Manufacturing process of a turbomachine blade

Abstract

Method for manufacturing a turbomachine blade The invention relates to a method for manufacturing a blade (100) comprising: - draping at least one surface of the internal portion of a spar (10) with fibers of a first nature intended to form a connecting skin (50), the draping being carried out by an automatic draping method; - draping a surface (201) of a mold (200) with fibers (31) of a second nature to form an aeronautical skin (30), the draping being carried out by an automatic draping method; - arranging a first structural portion (21); - arranging the draped spar (10) on the first structural portion; - arranging a second structural portion (22) on the draped spar; - closing the mold; - possibly a step of injecting a resin; and - applying to the mold conditions allowing the formation of a blade made of composite material. Figure for abstract: Fig. 2.

Description

Title of the invention: Method for manufacturing a turbine blade Technical field

[0001] This disclosure relates to the field of manufacturing propeller blades or vanes for aeronautical turbomachines. Prior art

[0002] Propeller blades or vanes for aeronautical turbomachines are generally made of metallic material. Although propeller blades or vanes made of metallic material have good mechanical strength, they nevertheless have the disadvantage of having a relatively high mass.

[0003] In order to obtain lighter propeller blades or vanes, it is known to produce propeller blades from composite material, that is to say by producing structural parts with fibrous reinforcement densified by a matrix.

[0004] In particular, it has been proposed to integrate metal feet into fiber preforms of the blade.

[0005] This ensures on the one hand good mechanical resistance of the root of the blade which is the place where the highest mechanical resistance is desirable and on the other hand a lightening of the external portion of the blade which ensures an aerodynamic role but for which a high mechanical resistance is not necessary.

[0006] However, this complicates the method of preparing the blade and requires in particular specific and complex weavings allowing the insertion of the spar part into the preform of the external part of the blade.

[0007] Furthermore, the step of inserting the spar portion into the preform is difficult to automate, so the resulting blades are expensive and complex to manufacture.

[0008] It remains desirable to further improve such blade geometries, and in particular to develop alternative methods for manufacturing variable-pitch blades, while ensuring that the latter nevertheless allow the blades to be made lighter than the metal blades of the prior art. Statement of the invention

[0009] This presentation aims to propose solutions to at least one of the problems set out above.

[0010] To this end, it proposes a method for manufacturing a turbomachine blade comprising: - draping of at least one surface of the internal portion of a spar by fibers of a first nature intended to form a connecting skin, the spar comprising a metallic blade root part and an internal part at least partially metallic, the draping being carried out by an automatic draping method - the draping of a surface of a mold by fibers of a second nature to form an aeronautical skin, the draping being carried out by an automatic draping method; - the arrangement of a first portion of structure on the fibers of a second nature arranged on the surface of the mold; - the arrangement of the draped spar on the first portion of the structure; - the arrangement of a second portion of structure on the draped spar; - closing the mold; - possibly a resin injection step; and - the application to the closed mold of temperature and pressure conditions allowing the formation of a blade in composite material.

[0011] Such a blade makes it possible to have a lower mass compared to a blade made entirely of metal of the prior art. In addition, it is obtained by simplified processes compared to composite blades of the prior art because the draping steps are carried out by automatic draping methods.

[0012] It is to the credit of the inventors to have succeeded in proposing methods for manufacturing blades made of composite material which are compatible with draping by automatic placement of fibres.

[0013] In fact, it is thus proposed processes that are more easily industrialized and represent a lower cost compared to the processes of the prior art in which manual removal is favored, in particular to guarantee the correct insertion of the spar in the opening provided for this purpose in the preform of the blade.

[0014] It will be noted that in the present application a draping method will be said to be automatic since it does not include a manual draping step.

[0015] For example, such automatic draping steps can be carried out by automatic fiber placement methods (or "AFP" for the acronym in English "Automated Fiber Placement"), automated tape placement (or "ATL" for the acronym in English "Automated Tape Layup"), filament winding (or "FW" for the acronym in English "Filament Winding") or even a so-called "pick and place" method.

