Rotor element for turbine engine having composite blade connected to metal disc

Abstract

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

Description

Technical Field

[0001] The present invention relates to the field of turbines, and more particularly to the field of rotors of the type for turbines including blades made of composite materials. Background Art

[0002] Generally, a turbine rotor (especially a fan rotor and a compressor rotor) is formed by blades mounted around a hub, which is generally referred to as a "disk" and is usually metallic. Generally, the blades themselves are also metallic and are attached to the disk by broaching fasteners that form the root of the blade, which is received in a unit provided on the outer periphery of the disk.

[0003] Various solutions have been proposed with the aim of reducing the weight of the turbine rotor, thereby improving consumption and contributing to healthy shaft dynamics.

[0004] For example, for a compressor rotor, the use of an integrally bladed rotor (more widely known by the acronym "DAM" or the name "blisk") has been proposed, in which the blades are integral with the disk. When using broaching fasteners, this solution reduces the weight of the rotor by reducing the number of parts and eliminating the need for a sealing system. To this end, the blades are generally machined from the same material block as the disk or welded to the disk.

[0005] For a fan rotor, the use of blades at least partially composed of a composite structure has been proposed, the composite structure including a fiber reinforcement densified by a polymer matrix, wherein the blades are always attached to the disk by broaching fasteners. This solution makes it possible to reduce the weight of the rotor by reducing the weight of the blades, since composite blades are lighter than metallic blades with equivalent propulsion characteristics.

[0006] However, these solutions are not entirely satisfactory. More specifically, compared with the solutions using composite materials, the single-piece blisk remains relatively dense, and the solutions using composite materials have a complex structure due to the necessity of using a sealing system. Summary of the Invention

[0007] An object of the present invention is to further reduce the weight of the turbine rotor. Other objects are to simplify the structure of the turbine rotor and ensure good mechanical strength.

[0008] To this end, a first aspect of the present invention relates to a rotor element for a turbine, the rotor element including a metallic disk and a plurality of blades mounted on the disk, wherein at least one of the blades is composed of a hybrid blade made of a composite material and including a metallic root connected to the disk.

[0009] According to a particular embodiment of the present invention, the rotor element further has one or more of the following features, which can be employed individually or in any technically feasible combination:

[0010] - Each blade in the blades is composed of a hybrid blade;

[0011] - At least one other blade in the blades is composed of a metal blade;

[0012] - At least one other blade in the blades is composed of a composite blade;

[0013] - The other blade or each other blade is composed of a metal blade;

[0014] - The other blade or each other blade is composed of a composite blade;

[0015] - The hybrid blade or each hybrid blade is mainly made of a composite material;

[0016] - The root is connected to the disk by welding or brazing;

[0017] - The disk has an edge, and for the hybrid blade or each hybrid blade, the disk has a connection surface arranged on the edge, and the root of the hybrid blade is connected to the connection surface;

[0018] - The connection surface is substantially planar;

[0019] - The connection surface is substantially perpendicular to the radial direction of the disk;

[0020] - For the hybrid blade or each hybrid blade, the disk includes a base projecting radially outward from the edge, the base having a distal end opposite to the edge, and the distal end constituting the connection surface for the hybrid blade;

[0021] - The base has an outer peripheral surface defining the distal end, and the hybrid blade is positioned relative to the base such that the belly, back, leading edge, and trailing edge of the hybrid blade are flush with the outer peripheral surface of the base respectively;

[0022] - The edge has a rotationally symmetric shape;

[0023] - The edge partially defines the air flow path of the turbine;

[0024] - The rotor element constitutes a compressor rotor stage or a fan rotor stage;

[0025] - The root has an aerodynamic profile; and

[0026] - The root is made of metal, and the hybrid blade includes a distal portion away from the disk, the distal portion being at least partially composed of a composite material structure, and the distal portion is an extension of the root at the belly and back.

[0027] According to a second aspect, another object of the present invention is a turbine, which comprises a rotor element according to the first aspect.

[0028] According to a third aspect, yet another object of the present invention is an aircraft, which comprises at least one turbine according to the second aspect.

[0029] Finally, according to a fourth aspect, another object of the present invention is a method for manufacturing a rotor element, which is the rotor element according to the first aspect, and the method comprises the following steps:

[0030] - providing a metal disk,

[0031] - providing at least one hybrid blade, which is made of a composite material and comprises a metal root, and

[0032] - connecting the root of the hybrid blade or the root of each hybrid blade to the disk.

