FIBER PREFORM WITH STIFFENING ELEMENTS SHAPED BY UNIDIRECTIONAL WIRE LAYERS

Description

Technical Field

[0001] The present invention relates to a fibrous preform intended to reinforce a turbomachine blade and equipped within an internal volume with one or more fibrous stiffeners formed by one or more unidirectional layers of fibers, which dampen vibration modes during operation. The invention also relates to a method for manufacturing a turbomachine blade comprising forming a matrix within a porosity of the fibrous preform. Previous technique

[0002] Aircraft parts are commonly made of composite materials, comprising a fibrous reinforcement and a matrix within the porosity of this reinforcement, with the aim of reducing mass while maintaining good mechanical properties. The fibrous reinforcement can be obtained by three-dimensional weaving or by stacking layers of two-dimensional fabric. US 2013017093, in particular, describes an aircraft propeller blade with a fibrous reinforcement comprising two fibrous portions forming an aerodynamic profile and defining an internal volume between them containing a lower-density conforming insert. This solution allows for further mass reduction while maintaining satisfactory mechanical properties.The challenge of reducing mass while maintaining mechanical properties through the use of composite materials arises particularly for turbomachine blades, especially outlet guide vanes (OGVs), which are subjected to significant mechanical loads during operation. EP 2781694 discloses a blade made of composite material.

[0003] It is desirable to supply turbomachine blades made of composite material with reduced mass while exhibiting optimized mechanical strength, particularly with regard to the vibratory stresses exerted on the part in operation. Description of the invention

[0004] The invention relates to a fibrous preform of a turbomachine blade comprising two fibrous portions forming an aerodynamic profile of the blade and defining between them an internal volume, in which each fibrous portion has at least one fibrous vibration damping stiffener comprising one or more unidirectional layers of yarn which is present in the internal volume.

[0005] For the sake of brevity, the expression "fibrous vibration damping stiffener" will be referred to hereafter as "fibrous stiffener".

[0006] Vibrational stresses during turbomachine operation can induce undesirable modes within a running blade. Therefore, the figure 1 This illustrates, at four successive instants shown from top to bottom, the displacement of the fibrous portions 14 forming the aerodynamic profile of a composite material outlet guide vane under the effect of vibrations encountered during operation. The fibrous portions 14 vibrate out of phase, with a change in the distance between them. This has been identified by the inventors as potentially damaging the component and / or negatively impacting engine performance due to the geometric modification of the aerodynamic profiles linked to changes in the modal deformations within the operating range. To address this problem, the invention proposes the incorporation of fibrous stiffeners that counteract the out-of-phase vibration of the fibrous portions and their relative displacement, and couple their movements, particularly at the antinodes.The use of unidirectional layers of yarns makes it possible to provide the desired stiffness to significantly dampen problematic vibration modes during operation, while limiting disturbances in the textile definition of the fibrous portions in order to reduce the risk that the stiffeners may lead to uncontrolled textile modifications that could result in weakening of the material.

[0007] In one embodiment, said at least one fibrous stiffener extends along a radial direction of the preform.

[0008] Alternatively, at least one fibrous stiffener extends along an axial direction of the preform. It should be noted that the invention remains within the scope of the invention if the fibrous preform has one fibrous stiffener extending along the radial direction and another fibrous stiffener extending along the axial direction.

[0009] In one example of implementation, the fibrous portions are formed by three-dimensional weaving.

[0010] In particular, said at least one fibrous stiffener may be present on a debonding zone on which the unidirectional layer or layers of yarn are separated from the corresponding fibrous portion, and the yarns of this or these unidirectional layer or layers may be woven with the corresponding fibrous portion outside the debonding zone.

[0011] Such a characteristic contributes advantageously to optimizing the bending stiffness of the blade, which is particularly of interest for certain applications, such as an output guide blade.

[0012] Alternatively, at least one fibrous stiffener is added to each fibrous portion.

[0013] In one embodiment, said at least one fibrous stiffener extends through the internal volume from one fibrous portion to the other fibrous portion.

[0014] Such a characteristic helps to further improve the damping of problematic vibration modes in operation such as membrane natural modes.

[0015] In one embodiment, each fibrous portion has at least two fibrous vibration damping stiffeners, each comprising one or more unidirectional layers of yarns and which are present in distinct regions of the internal volume.