[0016] In one embodiment, the draping steps are performed by an automatic fiber placement method.

[0017] This method ensures excellent fiber placement while increasing production throughput compared to prior art methods.

[0018] By "the application of temperature and pressure conditions allowing the formation of a blade in composite material", it is understood that the person skilled in the art knows, for a given resin and for a given nature of fibers, the temperature and pressure conditions which should be applied to the assembly to obtain the desired matrix binding the fibers together and thus forming the desired composite material.

[0019] In one embodiment, the fibers of the first and second nature can be pre-impregnated. In such a case, the resin injection step is not necessary, since the resin is already present during the draping of the spar and respectively of the aeronautical skin.

[0020] Alternatively, the fibers of the first nature and the second nature may be dry and the resin chosen for the resin injection step may be chosen from thermoplastic or thermosetting resins.

[0021] In one embodiment, the spar is entirely metallic, that is to say that the foot portion and the internal portion are both metallic.

[0022] In one embodiment, the spar comprises a metal blade root portion and an internal portion composed of a metal in the extension of the root and a composite material for the remainder of the internal portion of the spar.

[0023] In other embodiments, the spar comprises a metal blade root portion and an internal portion composed of a metal in the extension of the root and foam for the remainder of the internal portion of the spar.

[0024] Such spars allow a weight saving compared to a fully metallic spar, but in return complicate the production of such a spar because this adds an interface between the metal and a non-metallic material.

[0025] In one embodiment, the structural portion covers the entire surface of the internal part of the spar and is separated from the latter by the connecting skin.

[0026] In one embodiment, the fibers of a first nature may be carbon fibers.

[0027] In one embodiment, the fibers of a second nature may be chosen from glass, carbon, aramid fibers or a mixture of at least two of these elements.

[0028] In one embodiment, the structural portions may be selected from foam, wood, or a lightweight structure obtained by additive manufacturing.

[0029] In one embodiment, the bonding skin may be composed of carbon fiber composite material, glass, aramid or a mixture of at least two of these elements with a thermosetting or thermoplastic resin.

[0030] In one embodiment, the aeronautical skin may be composed of fiberglass, carbon or aramid composite material or a mixture of at least two. of these elements with a thermosetting or thermoplastic resin.

[0031] In a described method, the steps of injecting a resin and applying temperature and pressure conditions allowing the formation of a blade made of composite material may be in accordance with a resin injection molding method known as such.

[0032] Such a process (also called “RTM” for the English acronym “Resin Transfer Molding”) makes it possible to ensure excellent impregnation of the porosity of the fibers by the resin in order to ultimately obtain very reduced porosity.

[0033] For example, and in the case where the fibers are pre-impregnated fibers, the steps of injecting a resin and applying temperature and pressure conditions allowing the formation of a blade made of composite material may be in accordance with a resin injection molding process with a resin supply compatible with that used for the pre-impregnated material.

[0034] In such a process, although the fibers are pre-impregnated, resin, compatible with that contained in the pre-impregnated fibers, is injected into the mold.

[0035] Such a process, also known as SQ-RTM for the English acronym "Same Quality Resin Transfer Molding", makes it possible to apply pressure in the mold using the injected resin and thus guarantee better performance, particularly in terms of geometric strength and material health.

[0036] In one embodiment, the temperature of such a process may be between 20 and 450°C. In particular, the temperature may be variable during the process or even be chosen according to the resin used.

[0037] In one embodiment, the pressure applied during such a method may be between 1 bar and 50 bars.

[0038] In one embodiment, the duration of the step of applying the conditions capable of forming a composite material may be between a few minutes and several hours, for example 15 hours.

[0039] In one embodiment, the deposition angle for at least one automatic draping step may be between 10° and 20° between the deposition head and the substrate.

[0040] For example, the deposition angle for all automatic layup steps may be between 10° and 20° between the deposition head and the substrate.