[0033] According to a specific embodiment of the present invention, the manufacturing method further has one or more of the following features, which can be adopted alone or in any technically feasible combination:

[0034] - the root of each hybrid blade is connected to the disk by welding or brazing;

[0035] - the welding of the root of each hybrid blade to the disk is generated by friction, for example, by linear friction or orbital friction;

[0036] - the connection of the root of each hybrid blade to the disk generates at least one weld bead, and the manufacturing method comprises an additional step of processing the weld bead; and

[0037] - the method comprises the following additional steps: keeping the rotor element in static balance, for example, by processing a circumferential weld bead formed in the disk or by adding a weight attached to the disk. Description of the Drawings

[0038] Other features and advantages of the present invention will become apparent by reading the following description, which is provided only by way of example and with reference to the accompanying drawings, in which:

[0039] - Figure 1 is a view observed from above of an aircraft according to an exemplary embodiment of the present invention,

[0040] - Figure 2 is Figure 1 a longitudinal sectional view of the upper half of the turbine of the aircraft of

[0041] - Figure 3 isFigure 2 Three-quarter front perspective view of the rotor stage of the compressor of a turbine

[0042] - Figure 4 is Figure 3 Three-quarter front perspective view of the blade of the rotor stage

[0043] - Figure 5 is Figure 3 Front view of the disk of the rotor stage

[0044] - Figure 6 is Figure 5 View of the detail marked VI in

[0045] - Figure 7 is along Figure 3 Cross-sectional view of the rotor stage taken along the sectional plane marked VII-VII in Figure 3

[0046] - Figure 8 is Figure 7 View of the detail marked VIII in

[0047] - Figure 9 is a diagram showing the manufacturing method of the rotor of the rotor stage including Figure 3

[0048] - Figure 10 is during Figure 9 the first manufacturing stage of the method in Figure 3 Side view of the rotor stage

[0049] - Figure 11 is Figure 10 View of the detail marked XI in

[0050] - Figure 12 is during Figure 9 the second manufacturing stage of the method in Figure 3 Side view of the rotor stage

[0051] - Figure 13 is Figure 12 View of the detail marked XIII in

[0052] - Figure 14 is during Figure 9 the third manufacturing stage of the method in Figure 3 Side view of the rotor stage

[0053] - Figure 15 is Figure 14 View of the detail marked XV in

[0054] - Figure 16 ​​is taken in a plane oriented parallel to the extension axis of the blade in the first embodiment of the present invention Figure 4 a cross-sectional view of a portion of the blade, the portion of the blade including the weld bead of the blade,

[0055] - Figure 17 is taken in a plane orthogonal to the weld bead of the blade according to the first embodiment of the present invention Figure 4 a cross-sectional view of the portion of the blade,

[0056] - Figure 18 and Figure 19 are views similar to Figure 16 and Figure 17 respectively according to the second embodiment of the present invention, and

[0057] - Figure 20 and Figure 21 are views similar to Figure 16 and Figure 17 respectively according to the third embodiment of the present invention. Detailed Description

[0058] Figure 1 The illustrated aircraft 10 includes turbines 12 to propel the aircraft.

[0059] In the illustrated example, the aircraft 10 is an airplane. In a conventional manner, the aircraft includes a fuselage 14, a tail assembly 16, and two wings 18. Here, there are two turbines 12, each turbine 12 being mounted below a respective wing 18. In an alternative (not shown), the turbines 12 are disposed along the fuselage 14, for example, near the tail assembly 16. In another alternative (also not shown), the aircraft 10 includes a single turbine 12 or at least three turbines 12.

[0060] Figure 2 One of the turbines 12 is shown.

[0061] As can be seen in this figure, the turbine 12 is elongated along the longitudinal axis X. Generally, the turbine has angular symmetry about the longitudinal axis X, in other words, there is at least one angle at which the turbine remains unchanged when rotating about the longitudinal axis X.

[0062] Herein and hereinafter, the terms "inner" and "outer", "inside" and "outside" and their variants are determined with reference to the axis X, and an element regarded as "inner" or "inside" is oriented towards the axis X, while an "outer" or "outside" element is oriented in a direction opposite to the axis X. The term "radial" and its variants are determined with reference to a direction orthogonal to the axis X.