[0016] Such a feature helps to further improve the damping of problematic vibration modes during operation.

[0017] In particular, each fibrous portion may have a first fibrous stiffener extending along the radial direction and present in a region delimited by the planes of equation X1 = 0.15*C and X2 = 0.45*C, and a second fibrous stiffener extending along the radial direction and present in a region delimited by the planes of equation X3 = 0.50*C and X4 = 0.80*C, where C denotes the chord of the blade preform and the coordinates X1, X2, X3 and X4 are measured along the axial direction taking the leading edge as the origin.

[0018] Such an example constitutes a first example of a particular arrangement allowing the optimization of the damping of problematic vibration modes in operation.

[0019] In particular, each fibrous portion may have a first fibrous stiffener extending along the radial direction and present in a region delimited by the planes of equation X5 = 0.10*C and X6 = 0.40*C, a second fibrous stiffener extending along the radial direction and present in a region delimited by the planes of equation X7 = 0.35*C and X8 = 0.65*C, and a third fibrous stiffener extending along the radial direction and present in a region delimited by the planes of equation X9 = 0.60*C and X10 = 0.90*C, where C denotes the chord of the blade preform and the coordinates X5, X6, X7, X8, X9 and X10 being measured along the axial direction taking the leading edge as the origin.

[0020] Such an example constitutes a second example of a particular arrangement that optimizes the damping of problematic vibration modes during operation.

[0021] The two arrangement examples described above are applicable to various blade geometries, particularly whether they are movable or fixed. However, the first arrangement may be preferred for blades with a shorter chord length.

[0022] In one embodiment example, the fibrous preform is a preform of an outlet guide vane.

[0023] The invention also relates to a turbomachine blade comprising a fibrous preform as described above, and a matrix present in a porosity of the fibrous preform.

[0024] The blade could be an aircraft turbomachine blade.

[0025] The invention also relates to a turbomachine comprising a turbomachine blade as described above. The turbomachine may be an aircraft turbomachine.

[0026] The invention also relates to a method for manufacturing a turbomachine blade from composite material, comprising at least: the formation of a fibrous preform as described above, and the formation of a matrix in a porosity of the fibrous preform.

[0027] In one embodiment, a fugitive material conforming insert is present in the internal volume of the preform, and the process includes, after the formation of the matrix, the removal of the conforming insert.

[0028] In this case, a blade made of composite material is obtained, which is hollow, having an empty internal volume. Of course, this does not depart from the scope of the invention if the shaping insert is not intended to be removed but to remain in the blade, for example, to reduce its mass compared to a monolithic part without an internal volume. Brief description of the drawings

[0029] [ Fig. 1 ] There figure 1 represents, schematically, the out-of-phase displacements of fibrous portions of a non-inventory output guide vane for a vibratory mode in operation, at four successive instants (from top to bottom). Fig. 2 ] There figure 2 represents, schematically and partially, an example of an outlet guide vane according to the invention. Fig. 3 ] There figure 3 represents, schematically, a section of the dawn of the figure 2 according to III-III. Fig. 3A ] There figure 3A represents, schematically and partially, the extent of the fibrous stiffeners present in the blade of the figures 2 And 3 along the height. Fig. 4 ] There figure 4 represents, schematically, a section of a variant of an output guide vane according to the invention. Fig. 5 ] There figure 5 represents, schematically, a detail of the dawn of the figure 4 showing an example of the arrangement of unidirectional layers of fiber stiffener yarns relative to the fibrous portions. Fig. 6 ] There figure 6 represents, schematically, a variant of the figure 5 showing another example of the arrangement of unidirectional layers of fiber stiffener wires relative to the fibrous portions. Fig. 7 ] There figure 7 represents, schematically, a variant of the dawn of the figure 4 with fibrous stiffeners traversing the internal volume and extending from one fibrous portion to another. Fig. 8 ] There figure 8 represents, schematically and partially, a variant of the output guide vane according to the invention. Fig. 9 ] There figure 9 represents, schematically, another variant of the output guide vane according to the invention. Fig. 10 ] There figure 10 represents, schematically, an example of a flowchart of a manufacturing process for a part according to the invention. Description of the implementation methods

[0030] The invention will be described below and in the accompanying drawings with regard to its application to an outlet guide vane. The invention also applies to other types of turbomachinery blades, such as a fan blade, a compressor blade, or a turbine blade. It should be noted in particular that the invention can be applied to a static turbomachinery blade as well as to a moving turbomachinery blade, for example, a rotating one.

