Propeller blade with a hollow composite root

Abstract

The invention relates to a propeller blade (10) for a propeller engine made of composite material comprising a matrix densified fibrous reinforcement, the propeller blade comprising in its spanwise (D L ) direction a root (12) and an aerodynamic profile (11). The fibrous reinforcement comprises a fibrous preform with a three-dimensional weave, the fibrous preform having a root preform portion and an aerodynamic profile preform portion. The fibrous preform comprises a partition delimiting a recess forming a cavity (14) extending both into the root (12) and into the aerodynamic profile (11). A spar (130) is present in the cavity (14), the spar comprising an aerodynamic profile shaping portion (131) positioned in a first portion (142) of the cavity 14 and a root shaping portion (132) positioned in a second portion (141) of the cavity. The root (12) has a rotationally symmetrical shape.

Description

Technical Field

[0001] This invention relates to the field of propeller blades for aircraft, such as those present in turboprop engines. Background Technology

[0002] Propeller blades used in turboprop engines are typically made of metal. Although metal propeller blades have good mechanical strength, they still have the disadvantage of relatively large mass.

[0003] To obtain lighter propeller blades, it is known to produce propeller blades made of composite materials; in other words, to produce propeller blades by manufacturing structural components with matrix-densified fiber reinforcements.

[0004] Document US2013 / 0017093 describes the production of propeller blades from a fiber structure with an aerodynamic profile, a portion of which is incorporated into the propeller blade, one end of which is extended by a protrusion for forming the root of the propeller blade.

[0005] The next generation of engines requires more compact blade roots. This requirement stems from the need to enable blades to pivot about their vertical axis to adapt their angle of attack to flight conditions (for blades with variable pitch). This requirement, coupled with the fact that the blades must be coupled as low as possible to the disk, necessitates a significant reduction in the size of the root.

[0006] Therefore, compared with the roots of the prior art as described in document US2013 / 0017093, the roots of the new generation of blades have an axisymmetric or substantially axisymmetric shape and a reduced size, and these roots extend over the entire width of the lower part of the blade.

[0007] Such axisymmetric or quasi-axisymmetric shapes are more difficult to manufacture from composite materials, especially when using three-dimensional (3D) weaving to form the fiber reinforcement of the blade.

[0008] Furthermore, the mechanical loads borne by the new generation root impose additional constraints. More specifically, in addition to the mechanical traction and bending loads typically encountered (caused by centrifugal force and impact, respectively), the new generation root can be integrated into the rotor's disk using a metal housing, which generates additional mechanical loads during circumferential compression. Summary of the Invention

[0009] Therefore, it is desirable to propose a solution for producing aircraft propeller blades made of composite materials, which have a compact root and are capable of withstanding various mechanical loads.

[0010] To this end, the present invention provides a method for manufacturing propeller blades for turboprop engines, the propeller blades being made of composite materials including a matrix-densified fiber reinforcement, the method comprising:

[0011] - A fiber preform is produced as a single piece through three-dimensional weaving. The fiber preform has a flat shape extending in both a longitudinal and a transverse direction, the longitudinal direction corresponding to the spanwise direction of the propeller blade to be manufactured, and the transverse direction corresponding to the chordwise direction of the propeller blade to be manufactured. The fiber preform includes a root portion and an aerodynamic profile portion, the aerodynamic profile portion extending from the root portion along the longitudinal direction and extending along the transverse direction (D) between the leading and trailing edge portions. T )extend,

[0012] - The fiber preform is shaped to obtain a single fiber preform having an aerodynamic profile portion forming an aerodynamic profile preform and a root portion forming a root preform, and

[0013] - The fiber preform is densified by a matrix to obtain an intermediate component made of a composite material having a fiber reinforcement composed of the fiber preform and densified by the matrix, the intermediate component including an aerodynamic profile portion and a root portion.

[0014] The method is characterized in that the root portion of the fiber preform includes a separator defining a recess extending both into the root portion and into the aerodynamic profile portion of the fiber preform, the recess opening at the lower part of the fiber preform, wherein forming the fiber preform includes: inserting a spar into the recess, the spar including an aerodynamic profile forming portion positioned in the aerodynamic profile portion of the fiber preform and a root forming portion positioned in the root portion of the fiber preform to form an aerodynamic profile preform portion and a root preform portion, respectively; and the method further includes: after a densification step, machining the root portion of the intermediate component made of the composite material along a defined radius to form a propeller blade, the propeller blade including a root and an aerodynamic profile having a rotationally symmetrical shape.