[0063] Typically, the turbine 12 includes a nacelle 20, an internal flow path 22 for allowing an air flow to flow through the nacelle 20, a combustion chamber 24 accommodated in the flow path 22, a motor body 26, and an exhaust nozzle 28.

[0064] Hereinafter, the terms "upstream" and "downstream" are determined with reference to the flow direction of the air flow flowing through the flow path 22.

[0065] The motor body 26 includes a compressor 30, a turbine 32, and a drive shaft 34 that couples the turbine 32 to the compressor 30 to drive the compressor 30 by means of the turbine 32.

[0066] The compressor 30 is provided upstream of the combustion chamber 24 and supplies compressed air to the combustion chamber 24. The compressor includes a stator 35 and a rotor 36. The stator 35 is integral with the nacelle 20. The rotor 36 is rotatable relative to the stator 35 by means of the drive shaft 34; for this purpose, the rotor 36 is typically fixedly mounted on the drive shaft 34.

[0067] The stator 35 includes at least one (in this case, a plurality of) stator stages 37, each stator stage 37 being formed by blades (not shown) substantially arranged in the same radial plane. Similarly, the rotor 36 includes at least one (in this case, a plurality of) rotor stages 38, each rotor stage 38 being formed by blades (not shown) substantially arranged in the same radial plane. There is the same number of stator stages 37 and rotor stages 38, and these stator stages 37 and rotor stages 38 alternate with each other.

[0068] The turbine 32 is provided downstream of the combustion chamber 24 and receives the exhaust gas leaving the combustion chamber 24.

[0069] The longitudinal axis X is the axis of rotation of the drive shaft 34.

[0070] The drive shaft 34 is guided to rotate relative to the nacelle 20 by means of bearings (not shown).

[0071] In the example shown, the turbine 12 is a multi-body turbine, particularly a two-body turbine. The turbine 12 includes a low-pressure body 40 in addition to the engine body 26. Therefore, the engine body 26 constitutes the high-pressure body, the compressor 30 is a high-pressure compressor, the turbine 32 is a high-pressure turbine, and the drive shaft 34 is a high-pressure shaft.

[0072] The low-pressure body 40 includes a low-pressure compressor 41, a low-pressure turbine 42, and a low-pressure shaft 43 that couples the low-pressure turbine 42 to the low-pressure compressor 41 to drive the low-pressure compressor 41 by means of the low-pressure turbine 42.

[0073] The low-pressure compressor 41 is provided upstream of the high-pressure compressor 30 and supplies compressed air to the high-pressure compressor 30. The low-pressure compressor includes a stator 45 and a rotor 46. The stator 45 is integral with the nacelle 20. The rotor 46 is rotatable relative to the stator 45 by means of a low-pressure shaft 43; for this purpose, the rotor 46 is generally fixedly mounted on the low-pressure shaft 43.

[0074] The stator 45 includes at least one (in this case a plurality of) stator stages 47, each stator stage 47 being formed by blades (not shown) substantially arranged in the same radial plane. Similarly, the rotor 46 includes at least one (in this case a plurality of) rotor stages 48, each rotor stage 48 being formed by blades 49( Figure 3 ) forming. There are the same number of stator stages 47 and rotor stages 48, and these stator stages 47 and rotor stages 48 alternate with each other.

[0075] The low-pressure turbine 42 is provided downstream of the high-pressure turbine 32 and receives the exhaust gas leaving the high-pressure turbine 32.

[0076] The low-pressure shaft 43 is guided to rotate relative to the nacelle 20 by means of bearings (not shown).

[0077] The low-pressure shaft 43 is coaxial with the high-pressure shaft 34. Thus, the axis of rotation of the low-pressure shaft 43 is also the longitudinal axis X. In particular, the low-pressure shaft 43 extends inside the high-pressure shaft 34.

[0078] Here, the turbine 12 also includes a fan 50 for driving an air flow in an external circulation flow path 52 around the nacelle 20. Thus, the main air flow A (hot), which is constituted by a part of the air flow driven in the internal circulation flow path 22, can be distinguished from the secondary air flow B (cold), which is constituted by a part of the air flow driven in the external circulation flow path 52.