[0031] The following also provides a description of the invention applicable to the case of a fibrous preform formed by three-dimensional weaving, but we do not depart from the scope of the invention when the preform is obtained by other techniques, for example by stacking layers of two-dimensional fabric or unidirectional layers.

[0032] There figure 2 Figure 1 represents an example of an outlet guide blade 1 according to the invention. This blade 1 is a static turbomachine component that allows the flow exiting the fan into the secondary flow to be "straightened." The blade 1 comprises an aerodynamic profile 3, adapted to direct the flow downstream of the turbomachine, and platforms 5 and 7 that define a flow channel. The profile 3 extends between a leading edge 3a and a trailing edge 3b along an axial direction DA. The leading edge 3a is located on the upstream side and the trailing edge 3b on the downstream side. The platforms 5 and 7 are spaced along a radial direction DR. The profile 3 has, in cross-section with respect to the radial direction DR, a curved shape of variable thickness between its leading edge 3a and its trailing edge 3b, as schematically represented in Figure 1. figure 3 .

[0033] The blade 1 is made of composite material and comprises a fibrous preform 10 obtained by three-dimensional weaving, which is densified by a matrix present in its porosity. The fibrous preform 10 of the turbomachine part 1 corresponds to its fibrous reinforcement 10 and defines the aerodynamic profile. The fibrous preform 10 has the shape of the blade 1 and is formed by three-dimensional weaving of first and second yarns, as illustrated in the example. The formation of the fibrous preform 10 initially involves the formation of a fibrous blank by three-dimensional weaving using a Jacquard-type loom on which a bundle of weft yarns is arranged in a plurality of layers, the weft yarns being connected by warp yarns. By "three-dimensional weaving" or "3D weaving," we mean here a fabric in which at least some of the warp yarns connect weft yarns over several weft layers.It should be noted, however, that within the scope of the invention, the first yarns can be the warp yarns and the second yarns the weft yarns, or vice versa, the roles of warp and weft yarns being interchangeable. The fibrous preform can have an "interlock" weave, that is, a three-dimensional weave in which each layer of first yarns connects several layers of second yarns, with all yarns in the same column of first yarns having the same movement within the plane of the weave. It does not depart from the scope of the invention if another three-dimensional weave is used to form the preform. The material of the first and second yarns is chosen according to the application; the first and second yarns can be made of carbon or a ceramic material, such as silicon carbide or an oxide ceramic.

[0034] During the weaving of the fibrous rough corresponding to the illustrated preform, a debonding is made between two successive layers of second yarns on a debonding zone 12. This debonding makes it possible to define two fibrous portions 14 debonded and not woven together on the debonding zone 12 which can be separated from each other. The fibrous portions 14 define an internal volume 16, which here is in the form of a pocket 16 or a cavity 16. In particular, in the unlinking zone 12, no layer of yarns from one fibrous portion 14 is woven with a layer of yarns from another fibrous portion 14. In the unlinking zone 12, no yarn simultaneously weaves together a layer of yarns from one fibrous portion 14 with a layer of yarns from another fibrous portion 14. The fibrous portions 14 can form fibrous skins, of relatively thin thickness, for example less than or equal to 10 mm, in particular between 2 mm and 10 mm.The fibrous portions 14 are each obtained by three-dimensional weaving in the illustrated example.

[0035] The unbonding zone 12 extends from a first axial end 12a located on the leading edge 3a to a second axial end 12b located on the trailing edge 3b. The unbonding zone 12 can extend over at least 50%, for example at least 75%, of a dimension of the preform, for example its chord (which corresponds to the distance between the leading edge 3a and the trailing edge 3b). In the illustrated example, the unbonding zone 12 extends over substantially the entire height of the fibrous preform 10, but it does not depart from the scope of the invention if it extends over only a portion of this height. The weaving is continuous outside the unbonding zone 12 so as to join the fibrous portions 14 and form a single piece of fabric. The leading edge 3a is here woven continuously so as to connect the two fibrous portions 14 by weaving in a woven area 15.Such continuous weaving at the leading edge 3a is advantageous for increasing its resistance to potential impacts. The trailing edge 3b is also continuously woven here so as to connect the two fibrous portions 14 by weaving within a woven zone 11. Alternatively, the fibrous portions 14 can be connected by means other than weaving at the trailing edge 3b. In the illustrated example, radial unlinking can also be implemented to form the fibrous reinforcement of the platforms 5 to 7 from the fibrous reinforcement of the profile 3.