[0015] Therefore, the method of the present invention can produce propeller blades with a composite material root that is both compact and perfectly suited to withstand the various mechanical loads described above. More specifically, the fiber reinforcement portion of the root is produced by 3D weaving and has a rotationally symmetrical shape that connects at its center to the fiber reinforcement portion of the aerodynamic profile. Thus, a more compact composite material root than in the prior art is obtained, which typically extends across the entire width of the lower portion of the aerodynamic profile. In this composite material root, yarns, such as warp yarns, are oriented along the spanwise direction of the blade, which, through combination with 3D weaving, imparts good mechanical strength in terms of traction and bending. Furthermore, in the root composite material, yarns, such as weft yarns, are oriented along the tangential direction of the blade, which imparts good circumferential compressive mechanical strength.

[0016] Furthermore, the rotationally symmetrical shape of the root is compatible with the integration of propeller rotation or variable pitch systems.

[0017] By inserting the spar into the fiber reinforcement, wherein the root portion is integrally formed with the aerodynamic profile portion (in other words, woven into a single piece), very good mechanical strength of the entire component is ensured, especially the forces that the root may be subjected to.

[0018] According to one aspect of the invention, the separators present in the root portion of the fiber preform and the root forming portion of the spar have a width in the transverse direction that is larger than the machining radius of the root portion of the intermediate component made of composite material. This allows the entire spar to pass through the root portion of the fiber preform.

[0019] According to another aspect of the method of the present invention, the spar is made of a composite material, including a matrix-densified fiber reinforcement, or the spar is made of a metallic material.

[0020] According to another aspect of the method of the present invention, forming the fiber preform further includes inserting a forming component made of rigid honeycomb material around the aerodynamic profile forming portion of the spar.

[0021] According to another aspect of the method of the present invention, forming the fiber preform further includes injecting an expanded material around the aerodynamic profile-shaped portion of the spar.

[0022] Another object of the present invention is a propeller blade for a turboprop engine, the propeller blade being made of a composite material comprising a matrix-densified fiber reinforcement, the propeller blade including a root and an aerodynamic profile in its spanwise direction, the root and the aerodynamic profile extending in a chordal direction between a leading edge and a trailing edge, the fiber reinforcement comprising a fiber preform having a three-dimensional weave, the fiber preform having a root preform portion and an aerodynamic profile preform portion, the root preform portion being present in the root, the aerodynamic profile preform portion being present in the aerodynamic profile, the root and the aerodynamic profile preform portions being connected to each other by the three-dimensional weave, characterized in that the fiber preform includes a separator defining a recess forming a cavity extending both into the root and into the aerodynamic profile, a sparsity being present in the cavity, the sparsity including an aerodynamic profile forming portion positioned in a first portion of the cavity and a root forming portion positioned in a second portion of the cavity, and the root of the propeller blade having a rotationally symmetrical shape.

[0023] According to one aspect of the propeller blade of the present invention, the root forming portion of the spar is exposed at the leading edge and the trailing edge of the root of the propeller blade, and the remaining portion of the root forming portion of the spar is covered by the fiber preform.

[0024] According to another aspect of the propeller blade of the present invention, the spar is made of a composite material, including a matrix-densified fiber reinforcement, or the spar is made of a metallic material.

[0025] The present invention also relates to an aircraft engine comprising a plurality of propeller blades according to the invention, and an aircraft comprising at least one such engine. Attached Figure Description

[0026] Figure 1 This is a schematic diagram illustrating the 3D weaving of fiber preforms used to manufacture blades.

[0027] Figure 2 This is an enlarged sectional view of the yarn layer assembly along the weft direction, showing... Figure 1 The formation of the separator along section II-II at the root portion of the billet.

[0028] Figure 3 It shows Figure 1 A three-dimensional schematic diagram of the forming of the root preform and the aerodynamic profile preform in the fiber preform.

[0029] Figure 4This is an exploded perspective view showing the placement of an injection tool and a fiber preform in the injection tool according to an embodiment of the present invention.