[0041] Such an angle, called in the field "tilt angle" from the name of this parameter in English, makes it possible to ensure that the use of an automatic draping method does not damage the substrates and in particular the spar or the structural portions.

[0042] Indeed, when the deposition head has too small an angle with the substrate, it may happen that it damages said substrate.

[0043] Furthermore, such an angle ensures that it is possible to drape shapes even complex.

[0044] In one embodiment, the tension of the fibers deposited during the automatic draping steps may be non-zero and less than 20 daN.

[0045] This tension ensures excellent compaction of all the fibers deposited on the substrate and more generally of the turbomachine blade finally obtained. For example, good compaction can be defined by a small expansion, expansion being defined in the usual sense of this term in the field, namely the swelling of the fiber preform after their deposition. For example, the tension of the fibers allows an expansion of less than 8%, or even less than 7%, which ensures easy positioning of the fibers.

[0046] Also, such tension makes it possible to reduce the compaction pressure required when arranging the wires, and thus to reduce the risk of damaging the substrate.

[0047] In one embodiment, the blade does not comprise any layers other than the layers described above, i.e. the spar, the connecting skin, the structural portion and the aeronautical skin.

[0048] In one embodiment, the turbomachine blade is a variable-pitch blade, or a propeller blade.

[0049] Indeed, it is for these blades that the advantages of the invention are most useful.

[0050] In one embodiment, the method may comprise the following steps: - draping of at least one surface of the internal part of a spar by fibers of a first nature intended to form a connecting skin, the spar comprising a metallic blade root part and an internal part at least partially metallic, the draping being carried out by an automatic draping method; - draping a first surface of a mold with fibers of a second nature to form an aeronautical skin; - the draping of a second surface of a mold by fibers of a second nature to form an aeronautical skin, the draping being carried out by an automatic draping method; - the arrangement of a first portion of structure on the aeronautical skin arranged on the first surface of the mold; - the arrangement of the draped spar on the first portion of the structure; - the arrangement of a second portion of structure on the draped spar; - closing the mold by placing the second surface of the mold, draped by fibers of a second nature on the second portion of structure; - possibly a resin injection step; - the application of temperature and pressure conditions allowing the formation of a blade in composite material.

[0051] This method makes it possible to obtain the two surfaces of a turbomachine blade independently, each of the surfaces of the mold being draped independently of one another.

[0052] Furthermore, the draping steps are greatly simplified, because they are carried out independently of each other, each on a different surface and all carried out by an automatic draping method.

[0053] In one embodiment, the method of the invention comprises the following steps: - draping of at least one surface of the internal part of a spar by fibers of a first nature intended to form a connecting skin, the spar comprising a metallic blade root part and an internal part at least partially metallic, the draping being carried out by an automatic draping method; - the draping of a first surface of a mold by fibers of a second nature to form an aeronautical skin, the draping being carried out by an automatic draping method; - the arrangement of a first portion of structure on the aeronautical skin arranged on the first surface of the mold; - the arrangement of the draped spar on the first portion of the structure; - the arrangement of a second portion of structure on the draped spar, the second portion of structure; - draping the external surface of the second portion of structure using fibers of a second nature to form an aeronautical skin, the draping being carried out using an automatic draping method; - the provision of a retaining plate to ensure that the draped aeronautical skin is kept in shape; - possibly a step of injecting a resin into the load thus obtained; - the application of temperature and pressure conditions allowing the formation of a blade in composite material.

[0054] In this embodiment, the retaining plate ensures the sealing of the assembly, as well as maintaining the shape of the last draped aeronautical skin.

[0055] This allows, unlike a method comprising a mold closing step, obtaining a part with an even better surface condition than when the part is previously draped over a portion of the mold.

[0056] The best surface condition is important because it makes it possible to improve the aerodynamic properties of the blade thus obtained.

[0057] In one embodiment, the surfaces of the internal part of the spar and of the structural portion in contact with the connecting skin have a regular undulation.