[0036] The internal volume 16, defined by the fibrous portions 14, is designed to accommodate the shaping insert 18 before the matrix is ​​formed. This insert 18 gives the desired shape to the fibrous preform, and in particular to the aerodynamic profile 3. In the illustrated example, the shaping insert 18 is introduced through a radial end of the fibrous preform that belongs to the debonding zone 12. Within the preform, the fibrous portions 14 enclose the shaping insert 18. The shaping insert 18 is shown here as being a single piece, but it is not outside the scope of the invention if it consists of several distinct elements. The shaping insert 18 can be produced by molding or machining from a block of material. The conformation insert 18 here has ends 18a and 18b rounded to the shape respectively of the leading edge 3a and the trailing edge 3b.The conformation insert 18 may also include a median portion 18c between the ends 18a and 18b shaped to form at least a part of the intrados and extrados of the blade 1.

[0037] According to a first example corresponding to figures 2 And 3 The blade 1, which is intended to be mounted in the turbomachine, still includes the shaping insert 18. In this case, it is advantageous to choose a shaping insert 18 with a lower density than the density of the fibrous portions 14 densified by the matrix, with a view to reducing the weight of the part. The shaping insert 18 can be made of a porous material, such as a foam or a cellular material, for example, as described in document WO 2011 / 083250. The shaping insert 18 can be rigid so as not to deform during operation. According to a second example corresponding to the figure 4 The blade intended to be mounted in the turbomachine is hollow, with the internal volume 16 being empty. In this case, the shaping insert 18 introduced to form the preform was removable or made of fugitive material and was removed after densification by the matrix.

[0038] As mentioned above, fiber stiffeners were integrated into the preform to improve the blade's resistance to vibrational stresses during operation and thus increase its lifespan. The following section describes details of these fiber stiffeners using several embodiment examples.

[0039] In this respect figures 3, 3A And 4 illustrate a first example of fibrous stiffeners that can be used within the scope of the invention. The case of figures 3 et 3A does not differ from the case of the figure 4 that the insert 18 is or is not present in the blade. In the illustrated example, each fibrous portion 14 has two fibrous stiffeners 19a-19b which are located in distinct regions of the internal volume 16. However, the invention is not limited to this precise number of stiffeners. According to one embodiment, each fibrous portion may have more than two fibrous stiffeners, regularly spaced or not, for example, at least three or four fibrous stiffeners. Alternatively, each fibrous portion has only one fibrous stiffener.

[0040] Each fibrous stiffener 19a-19b is formed of several unidirectional layers of yarns 191a-191b, that is, by several layers of yarns 191a-191b extending substantially in the same direction. The invention remains within the scope of this invention if the fibrous stiffeners comprise only a single unidirectional layer. The yarns 191a-191b are not woven or bonded together in each fibrous stiffener 19a-19b. The yarns 191a-191b are present in the internal volume 16, on an internal surface SI of the fibrous portions 14 delimiting this internal volume 16. The illustrated example concerns the case of fibrous stiffeners 19a-19b in which the yarns 191a-191b extend along the radial direction DR. An orientation of the unidirectional layers along the axial direction DA is also possible and will be discussed later in connection with the figure 9 In the example of figures 3 And 4, the fibrous stiffeners 19a-19b form protruding reliefs of the fibrous portions 14 in the internal volume 16. The wires 191a-191b are arranged in each stiffener 19a-19b so as to give the latter a variable thickness with possibly a maximum thickness positioned in a region of maximum relative displacement of the fibrous portions 14 associated with a vibrational mode of the blade in operation.

[0041] The person skilled in the art will recognize that the number and positioning of the stiffeners 19a-19b depends on the part and the vibrational stresses encountered in operation and will be able to seek a compromise between the desired damping effect and the control of the mass of the blade incorporating the stiffeners.