[0030] Figure 5 It shows Figure 4 A three-dimensional diagram of a closed injection tool.

[0031] Figure 6 This is a three-dimensional schematic diagram of an intermediate component made of composite material obtained according to an embodiment of the present invention.

[0032] Figure 7 In machining Figure 6 A three-dimensional schematic diagram of a blade made of composite material obtained after the root portion of the intermediate component. Detailed Implementation

[0033] This invention is generally applicable to various types of propeller blades used in aircraft engines. It has advantageous, but not exclusive, applications in large-diameter propeller blades used in oscillating or variable-pitch systems. These propeller blades typically have a so-called "cylindrical" root, meaning the root has a rotationally symmetrical shape and good resistance to traction, bending, and circumferential compressive forces. In particular, blades according to the invention can be configured as blades of enclosed impellers (such as fan blades) or open impellers (e.g., in so-called "open rotor" aircraft engines).

[0034] In the remainder of this description, exemplary embodiments are described with respect to the blades of a turboprop engine. However, these exemplary embodiments are also applicable to aircraft propeller blades.

[0035] Figure 1 The fiber preform 100 used to form the fiber preform to be produced is shown schematically.

[0036] like Figure 1 As illustrated schematically, the fiber structure prefabricated material 100 is obtained by three-dimensional (3D) weaving in a known manner using a jacquard loom, on which a bundle of warp yarns 101 or strands has been arranged into a multi-layered warp with several hundred yarns per layer, the warp yarns being connected by weft yarns 102. The fiber structure prefabricated material 100 is woven into a single piece, the prefabricated material along the longitudinal direction D... L and horizontal direction D T Extend, in the vertical direction D L Corresponding to the spanwise direction of the blade to be manufactured between the lower part 100c and the upper part 100d, the transverse direction D TCorresponding to the chordal direction between the leading edge 100a and the trailing edge 100b of the blade to be manufactured, the blank includes: an aerodynamic profile portion 111 defining two surfaces 111e and 111f for forming the suction and pressure surfaces of the blade, respectively; and a root portion 112 for subsequently forming the blade root, and in the longitudinal direction D L Extending upwards to the outside of the aerodynamic profile blank 111 and in the transverse direction D T The upper edge recedes inward from the leading edge 100a and the trailing edge 100b.

[0037] Here, the term "three-dimensional weaving" or "3D weaving" refers to a weaving method in which at least some warp yarns are connected to weft yarns on multiple weft layers, such as "interlocking weaving." Here, the term "interlocking" weaving refers to a weaving in which each layer of weft yarns connects to multiple layers of warp yarns, and all yarns in the same weft row have the same movement within the weaving plane.

[0038] Other types of known three-dimensional weaving can be used, for example, in particular those described in document WO 2006 / 136755, the contents of which are incorporated herein by reference. In particular, this document describes the production of fiber-reinforced structures for components such as blades by single-piece weaving, the fiber-reinforced structures having a first type of core weaving and a second type of skin weaving, which can impart the desired mechanical and aerodynamic properties to such components.

[0039] In particular, the fiber preform according to the invention can be woven from carbon fiber or ceramic yarn (e.g., silicon carbide).

[0040] During the weaving process of the fiber preform, its thickness and width vary, and a certain number of warp yarns are not woven, which allows for the definition of the continuously variable profile and thickness required for the preform 100. An example of progressive 3D weaving is described in document US2006 / 257260, which in particular allows the thickness of the preform to vary between a first edge used to form the leading edge and a second edge with a smaller thickness used to form the trailing edge.

[0041] According to the invention, during weaving, a separator 106 is created between two consecutive warp layers within the fiber preform 100. The separator 106 extends in a plane parallel to the surface of the fiber preform and over a region defined by a profile 106a that partially divides the fiber preform 100 into two weaving portions 113 and 114. In the longitudinal direction D... L Above, the separator 106 passes through the root portion 112 of the fiber preform 100 and partially penetrates into the aerodynamic profile portion 111 of the fiber preform 100. Furthermore, the separator 106 extends in the transverse direction D between the leading edge 100a and the trailing edge 100b of the preform 100. TExtending and receding inward from these edges, in other words, the separator 106 is not open at the leading edge 100a and trailing edge 100b to retain the connecting portions 105 and 107 adjacent to the leading edge 100a and trailing edge 100b, respectively. Furthermore, the separator 106 opens at the lower portion 100c. Thus, the separator 106 forms a recess 140 that can enter through the lower portion 100c.