[0079] The fan 50 includes a fan rotor 54. The fan rotor 54 can be rotatably mounted relative to the nacelle 20 about the longitudinal axis X. The fan rotor 54 includes a disk 55, and the disk 55 is provided with fan blades 56 that extend substantially radially outward from the disk 55. These blades 56 drive the air flow in the external circulation flow path 52 when rotating.

[0080] The low-pressure turbine 42 rotates the fan rotor 54 via the low-pressure shaft 43. In the example shown, this drive is achieved by means of a speed reducer 57, such that the fan rotor 54 can rotate at a speed lower than the speed of the low-pressure shaft 43. Alternatively (not shown), this drive is direct, in other words, the fan rotor 54 is constrained to rotate with the low-pressure shaft 43.

[0081] In particular, the fan rotor 54 is arranged upstream of the internal flow path 22 and also drives the air flow in the internal flow path 22.

[0082] In the example shown, the fan 50 also includes a fan stator 58 (also referred to as a stator vane), the fan stator 58 including fixed vanes 59 which are arranged in the external flow path 52 at the outer periphery of the nacelle 20 in a plane orthogonal to the longitudinal axis X. Here, the fan stator 58 is arranged downstream of the fan rotor 54.

[0083] Here, the external flow path 52 is defined between the nacelle 20 and a fan housing 60 surrounding the fan 50. The turbine 12 generally consists of a turbofan engine with a high bypass ratio, the bypass ratio being defined as the ratio of the flow rate of the secondary flow B (cold) divided by the flow rate of the primary flow A (hot).

[0084] Alternatively (not shown), the turbine 12 does not have a fan housing 60, in other words, the external flow path 52 does not have an outer periphery defined. Thus, the turbine 12 consists of a turbofan engine or a turboprop engine with an unducted fan.

[0085] Figure 3 The rotor stage 48 of the low-pressure compressor 41 is shown in more detail in. As can be seen in this figure, in addition to the blades 49, the rotor stage 48 also includes a metal disk 62 for supporting the blades 49.

[0086] Reference Figure 4 , each blade 49 extends along a substantially radial elongation direction Y (in other words, perpendicular to the longitudinal axis X). Hereinafter, "height" means the distance along the elongation axis Y.

[0087] Each blade 49 has an aerodynamic profile shaped to generate lift when the blade 49 moves in an air flow. Thus, as can be seen in Figure 4 , each blade 49 includes a belly 70, a back 72, a leading edge 74 forming the upstream edge of the blade 49, a trailing edge 76 forming the downstream edge of the blade 49, and a weld bead (not shown) which is orthogonal to the elongation axis Y and connects the leading edge 74 to the trailing edge 76.

[0088] In the example shown, each blade 49 is twisted about its elongation axis Y, in other words, when each blade 49 moves along the elongation axis Y, the weld bead of each blade 49 pivots about the elongation axis Y.

[0089] Still referring to Figure 4 , each blade 49 also has a proximal end 77 for connection to the disk 62 and an opposite free distal end 78.

[0090] At least one of the blades 49 (here a plurality of blades 49) consists of hybrid blades 79, the hybrid blades 79 comprising a proximal portion 80 and a distal portion 82, the proximal portion 80 forming a root, being close to the disk 62 and being made of metal, and the distal portion 82 being remote from the disk 62 and being made at least in part of a composite material structure (not shown). The proximal portion 80 and the distal portion 82 each have an aerodynamic profile.

[0091] The root 80 defines the proximal end 77 of the blade 79, and the proximal end 77 is substantially planar. The root 80 typically extends at a height between 10% and 20% of the height of the blade 79. Alternatively, the root 80 extends at a height less than 10% of the blade height or greater than 20% of the blade height.

[0092] The metal forming the root 80 is, for example, titanium or a titanium alloy.

[0093] In a first embodiment, as Figure 16 and Figure 17 shown, the root 80 is attached to the composite material structure by means of at least one pin 81 (in this case, two pins 81). Each pin 81 is embedded in the composite material structure and is shrink-fitted into the root 80. Advantageously, one or more pins 81 have an optimized geometry to ensure good strength of the distal portion 82 at the root 80, regardless of the centrifugal forces to which the blade 79 is subjected.

[0094] In a second embodiment, in Figure 18 and Figure 19 it is shown that the root 80 is attached to the composite material structure by means of a metal beam 83, the metal beam 83 being integral with the root 80 and extending in the composite material. In this case, the metal beam 83 extends along the elongation direction Y. The metal beam 83 is typically integral with the root 80. Advantageously, the metal beam 83 has an optimized geometry to ensure good strength of the distal portion 82 at the root 80, regardless of the centrifugal forces to which the blade 79 is subjected.

