed along the section 92, that is to say that each base 96 is substantially equidistant from each adjacent base 96. A method 200 for manufacturing the rotor 46 will now be described, with reference to Figures 9 to 15. The method 200 begins with a first step 210 for manufacturing a rotor stage 48 of the rotor 46. This first step 210 comprises the supply 212 of the disk 62 and the supply 213 of the hybrid blades 79. These supply steps 212, 213 are concomitant or, as shown,successive to one another. The supply 212 of the disc 62 typically comprises the machining of the entire disc 62 in a single block of metal. The supply 213 of the hybrid blades 79 typically comprises the machining of the heel 80 and the metal parts of the distal portion 82 in a single block of metal, the manufacture of the composite structure, and the assembly of the heel 80 and the metal parts of the distal portion 82 to the composite structure. The supply sub-steps 212, 213 are followed by a sub-step 214 of connecting the heel 80 of each hybrid blade 79 to the disc 62. During this sub-step 214, each blade 79 is in turn placed opposite a respective base 96 of the disc 62, as shown in Figures 10 and 11, then the proximal end 77 of the blade 79 is brought into contact with the connecting surface 106 delimited by the base 96 and welded to the latter. This welding is advantageously carried out by friction,for example by linear or orbital friction. Optionally, the welding uses a filler metal; this filler metal is then of the same nature as the metals making up the blade 79 and the disc 62. Thanks to the flatness of the connecting surface 106 and that of the proximal end 77, this welding is facilitated. Indeed, the plane-to-plane contact of the two surfaces to be welded 106, 77 allows friction welding which would be difficult to implement if the surfaces had more complex shapes, for example if one of the two surfaces had a shape of revolution like the edge 92. It will be noted that this flatness of the surfaces is allowed by the fact that the connecting surface 106 is not directly carried by the edge 92 but by a base 96 projecting from the edge 92. Alternatively, the proximal end 77 of the blade 79 is not welded but brazed to the connecting surface 106. As visible in Figures 12 and 13,the welding (or brazing) of the heel 80 to the base 96 produces a weld bead 110 at the periphery of the connecting surface 106. This weld bead 110 would risk disturbing the aerodynamic flow at the foot of the blade 79. To remedy this, the connecting sub-step 214 is followed by a sub-step 215 of machining said weld bead 110. During this sub-step 215, the weld bead 110 is machined so as to make the periphery of the junction between the blade 79 and the base 96 smooth and regular, as visible in Figures 14 and 15. The weld bead 110 being placed at a distance from the edge 92 by the base 96, its machining is facilitated. Advantageously, sub-step 215 is implemented in one go after all the hybrid blades 79 have been welded or brazed to the disc 62. Alternatively, sub-step 215 is implemented in several goes, after each connection of a hybrid blade 79 to the disc 62. The manufacturing 210 of the rotor stage 48 also comprises, where appropriate,the supply 216 of the non-hybrid blades 86 and the assembly 217 of the latter to the disc 62 by inserting their roots into the cells 108 of the disc 62. The manufacture 210 of the rotor stage 48 concludes with a sub-step 218 of static balancing of said rotor stage 48, during which material is added and / or removed from the rotor stage 48 so as to ensure its static balancing. This static balancing of the rotor stage 48 is for example obtained by machining a circumferential bead (not shown) formed in the disc 62 or adding weights (not shown) fixed to the disc 62. Step 210 is repeated for each rotor stage 48 of the rotor 46. These rotor stages 48 are then assembled to each other during a subsequent assembly step 220. The rotor 46 is thus obtained. Thanks to the embodiment described above, it is thus possible to obtain a low-pressure compressor rotor 46 which is both light,resistant, and simple to produce. Furthermore, because the hybrid blades 79 comprise a composite material structure, it is easier to adjust their mechanical properties. It will be noted that, although the embodiment described above relates to a low-pressure compressor rotor, the invention is in no way limited to this single embodiment and applies to any type of turbomachine rotor. In particular, the above description of the rotor stage 48 and the manufacturing method 200 is applicable to a fan rotor such as the rotor 54.,

Claims

CLAIMS 1. Rotor element (48) for a turbomachine comprising a metal disc (62) and a plurality of blades (49) mounted on the disc (62), wherein at least one of said blades (49) is constituted by a hybrid blade (79) composed of composite material and comprising a metal heel (80) bonded to the disc (62) by welding or brazing.

2. Rotor element (48) according to claim 1, wherein the disc (62) has an edge (92) and, for the or each hybrid blade (79), a connecting surface (106), arranged on said edge (92), to which the heel (80) of said hybrid blade (79) is bonded.

3. Rotor element (48) according to claim 2, wherein the connecting surface (106) is substantially planar.

4. A rotor element (48) according to claim 3, wherein the connecting surface (106) is substantially perpendicular to a radial direction of the disc (62). 5.Rotor element (48) according to any one of claims 2 to 4, wherein the disc (62) comprises, for the or each hybrid blade (79), a base (96) projecting radially outwardly from the edge (92), said base (96) having, opposite the edge (92), a distal end (100) constituting the connecting surface (106) of said hybrid blade (79).

6. Rotor element (48) according to claim 5, wherein the base (96) has a peripheral surface (102) bordering the distal end (100) and the hybrid blade (79) is positioned relative to the base (96) so that its intrados (70), its extrados (72), its leading edge (74) and its trailing edge (76) are each flush with the peripheral surface (102) of the base (96).

7. A rotor element (48) according to any one of claims 2 to 6, wherein the wafer (92) has a revolution shape. 8.Rotor element (48) according to any one of claims 2 to 7, in which the edge (92) partly delimits an air stream (22) of the turbomachine (12).

9. Rotor element (48) according to any one of the preceding claims constituting a rotor stage of a compressor or fan.

10. Rotor element (48) according to any one of the preceding claims, wherein the heel (80) has an aerodynamic profile.

11. Rotor element (48) according to any one of the preceding claims, wherein the heel (80) is made of metal and the hybrid blade (79) comprises a distal portion (82), remote from the disc (62), composed at least in part of a structure made of composite material.

12. Rotor element (48) according to claim 11, wherein said distal portion (82) is in the extension of the heel (80) at the intrados (70) and at the extrados (72).

13. Rotor element (48) according to claim 11 or 12, in which the heel (80) is fixed to the composite material structure by means of at least one pin (81) embedded in the composite material structure and shrunk into the heel (80). 14.Rotor element (48) according to any one of claims 11 to 13, wherein the heel (80) is fixed to the composite material structure by means of a metal spar (83) integral with the heel (80) and extending into the composite material.

15. Rotor element (48) according to any one of claims 11 to 14, wherein the heel (80) is fixed to the composite material structure by a bathtub-type fixing.

16. Turbomachine (12) comprising a rotor element (48) according to any one of the preceding claims.

17. Aircraft (10) comprising at least one turbomachine (12) according to claim 16.

18. Method (210) for manufacturing a rotor element (48) according to any one of claims 1 to 15, comprising the following steps: - providing (212) a metal disc (62), - providing (213) at least one hybrid blade (79) composed of composite material and comprising a metal heel (80), and. - connection (214) of the heel (80) of the or each hybrid blade (79) to the disc (62), said connection (214) being made by welding or brazing.

19. Manufacturing method (210) according to claim 18, comprising an additional step of static balancing (218) of the rotor element (48), for example by machining a circumferential bead formed in the disc (62) or adding weights fixed to the disc (62).

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