[0042] Figure 2 The 3D weaving pattern of the blank 100 with interlocking weaving is schematically shown. Figure 2 It is the warp cross section of the region including the separator 106 in a part of the blank 100. Figure 1 A magnified partial view of section II-II in the figure. In this example, billet 100 includes: substantially along the longitudinal direction D L Extended 8 layers of warp yarns 101. In Figure 2 In the fiber preform 100, eight layers of warp yarns are connected by weft yarns T1 to T8 in connection areas 105 and 107, the weft yarns being substantially along the transverse direction D. T Extending further, at the separator 106, the braided portion 113 includes four layers of warp yarns 101 connected together by four weft yarns T1 to T4, while the braided portion 114, which includes four layers of weft yarns forming the yarn layer group 109, is connected by four weft yarns T5 to T8.

[0043] In other words, the fact that weft yarns T1 to T4 do not extend into the warp layer of the braided portion 114 and that weft yarns T5 to T8 do not extend into the warp layer of the braided portion 113 ensures that the separator 106 separates the braided portions 113 and 114.

[0044] Once weaving is complete, the unwoven yarns present around the fiber preform 100 are cut to extract the preform, and then the root portion of the preform is shaped. In the example described here, the root portion 112 is shaped by separating the woven portions 113 and 114 and introducing the spar 130 into the recess 140 formed by the separator 106, as shown. Figure 3 As illustrated in the illustration. The spar 130 includes an aerodynamic profile-shaped portion 131 positioned in the upper portion or base 140a of a recess 140 present in the aerodynamic profile portion 111 of the fiber preform 100. The spar 130 also includes a root-shaped portion 132 positioned in the lower portion or starting portion 140b of a recess 140 present in the root portion 112 of the fiber preform 100. The recess 140 is located at its lower portion 140b and in the transverse direction D. T The upper extension exceeds the width l 140 The width l 140 It is larger than the final diameter at the root of the propeller blade to be produced, as described below. Width l 140 With fiber preform 100 ( Figure 1 The width l of the separator 106 in ) 106 Correspondingly, this width is necessary to allow the aerodynamic profile forming portion 111 to pass through the lower portion 140b of the recess 140. The root forming portion 132 of the spar 130 is in the lateral direction D... T The upper part has an elongated shape to accommodate the width l of the recess in the lower part 140b of the recess. 140 This is especially important in order to control the shape retention of the preform in the injection tool.

[0045] The spar 130 can be made of various materials. In particular, it can be made of composite materials, including fiber reinforcements obtained by three-dimensional weaving or stacking of two-dimensional fiber layers and densified by a matrix. The spar can also be made of metallic materials.

[0046] In the example described here, a shaped part 150 made of a rigid honeycomb material (e.g., rigid foam) is positioned around the aerodynamic profile shaped portion 131 of the spar.

[0047] Therefore, as Figure 4 As shown, a fiber preform 200 was obtained, which has a longitudinal direction D L The upper part includes: an aerodynamic profile preform portion 211, and a root preform portion 212 having a raised shape and a recess 240 (including the spar 130). The aerodynamic profile preform portion 211 extends in the lateral direction D between the leading edge portion 211a and the trailing edge portion 211b. T extend.

[0048] Then, the fiber preform is densified. Densification of the fiber preform used to form the fiber reinforcement of the part to be manufactured involves filling all or part of the pores in the preform with the material constituting the matrix. This densification is carried out in a known manner (i.e., according to the liquid method (CVL)). The liquid method involves impregnating the preform with a liquid composition containing a precursor of the matrix material. The precursor is typically in polymer form, such as a high-performance epoxy resin optionally diluted in a solvent. The preform is placed in a mold that is closed in a sealed manner, the mold having a recess with the same shape as the molded final blade. The mold is then closed, and a matrix precursor liquid (e.g., resin) is injected into the entire recess to impregnate all the fiber portions of the preform.

[0049] The conversion of the precursor into the matrix (i.e., its polymerization) is carried out by heat treatment, typically by heating the mold after removing any solvents and crosslinking of the polymer. The preform is always held in a mold with a shape corresponding to the part to be produced.