[0095] In a third embodiment, as Figure 20 and Figure 21As shown, the heel 80 is attached to the composite structure by a bathtub-type fastener. In other words, the heel 80 is attached to the composite structure by means of ribs 85 integral with the heel 80 (the composite structure being embedded between the ribs 85). Each rib 85 projects from a face 87 of the heel 80 that is oriented towards the distal portion 82 along the elongation direction Y. Preferably, as shown, each rib 85 is positioned at the outer periphery of the face 87 and flush with the edge of the face 87. Here, the ribs 85 jointly define a closed contour in a plane orthogonal to the elongation axis Y. The ribs 85 are generally integral with the heel 80. Advantageously, the ribs 85 have an optimized geometry to ensure good strength of the distal portion 82 at the heel 80, regardless of the centrifugal forces to which the blade 79 is subjected.

[0096] Alternatively, the heel 80 is attached to the composite structure by conventional means known to those skilled in the art. For example, the heel 80 is screwed onto the composite structure or directly introduced into a mold in which the fiber reinforcement of the composite structure is deposited for co-injection during the injection step of the matrix of the composite structure (usually in the case where the composite structure is produced by resin transfer molding).

[0097] Returning to Figure 4 , the distal portion 82 extends over most of the height of the blade 79. Advantageously, the distal portion 82 extends over at least 80% (e.g., at least 85%) of the height of the blade 79. The distal portion 82 extends from the heel 80 to the distal end 78 of the blade 79.

[0098] The distal portion 82 is an extension of the heel 80 to the soffit 70 and the back 72. In other words, the portions of the soffit 70 and the back 72 carried by the distal portion 82 are flush with the portions of the soffit 70 and the back 72 carried by the heel 80. Thus, there is no sharpness or shoulder at the junction between the distal portion 82 and the heel 80, which avoids disturbing the aerodynamic flow and ensures good mechanical strength.

[0099] The composite structure extends over the entire height of the distal portion 82. In other words, the composite structure extends from the heel 80 to the distal end 78 of the blade 79.

[0100] The composite structure includes a fiber reinforcement (not shown) and a matrix (not shown) in which the fiber reinforcement is embedded.

[0101] The fiber reinforcement is obtained, for example, by three-dimensional weaving. To this end, the fiber reinforcement is generally formed from a one-piece fiber preform having a varying thickness, the one-piece fiber preform comprising warp yarns and weft yarns, these yarns comprising, for example, carbon, glass, basalt and / or aramid fibers. Advantageously, the fiber preform is obtained by three-dimensional weaving or multi-layer weaving (i.e., the warp yarns follow a curved path to interconnect the weft yarns belonging to different weft yarn layers), it being noted that the three-dimensional weaving may include two-dimensional surface weaving. For example, as described in particular in document WO 2006 / 136755, various three-dimensional weavings may be used, such as interlock weaving, multi-satin weaving or multi-mesh weaving. The fiber reinforcement is then embedded in the matrix, for example using a technique known as resin transfer molding (more widely known by the abbreviation RTM).

[0102] Alternatively, the composite structure is obtained by covering individual prepreg layers, each prepreg layer comprising a fiber reinforcement and a matrix.

[0103] Preferably, the composite structure forms the majority of the blade 79, such that the blade 79 thus consists mainly of the composite material. The expression "consisting mainly of" should be understood herein and hereinafter as by mass ratio, in other words, the main component (here the composite material) constitutes more than 50% of the weight of the object (here the blade 79). Advantageously, the composite structure constitutes at least 70% of the weight of the blade 79.

[0104] Alternatively, the composite structure is a minor component, and thus the blade 79 consists mainly of metal. As another variant, the blade 79 is substantially half composed of metal and half composed of composite material.

[0105] Advantageously, the distal portion 82 is also composed of metal parts, the distal portion 82 including herein a beam 83 or a rib 85, extending inside and / or around the composite structure along the root 80. Preferably, this metal part is integral with the root 80. This metal part extends over all or part of the height of the distal portion 82.

[0106] Preferably, the distal portion 82 includes a transition section 84 in contact with the root 80, wherein the density of the metal part decreases as the distance from the root 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.