[0042] In order to optimize the damping, it is possible to position each fibrous stiffener 19a-19b in a region comprising an antinode of the vibrational mode and delimited by planes whose positions correspond to amplitudes of relative displacement of the fibrous portions 14 equal to 20% of an amplitude of said antinode.

[0043] According to one example, each fibrous portion 14 may have (i) a first fibrous stiffener 19a located in a region delimited by planes P1 and P2, respectively, with equations X1 = 0.15*C and X2 = 0.45*C, and (ii) a second fibrous stiffener 19b located in a region delimited by planes P3 and P4, respectively, with equations X3 = 0.50*C and X4 = 0.80*C, where C denotes the chord of the blade or blade preform, and the coordinates X1, X2, X3, and X4 are measured along the axial direction DA, taking the leading edge 3a as the origin. The majority (more than 50%), or even all, of the fibers 191a of the first fibrous stiffener 19a may be located between planes P1 and P2. The majority (more than 50%), or even all, of the 191b wires of the second fibrous stiffener 19b may be present between planes P3 and P4.

[0044] It should be noted that fibrous stiffeners present on two distinct fibrous portions 14 can be opposite each other (the first fibrous stiffeners 19a are opposite each other, as are the second fibrous stiffeners 19b in the illustrated example). These fibrous stiffeners 19a or 19b can be located on the same plane orthogonal to the axial direction DA.

[0045] The 191a-191b fibers of the unidirectional layers (or the stiffeners 19a and 19b) may each extend over at least 50% of the height H of the fibrous preform or blade, or even over at least 75% of this height. Unless otherwise specified, the blade height is measured along the radial direction DR.

[0046] According to an example and as illustrated in the figure 3A , the wires 191a-191b can extend at least from plane Q1 to plane Q2 where planes Q1 and Q2 are respectively of equation Y1 = 0.1*H and Y2 = 0.9*H, with possibly Y1 = 0.05*H and / or Y2 = 0.95*H, H denoting the height of the blade or blade preform and the coordinates Y1 and Y2 being measured along the radial direction DR taking as origin the internal radial end of the blade or blade preform.

[0047] THE figures 5 And 6 The following will now be described, showing two variants of integrating the fibrous stiffeners 19a-19b into the textile of the blade's fibrous preform. These variants apply equally regardless of the number or position of the stiffeners. In the example illustrated in the figure 5 The yarns 191a can be co-woven with the fibrous portions 14 in the woven zones 194 and separated from them in a debonding zone 190. The debonding zone 190 is located between the woven zones 194. The woven zones 194 can form a single piece of fabric. The debonding 190 is carried out here along the radial direction DR, but the same principle remains applicable to the case of stiffeners extending along the axial direction DA, as described below. Thus, in the example of the figure 5 The unbinding zone 190 is created during the weaving of the blank in order to release the yarns 191a and form the stiffener 19a. The integration of the stiffeners thus does not require an additional step after weaving. In this case, an insert 192 can be positioned to hold the yarns 191a in place during matrix formation; this insert can advantageously be made of a fugitive material so that it can be removed after matrix formation. In the variant of the figure 6 The 191a fibers are attached to the corresponding fibrous portion 14 after its formation. They can be held in place by gluing, for example, or mechanically before matrix formation. figures 5 And 6 illustrate the case of wires 191a of stiffener 19a but remain, of course, applicable regardless of the stiffener, and in particular for stiffener 19b.

[0048] The examples just described, in connection with the figures 3 à 6 , concern fibrous stiffeners 19a-19b that do not extend from one fibrous portion 14 to another. The figure 7 illustrates a variant in which the fibrous stiffeners 19c and 19d comprise unidirectional layers of yarns 191c-191d extending through the internal volume 16 from one fibrous portion 14 to the other fibrous portion 14. Those skilled in the art will prefer the use of a through stiffener such as that of the figure 7 or not traversing as illustrated in figures 3 And 4 depending on the desired compromise between vibration damping and controlling the blade mass. The use of through-split stiffeners 19c and 19d may be preferred for large blades, such as fan blades. For smaller blade geometries, non-through-split stiffeners of the type illustrated in the diagrams are suitable. figures 3 And 4 may be given preference. figure 7 illustrates a case without conformation insert 18 but, according to an unillustrated variant, a multi-part conformation insert may be present in the internal volume 16.