[0050] In the case of forming a carbon ceramic matrix, heat treatment includes pyrolyzing the precursor to transform the matrix into a carbon or ceramic matrix, depending on the precursor used and the pyrolysis conditions. For example, the ceramic liquid precursor (especially SiC) can be a polycarbosilane (PCS), polytitanium carbosilane (PTCS), or polysilazane (PSZ) resin, while the carbon liquid precursor can be a resin with a relatively high coke content, such as a phenolic resin. To achieve the desired degree of densification, several consecutive cycles from impregnation to heat treatment can be performed.

[0051] According to one aspect of the invention, particularly in the case of forming an organic matrix, densification of the fiber preform can be achieved by a well-known resin transfer molding (RTM) method. According to this RTM method, the fiber preform is placed in a mold having the external shape of the part to be produced. A thermosetting resin is injected into the internal space of the mold containing the fiber preform. Typically, a pressure gradient is established in the internal space between the location of the resin injection and the orifice for removing the resin, in order to control and optimize the impregnation of the preform by the resin.

[0052] like Figure 4 As illustrated, the process of injecting the liquid matrix precursor composition into the fiber structure and converting it into a matrix is ​​carried out in an injection tool 300, which includes a first housing 310 and a second housing 320. The first housing 310 includes, at its center, a first impression 311 that partially corresponds to the shape and size of the blade to be produced, while the second housing includes, at its center, a second impression 321 that partially corresponds to the shape and size of the blade to be produced.

[0053] like Figure 5 As illustrated, once the tool 300 is closed, the first impression 311 and the second impression 321 of the first housing 310 and the second housing 320 together define an internal volume 301 having the shape of the blade to be produced, and in which the fiber preform 200 is placed. Compaction of the fiber preform 200 can be performed with the tool 300 closed to obtain a defined fiber content in the preform. In this case, compaction pressure is applied to the housings 310 and 320, for example, by a press. Compaction of the fiber preform can also be performed in a separate tool before the preform is introduced into the injection tool.

[0054] Tool 300 also includes means for injecting a liquid matrix precursor and means for converting the precursor into a matrix. More specifically, in the example described herein, the first housing 310 of tool 300 includes an injection port 313 for injecting the liquid matrix precursor composition into the fiber preform, while the second housing includes a discharge port 323 for interacting with a pumping system to place the tool under vacuum and remove air during injection. Injection tool 300 also includes a lower portion 340 and an upper portion 350, with the first housing 310 and the second housing 320 disposed therebetween, and the lower portion 340 and the upper portion 350 equipped with heating devices. Figure 5 (Not shown in the image).

[0055] Once tool 300 is closed, the blade is molded by impregnating the preform 200 with a thermosetting resin (polymerized through heat treatment). For this purpose, the well-known resin transfer molding (RTM) method is used. According to this RTM method, resin 360 (e.g., thermosetting resin) is injected via injection port 313 of the first housing 310 into the internal space occupied by the preform 200. Port 323 of the second housing 320 is connected to an exhaust pipe maintained under pressure. Figure 6 (Not shown in the image). This configuration allows a pressure gradient to be established between the lower part of the preform 200 (the location of resin injection) and the upper part of the preform located near port 323. In this way, the resin 360 injected primarily at the lower part of the preform will gradually impregnate the entire preform by flowing therein until excess resin is discharged through its discharge port 323. Of course, the first housing 310 and the second housing 320 of the tool 300 may each include multiple injection ports and discharge ports. The RTM method can also be performed under vacuum (VA-RTM).

[0056] For example, the resin used can be an epoxy resin with a temperature rating of 180°C. Resins suitable for the RTM method are well known. Preferably, they have low viscosity to facilitate their injection into the fibers. The choice of the resin's temperature rating and / or chemical properties depends on the thermomechanical stresses the component must withstand. Once the resin is injected into the entire reinforcement, its polymerization is carried out by heat treatment according to the RTM method.

[0057] The densification method described above enables the production of propeller blades made primarily of organic matrix (CMO), carbon matrix (C / C), and ceramic matrix (CMC) composite materials from the fiber preforms of the present invention.