[0107] Return to Figure 3, in the example shown, at least one of the vanes 49 (here two of the vanes 49) is formed by a vane 86 that is not a hybrid vane 79. The vane 86 includes an airfoil 88 intended to extend into the air flow, the airfoil 88 defining a distal end 78 of the vane 86 and forming the aerodynamically profiled portion of the vane 86. The vane 86 also includes a root (not shown) that forms a proximal end 77 of the vane 86. The root generally includes a bulbous portion and a strut connecting the bulbous portion to the airfoil 88, and the bulbous portion is connected to the strut by a neck that defines a locally minimum portion of the root cross-section.

[0108] The vane 86 is, for example, a metal vane, in other words, the vane 86 consists only of metal, or the vane 86 is a composite vane, in other words, the structure of the vane 86 consists only of composite materials. For example, the vane 49 includes a plurality of vanes 86, all of the vanes 86 being metal vanes, all being composite vanes, or some being composite vanes and others being metal vanes.

[0109] Alternatively (not shown), all of the vanes 49 are hybrid vanes 79.

[0110] Reference Figure 5 , the disk 62 includes a large upstream face 90, a large downstream face 91 ( Figure 7 ), and an edge 92 located at the outer periphery of the disk 62 that connects the large faces 90, 91 to each other. The disk 62 also has a central through-hole 94 that opens on each of the large faces 90, 91, such that the disk 62 thus has an annular shape.

[0111] The edge 92 defines the outer peripheral edge of the disk 62. The edge 92 has a circular shape, such as a cylindrical or frustoconical shape. In particular, the edge 92 partially defines an internal flow path 22.

[0112] The disk 62 also includes a plurality of bases 96, each base 96 protruding radially outward from the edge 92. The number of bases 96 is equal to the number of hybrid vanes 79. Thus, the disk 62 includes one base 96 for each hybrid vane 79.

[0113] Preferably, the base 96 and the rest of the disk 62 are an integral part. The base 96 is generally machined from the same block of material as the rest of the disk 62.

[0114] As can be seen in Figure 6 , each base 96 has a distal end 100 opposite the edge 92. The distal end 100 is substantially planar and substantially perpendicular to the radial direction of the disk 62, in other words, there is a radial direction of the disk 62 that passes through the distal end 100 and is substantially orthogonal to the distal end 100.

[0115] Each base 96 also has an outer peripheral surface 102 that defines a distal end 100. The outer peripheral surface 102 is connected to the edge 92 by a rounded corner 104. The rounded corner 104 extends around the base 96. Thus, the transition between the edge 92 and the outer peripheral surface 102 of the base 96 is gradual, which avoids aerodynamic interference.

[0116] The distal end 100 constitutes a connection surface 106 to which the root 80 (more specifically, the proximal end 77) of the corresponding hybrid blade 79 is connected (preferably, welded). Thus, the joint between the disk 62 and the hybrid blade 79 is at a distance from the edge 92, which ensures good mechanical strength. In fact, the main stress is thus concentrated in the base 96, which has greater strength since the base 96 and the rest of the disk 62 are in one piece.

[0117] The shape of the connection surface 106 is substantially the same as the shape of the proximal end 77 of the hybrid blade 79, and the hybrid blade 79 is positioned relative to the base 96 such that the belly 70, the back 72, the leading edge 74, and the trailing edge 76 of the hybrid blade 79 are flush with the outer peripheral surface 102 of the base 96, respectively. Thus, the joint between the hybrid blade 79 and the disk 62 has no sharpness, which avoids disturbing the aerodynamic flow and ensures good mechanical strength.

[0118] The disk 62 is made of metal. Advantageously, the metal has the same properties as the metal constituting the root 80 of the hybrid blade 79, in other words, the main metal element of each of the metals constituting the metal is the same as the main metal element of the other metals. Thus, the metal constituting the disk 62 and the metal constituting the root 80 of the hybrid blade 79 are, for example, the same metal or different alloys of the same base metal. This ensures good mechanical strength of the hybrid blade 79 at the disk 62.

[0119] In the example shown, the disk 62 also includes units 108 formed in the edge 92. The number of units 108 is equal to the number of blades 86. Thus, the disk 62 includes one unit 108 for each blade 86.