[0049] We have just described, in connection with the figures 3 à 7 Examples in which each fibrous portion 14 has two fibrous stiffeners. The example of the figure 8 which will now be described concerns a variant in which each fibrous portion 14 has three fibrous stiffeners.

[0050] In the example of the figure 8 , each fibrous portion 14 has three fibrous stiffeners 19e, 19f and 19g which can each have a structure as described above. According to an example, each fibrous portion 14 can have (i) a first fibrous stiffener 19e present in a region delimited by planes P5 and P6 respectively with equation X5 = 0.10*C and X6 = 0.40*C, (ii) a second fibrous stiffener 19f present in a region delimited by planes P7 and P8 respectively with equation X7 = 0.35*C and X8 = 0.65*C, and (iii) a third fibrous stiffener 19g present in a region delimited by planes P9 and P10 respectively with equation X9 = 0.60*C and X10 = 0.90*C, where C denotes the chord of the blade or the blade preform. The coordinates X5, X6, X7, X8, X9 and X10 are measured along the axial direction DA taking the leading edge 3a as the origin.

[0051] According to an example, the following characteristics can also be verified: The wires of the unidirectional layer(s) of the first stiffener 19e may extend at least from plane Q3 to plane Q4, where planes Q3 and Q4 have equations Y3 = 0.2H and Y4 = 0.7H, respectively, with Y3 possibly varying between 0.00H and 0.40H and / or Y4 varying between 0.50H and 0.90H; the wires of the unidirectional layer(s) of the second stiffener 19f may extend at least from plane Q5 to plane Q6, where planes Q5 and Q6 have equations Y5 = 0.1H and Y6 = 0.9H, respectively, with Y5 possibly varying between 0.05H and 0.1H and / or Y6 varying between 0.90H and 0.95H; and the wires of the unidirectional layer(s) of the third stiffener 19g may extending at least from plane Q7 to plane Q8 where planes Q7 and Q8 have equations Y7 = 0.3*H and Y8 = 0.7*H respectively, with Y7 possibly varying between 0.10*H and 0.50*H, and / or Y8 varying between 0.50*H and 0.90*H, where H denotes the height of the blade or blade preform and the coordinates Y3, Y4, Y5, Y6, Y7 and Y8 are measured along the radial direction DR taking as origin the inner radial end of the blade or blade preform.

[0052] We have represented on the figure 9 The case of a 19h fiber stiffener extending along the axial direction DA of the preform. Similar to what was described above, several fiber stiffeners can extend along the axial direction DA and be positioned at different heights of the blade preform.

[0053] To optimize damping, each fiber portion 14 may have at least one fiber stiffener 19h located in a region delimited by planes P5 and P6 defined above, and / or at least one fiber stiffener 19h located in a region delimited by planes P7 and P8 defined above, and / or at least one fiber stiffener 19h located in a region delimited by planes P9 and P10 defined above. Alternatively, each fiber portion 14 may have at least one fiber stiffener 19h located in a region delimited by planes P1 and P2 defined above, and / or at least one fiber stiffener 19h located in a region delimited by planes P3 and P4 defined above. The fiber stiffeners 19e may extend over more than 50% of the blade or preform chord, as illustrated.

[0054] We have just described various possible structures for blades and blade preforms according to the invention. The following section describes details of the blade manufacturing process that can be implemented within the scope of the invention.

[0055] There figure 10 is a flowchart of an example of a method according to the invention. In a first step E10, a fibrous blank of the blade is obtained. The blank can be obtained by three-dimensional weaving or by stacking layers of two-dimensional fabric or unidirectional layers. The blank includes the fibrous stiffeners described above. As indicated above, these stiffeners can be formed during the weaving of the blank by bringing out yarns 191a over a debinding zone 190 (example of the figure 5 ) or be integrated after obtaining the rough shape by attaching them to the fibrous portions (example of the figure 6The blank is then shaped by introducing the conforming insert 18 into the internal volume 16 (step E20). The fibrous portions 14 enclose the conforming insert 18 and take its shape to define the aerodynamic profile. The conforming insert 18 may be intended to be removed after the matrix is ​​formed to obtain a hollow blade, or it may be present in the blade intended to be mounted in the turbomachine. Once the blank has been shaped and the conforming insert 18 introduced, the matrix is ​​formed within its porosity by a method known per se (step E30). For example, in the case where the part is made of an organic matrix composite, the matrix formation may involve injecting a resin into the porosity of the preform and curing it, for example, cross-linking it, so as to form the matrix.The matrix is ​​formed within the porosity of the fibrous portions 14, as well as around the fibers 191a-191d of the unidirectional layers and in the inter-fiber spaces 191a-191d. The invention remains within the scope of this invention when the matrix is ​​not made of an organic material but of a ceramic material, for example, when the blade is intended to be mounted in the hot section of a turbomachine, such as a turbine blade. In this case, the ceramic matrix can be formed by chemical vapor infiltration (CVI), melt infiltration (MI), or polymer impregnation and pyrolysis (PIP). The matrix formation techniques are known per se and do not require further description.