[0058] After injection, polymerization, and demolding, such as Figure 6As illustrated, an intermediate component 20 made of composite material was obtained, which has a fiber reinforcement consisting of a matrix-densified fiber preform 200, the intermediate component 20 including an aerodynamic profile portion 21 and a root portion 22.

[0059] Then, from the point with a defined radius R U The intermediate part 20, made of composite material, is machined to the root portion 22, the defined radius R. U Limiting the machining profile C U This is done to form a root with a rotationally symmetrical shape. The root portion 22 is then machined to remove material present in the machining contour C. U The external material forms a root with a rotationally symmetrical shape. From Figure 6 As can be seen, a portion of the root forming part 132 of the spar 130 extends into the machined profile C. U In addition. Therefore, the machining here includes: the machining of the machining contour C. U Part of the externally densified fiber preform and also present in contour C U Part of the root forming portion 132 of the outer wing spar 130 is removed.

[0060] Finally, trim the blades to remove excess resin and machine the chamfers. Since the part is molded, no further machining is required as it conforms to the required dimensions.

[0061] like Figure 7 As illustrated, a blade 10 is obtained, which is formed from a matrix-densified fiber reinforcement. The blade includes, at its lower portion, a root 12 formed by a machined root preform portion 212 and a blade 11 formed by a blade preform portion 211 of the fiber preform 200. The blade 10 has a leading edge 11a and a trailing edge 11b, which correspond to the leading edge 211a and trailing edge 211b portions of the fiber preform 200, respectively. The root portion 12 includes a cavity 14 formed by the recess 240 of the fiber preform 200. The cavity 14 includes a spar 130 bonded to the interior of the cavity 14. An aerodynamic profile shaped portion 131 of the spar 130 is present and bonded in a first portion 142 of the cavity 14, which corresponds to the upper portion 140a of the recess 140 present in the aerodynamic profile portion 111 of the fiber preform 100. The root shaped portion 132 of the spar 130 is present and bonded in a second portion 141 of the cavity 14, which corresponds to the lower portion 140b of the recess 140 present in the root portion 112 of the fiber preform 100.

[0062] like Figure 7As can be seen, the root forming portion 132 of the sparsor 13 is exposed at the leading edge 12a and trailing edge 12b of the propeller blade root, while the remainder of the root forming portion of the sparsor is covered by a dense fiber preform. This localized exposure of the sparsor at the propeller blade root is caused by machining the root portion 22 of the intermediate component 20 to a diameter smaller than the excess width of the separator 106 at the portion of the fiber preform 100 used to form the propeller blade root. However, due to the elongated shape of the root forming portion 132 of the sparsor 130, most of the outer periphery of the root 12 is formed by the root preform portion 212 of the fiber preform 200. By retaining most of the fiber reinforcement on the outer periphery of the root 12, the strength of the propeller blade is improved in this region subjected to greater bending forces due to the aerodynamic loads on the propeller blade. Because the fiber reinforcement has a continuous three-dimensional weave from the root to the tip of the aerodynamic profile, it is well-suited for transferring localized forces to the rest of the propeller blade, thereby improving its mechanical strength.

[0063] In the example above, a shaped component made of rigid honeycomb material is positioned around the aerodynamic profile shaped portion of the spar. However, the use of such a shaped component is optional, and the spar can have a shape suitable for filling the entire volume of the recesses present in the aerodynamic profile portion. Using an additional shaped component made of rigid honeycomb material can reduce the overall mass of the propeller blade. A shaped component can also be produced in situ around the aerodynamic profile shaped portion by injecting an expanding material. In this case, a soluble filler element, such as a salt core, is temporarily positioned in the fiber preform before injection into the matrix. Once the intermediate component has been produced, in other words, after the fiber preform has densified, the filler element is removed, and the expanding material is injected into the vacated volume.