[0120] These units 108 have a shape that is substantially complementary to the shape of the root of the blade 86, and when the blade 86 is attached to the disk 62, these units 108 receive the root. The root of the blade 86 cooperates with the units 108 to hold the blade 86 attached to the disk 62.

[0121] The bases 96 and the units 108 are regularly distributed along the edge 92, in other words, each base 96 or unit 108 is substantially equidistant from each adjacent base 96 or unit 108.

[0122] In an alternative where all of the blades 49 are hybrid blades 79, the disk 62 does not include any units 108. Thus, the bases 96 are regularly distributed along the edge 92, in other words, each base 96 is substantially equidistant from each adjacent base 96.

[0123] Reference will now be made to Figures 9 to 15 describe a method 200 for manufacturing a rotor 46.

[0124] The method 200 begins with a first step 210 of manufacturing a rotor stage 48 of the rotor 46.

[0125] This first step 210 includes providing 212 the disk 62 and providing 213 the hybrid blades 79. These providing steps 212, 213 occur simultaneously, or, as shown, occur sequentially with respect to each other.

[0126] Providing 212 the disk 62 generally includes: machining the entire disk 62 from a single metal block.

[0127] Providing 213 the hybrid blades 79 generally includes: machining the metal parts of the root 80 and the distal portion 82 from a single metal block; manufacturing a composite structure; assembling the metal parts of the root 80 and the distal portion 82 to the composite structure.

[0128] After the providing sub-steps 212, 213 is a sub-step 214 of connecting the root 80 of each hybrid blade 79 to the disk 62. During this sub-step 214, each blade 79 is placed opposite a corresponding base 96 of the disk 62, as Figure 10 and Figure 11 shown, and then the proximal end 77 of the blade 79 is brought into contact with and welded to the connection surface 106 defined by the base 96. Advantageously, this welding is carried out by friction, for example by linear friction or orbital friction. Optionally, the welding uses a filler metal; the filler metal has the same properties as the metals constituting the blade 79 and the disk 62.

[0129] The flatness of the connection surface 106 and the proximal end 77 facilitates this welding. In fact, the plane-to-plane contact of the two surfaces 106, 77 to be welded allows friction welding, and if the surfaces have a more complex shape (for example, if one of the two surfaces has a rotational shape similar to the edge 92), friction welding will be difficult to implement. It should be noted that this surface flatness is achieved by the fact that the connection surface 106 is not directly supported by the edge 92, but by the base 96 protruding from the edge 92.

[0130] Alternatively, the proximal end 77 of the blade 79 is not welded to the connection surface 106, but brazed to the connection surface 106.

[0131] As in Figure 12and Figure 13 As can be seen in Figure 14 and Figure 15 , the welding (or brazing) of the root 80 to the base 96 forms a weld bead 110 at the outer periphery of the connection surface 106. This weld bead 110 may interfere with the aerodynamic flow at the root of the blade 79. To solve this problem, after the connection sub-step 214, there is a sub-step 215 of machining the said weld bead 110. During this sub-step 215, the weld bead 110 is machined so that the outer periphery of the joint between the blade 79 and the base 96 is smooth and regular, as can be seen in Figure 14 and Figure 15 . Figure 14 and Figure 15 As can be seen in Figure 14 and Figure 15 .

[0132] Since the weld bead 110 is spaced from the edge 92 by a distance through the base 96, it helps with the machining of the weld bead 110.

[0133] Advantageously, after all the hybrid blades 79 have been welded or brazed to the disk 62, the sub-step 215 is carried out in one go. Alternatively, the sub-step 215 is carried out in several steps after each connection of the hybrid blade 79 to the disk 62.

[0134] In appropriate cases, the manufacture 210 of the rotor stage 48 further includes: providing 216 non-hybrid blades 86; and assembling 217 the non-hybrid blades 86 to the disk 62 by inserting the roots of the non-hybrid blades 86 into the cells 108 of the disk 62.

[0135] The manufacture 210 of the rotor stage 48 ends with a sub-step 218 of keeping the said rotor stage 48 in static balance. During the sub-step 218, material is added to and / or removed from the rotor stage 48 to ensure the static balance of the rotor stage 48. This static balance of the rotor stage 48 is achieved, for example, by machining a circumferential weld bead (not shown) formed in the disk 62 or by adding weights (not shown) attached to the disk 62.