[0056] After the die has formed, the conforming insert 18 can be removed if desired, for example by melting or dissolving it with a solvent (optional step E40), or the insert can simply be removed. If desired, machining or finishing treatments can also be carried out after the die has formed to obtain the blade ready for mounting in the turbomachine.

[0057] The expression "between ... and ..." should be understood as including the boundaries.

Claims

1. A fibrous preform (10) of a turbomachine vane or blade (1) comprising two fibrous portions (14) forming an aerodynamic profile of the vane or blade and defining therebetween an interior volume (16), wherein each fibrous portion has at least two fibrous vibration damping stiffeners (19a-19g) each comprising one or several unidirectional layers of yarns (191a-191d) and present in distinct regions of the interior volume, and wherein: - each fibrous portion has a first fibrous stiffener (19a ; 19c) extending along a radial direction (DR) and present in a region delimited by the planes (P1 ; P2) of equation X1 = 0.15*C and X2 = 0.45*C, and a second fibrous stiffener (19b ; 19d) extending along the radial direction and present in a region delimited by the planes (P3 ; P4) of equation X3 = 0.50*C and X4 = 0.80*C, where C refers to the chord of the vane or blade preform and the coordinates X1, X2, X3 and X4 being measured along the axial direction (DA) by taking the leading edge (3a) as the origin, or - each fibrous portion (14) has a first fibrous stiffener (19e) extending along the radial direction (DR) and present in a region delimited by the planes (P5 ; P6) of equation X5 = 0.10*C and X6 = 0.40*C, a second fibrous stiffener (19f) extending along the radial direction and present in a region delimited by the planes (P7 ; P8) of equation X7 = 0.35*C and X8 = 0.65*C; and a third fibrous stiffener (19g) extending along the radial direction and present in a region delimited by the planes (P9 ; P10) of equation X9 = 0.60*C and X10 = 0.90*C, where C refers to the chord of the vane or blade preform and the coordinates X5, X6, X7, X8, X9 and X10 being measured along an axial direction (DA) by taking the leading edge (3a) as the origin.

2. The fibrous preform (10) according to claim 1, wherein the fibrous portions (14) are formed by three-dimensional weaving.

3. The fibrous preform (10) according to claim 2, wherein said at least two fibrous stiffeners (19a-19h) are present on a non-interlinked area (190) on which the unidirectional layer(s) of yarns (191a-191d) are separated from the corresponding fibrous portion (14), and wherein the yarns of this or these unidirectional layers are woven with the corresponding fibrous portion outside the non-interlinked area.

4. The fibrous preform (10) according to any one of claims 1 or 2, wherein said at least two fibrous stiffeners (19a-19h) are added onto each fibrous portion (14).

5. The fibrous preform (10) according to any one of claims 1 to 4, wherein said at least two fibrous stiffeners (19c ; 19d) extend through the interior volume (16) from one fibrous portion (14) to the other fibrous portion (14).

6. The fibrous preform (10) according to any one of claims 1 to 5, wherein the fibrous preform is a preform of an outlet guide vane (1).

7. A turbomachine vane or blade comprising a fibrous preform according to any one claims 1 to 6, and a matrix present in a porosity of the fibrous preform.

8. A method for manufacturing a turbomachine vane or blade (1) made of composite material, comprising at least: - the formation of a fibrous preform (10) according to any one claims 1 to 6, and - the formation of a matrix in a porosity of the fibrous preform.

9. The method according to claim 8, wherein a conformation insert (18) made of fugitive material is present in the interior volume of the preform, and the method comprising, after the formation of the matrix, the elimination of the conformation insert.