Claims

1. A method for manufacturing a propeller blade (10) for use in a turboprop engine, the propeller blade being made of a composite material comprising a matrix-densified fiber reinforcement, the method comprising: A single piece of fiber preform (100) is produced by three-dimensional weaving, the fiber preform having respectively along the longitudinal direction (D L ) and lateral direction (D) T The fiber preform (100) has an extended flat shape, the longitudinal direction of which corresponds to the spanwise direction between the lower (100c) and upper (100d) portions of the propeller blade to be manufactured, and the transverse direction of which corresponds to the chordwise direction of the propeller blade to be manufactured. The fiber preform (100) includes a root portion (112) and an aerodynamic profile portion (111), the aerodynamic profile portion extending from the root portion along the longitudinal direction (D). L ) extends, and between the leading edge portion (100a) and the trailing edge portion (100b) along the said lateral direction (D T )extend, The fiber preform (100) is formed to obtain a single fiber preform (200), the fiber preform having an aerodynamic profile portion (111) forming an aerodynamic profile preform portion (211) and a root portion (112) forming a root preform portion (212), and The fiber preform (200) is densified by a matrix to obtain an intermediate component (20) made of composite material, the intermediate component having a fiber reinforcement formed by the fiber preform (200) and densified by the matrix, the intermediate component including an aerodynamic profile portion (21) and a root portion (22). The method is characterized in that the root portion (112) of the fiber preform (100) includes a separator (106) defining a recess (140) extending both into the root portion of the fiber preform and into the aerodynamic profile portion, the recess (140) opening at the lower portion (100c) of the fiber preform. Forming the fiber preform (100) includes: inserting a spar (130) into the recess (140), the spar including an aerodynamic profile forming portion (131) positioned in the aerodynamic profile portion (111) of the fiber preform (100) and a root forming portion (132) positioned in the root portion (112) of the fiber preform, to form an aerodynamic profile preform portion (211) and a root preform portion (212), respectively; and the method further includes: after the densification step, along a defined radius (R... U The root portion (22) of the intermediate component (20) made of composite material is machined to form a propeller blade (10), the propeller blade including a root (12) having a rotationally symmetric shape and an aerodynamic profile (11).

2. The method according to claim 1, wherein, The separator (106) present in the root portion (112) of the fiber preform (100) and the root forming portion (132) of the wing beam (130) in the transverse direction (D) T The machining radius (R) of the intermediate part (20) made of composite material is greater than that of the root portion (22). U Large width (L) 106 ).

3. The method according to claim 1 or 2, wherein, The spar (130) is made of composite material, including a matrix-densified fiber reinforcement, or the spar (130) is made of metallic material.

4. The method according to claim 1, wherein, The process of forming the fiber preform (100) further includes inserting a forming component (150) made of rigid honeycomb material around the aerodynamic profile forming portion (131) of the spar (130).

5. The method according to claim 1, wherein, Shaping the fiber preform (100) further includes injecting an expanded material around the aerodynamic profile shaping portion (131) of the spar (130).

6. A propeller blade (10) for a turboprop engine, said propeller blade being made of a composite material comprising a matrix-densified fiber reinforcement, said propeller blade being disposed in its spanwise direction (D L The structure includes a root (12) and an aerodynamic profile (11), the root and the aerodynamic profile being located in a chordal direction (D) between the leading edge (11a) and the trailing edge (11b). T The fiber reinforcement comprises a fiber preform (200) having a three-dimensional weave, the fiber preform having a root preform portion (212) and an aerodynamic profile preform portion (211), the root preform portion (212) being present in the root (12), the aerodynamic profile preform portion (211) being present in the aerodynamic profile (11), the root and aerodynamic profile preform portions being connected to each other by the three-dimensional weave, characterized in that... The fiber preform includes a separator defining a recess (140) opening at the root preform portion (212), the recess forming a cavity (14) extending both into the root (12) and into the aerodynamic profile (11), a sparsity (130) existing in the cavity (14), the sparsity including an aerodynamic profile forming portion (131) positioned in a first portion (142) of the cavity (14) and a root forming portion (132) positioned in a second portion (141) of the cavity, and the root (12) of the propeller blade has a rotationally symmetrical shape.

7. The propeller blade according to claim 6, wherein, The root forming portion (132) of the spar (130) is exposed at the leading edge (12a) and trailing edge (12b) of the root (12) of the propeller blade, and the remainder of the root forming portion of the spar is covered by the fiber preform.

8. The propeller blade (10) according to claim 6 or 7, wherein, The spar (130) is made of composite material, including a matrix-densified fiber reinforcement, or the spar (130) is made of metallic material.

9. An aircraft engine, the aircraft engine comprising: Multiple propeller blades according to any one of claims 6 to 8.

10. An aircraft, the aircraft comprising: At least one aircraft engine according to claim 9.

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