[0136] Step 210 is repeated for each rotor stage 48 of the rotor 46.

[0137] Then, during a subsequent assembly step 220, these rotor stages 48 are assembled together. In this way, the rotor 46 is obtained.

[0138] Thus, through the above embodiments, a lightweight, robust and simply manufactured low-pressure compressor rotor 46 can be obtained. Additionally, since the hybrid blade 79 includes a composite material structure, it is easier to adjust the mechanical properties of the hybrid blade 79.

[0139] It should be noted that although the above embodiments relate to a low-pressure compressor rotor, the present invention is in no way limited to this single embodiment, but is applicable to any type of turbine rotor. In particular, the above description of the rotor stage 48 and the manufacturing method 200 is applicable to fan rotors, such as the rotor 54.

Claims

1. A rotor element (48) for a turbine, the rotor element comprising a metal disk (62) and a plurality of blades (49) mounted on the disk (62), wherein, At least one of said blades (49) consists of a hybrid blade (79) made of a composite material and including a metallic root (80), said metallic root being connected to said disk (62) by welding or brazing.

2. The rotor element (48) according to claim 1, wherein, Said disk (62) has an edge (92), and for said hybrid blade or each hybrid blade (79), the disk has a connection surface (106) arranged on said edge (92), to which the root (80) of the hybrid blade (79) is connected.

3. The rotor element (48) according to claim 2, wherein, Said connection surface (106) is substantially planar.

4. The rotor element (48) according to claim 3, wherein, Said connection surface (106) is substantially perpendicular to the radial direction of said disk (62).

5. The rotor element (48) according to any one of claims 2 to 4, wherein, For said hybrid blade or each hybrid blade (79), said disk (62) includes a base (96) projecting radially outwards from said edge (92), said base (96) having a distal end (100) opposite to said edge (92), said distal end constituting the connection surface (106) for said hybrid blade (79).

6. The rotor element (48) according to claim 5, wherein, Said base (96) has an outer peripheral surface (102) defining said distal end (100), and the hybrid blade (79) is positioned relative to said base (96) such that the soffit (70), the back (72), the leading edge (74) and the trailing edge (76) of the hybrid blade are flush with the outer peripheral surface (102) of said base (96) respectively.

7. The rotor element (48) according to any one of claims 2 to 6, wherein, Said edge (92) has a rotationally symmetric shape.

8. The rotor element (48) according to any one of claims 2 to 7, wherein, Said edge (92) partially defines the air flow path (22) of said turbine (12).

9. The rotor element (48) according to any one of the preceding claims, wherein, Said rotor element constitutes a compressor rotor stage or a fan rotor stage.

10. The rotor element (48) according to any one of the preceding claims, wherein, Said root (80) has an aerodynamic profile.

11. The rotor element (48) according to any one of the preceding claims, wherein, Said root (80) is made of metal, and said hybrid blade (79) includes a distal portion (82) remote from said disk (62), said distal portion being formed at least in part by a composite structure.

12. The rotor element (48) according to claim 11, wherein, Said distal portion (82) is an extension of said root (80) at the soffit (70) and the back (72).

13. The rotor element (48) according to claim 11 or 12, wherein, Said root (80) is attached to said composite structure by means of at least one pin (81), said at least one pin being embedded in the composite structure and shrink-fitted in said root (80).

14. The rotor element (48) according to any one of claims 11 to 13, wherein, Said root (80) is attached to said composite structure by means of a metal beam (83), said metal beam being integral with said root (80) and extending in the composite material.

15. The rotor element (48) according to any one of claims 11 to 14, wherein, Said root (80) is attached to said composite structure by a bathtub fastener.

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

17. An aircraft (10), the aircraft comprising at least one turbine (12) according to claim 16.

18. A method (210) for manufacturing a rotor element (48), the rotor element being the rotor element according to any one of claims 1 to 15, the method comprising the following steps: - Provide (212) a metallic disk (62), - Provide (213) at least one hybrid blade (79) made of a composite material and including a metallic root (80), and - Connect (214) the root (80) of said hybrid blade or each hybrid blade (79) to said disk (62), said connection (214) being effected by welding or brazing.

19. The manufacturing method (210) according to claim 18, the manufacturing method comprising the following additional steps: making the rotor element (48) statically balanced (218), for example by machining the circumferential weld beads formed in the disc (62) or by adding weights attached to the disc (62).

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