[0058] Such a regular undulation is understood in the direction of the thickness of the blade, that is to say from the spar towards the outside and this in a manner normal to the external surface of the blade.

[0059] Such a blade makes it possible on the one hand to have a significantly reduced mass compared to a blade made entirely of metal of the prior art, and on the other hand to ensure excellent continuity of forces between the blade root part and the rest of the blade, thanks to the undulations of the connecting skin.

[0060] Indeed, it is to the credit of the inventors to have managed to identify that the automatic draping methods allow excellent filling of the corrugations, while ensuring that the blade has mechanical characteristics superior to those of the blades of the prior art.

[0061] The undulation of the surface of the blade root and of the surface of the facing structural portion makes it possible to significantly improve the transfer of forces between the blade root portion and the structural portion.

[0062] Furthermore, unlike a flat surface, the corrugation allows mechanical retention between the metallic and composite components allowing the transfer of forces not simply through a flat shear interface but through a shape effect which allows the transfer of part of the force by matting.

[0063] In one embodiment, the undulation of the surface of the inner portion of the spar and the undulation of the surface of the structural portion may comprise a period of between 20 mm and 70 mm.

[0064] In one embodiment, the undulation of the surface of the inner portion of the spar and the undulation of the surface of the structural portion may comprise an amplitude of between 2.0 mm and 8.0 mm.

[0065] The “amplitude” and the “period” of the surface of the internal part of the spar and the undulation of the surface of the portion of structure here takes the usual meaning of these terms in the field of waves.

[0066] The amplitude characterizes the distance between a peak and a trough, and the period the distance between two successive peaks (or two successive troughs).

[0067] These parameters proposed for the amplitude and period of the surface undulation of the blade root portion form an excellent compromise between the ease of draping the surface of the blade root by means of an automatic draping method and excellent transmission of forces between the blade root and the rest of the blade.

[0068] Preferably, the undulation of the surface of the internal part of the spar and the undulation of the surface of the structural portion have the same period and the same amplitude.

[0069] Indeed, having complementarity between the undulations makes it possible to ensure excellent continuity of efforts.

[0070] Regular undulation is understood in the direction of the thickness of the layers concerned, that is to say in the direction which goes from the spar towards the outside of the blade. In other words, the regular undulation causes a variation in the thickness of the spar and of the structural portion.

[0071] Nevertheless, the corrugation does not impair the achievement of a smooth outer skin because the corrugations are filled by the arranged aeronautical skin. For example, the surfaces of the inner portion of the spar and the aeronautical skin in contact with the connecting skin are preferably corrugated and specifically oriented folds are interposed in the hollows of the corrugations to enable a smooth outer surface to be achieved for the aeronautical skins.

[0072] By "regular" designating the undulation, it is understood that the period and the amplitude of the undulation do not vary by more than 25% for a blade.

[0073] In one embodiment, a portion of the surface of the internal part of the spar is not covered by the structural portion and the connecting skin is then in direct contact with the internal part of the spar on the one hand and the aeronautical skin on the other hand.

[0074] In particular, the portion of the spar forming the intrados and / or the extrados of the turbomachine blade can be in direct contact with the connecting skin, then the aeronautical skin.

[0075] Indeed, in certain embodiments, the presence of the structural portion is not necessary for all zones of the blade.

[0076] The intrados and the extrados of the turbomachine blade play mainly an aerodynamic role, and serve to guide the air flow passing through the turbomachine. Thus, it may be advantageous not to place any structural portion there in order to further reduce the weight of the blade by avoiding the presence of structural portions where they are not strictly necessary.

[0077] In another embodiment, the structure portion is present between all the surfaces in contact with the connecting skin, and the structure portion further comprises the undulations which have just been described.

[0078] This embodiment ensures that the entire blade obtains the advantages described for the presence of the undulations, i.e. excellent transmission of forces. Brief description of the drawings

[0079] [Fig.l] [Fig.l] is a schematic representation of a turbomachine.

[0080] [Fig.2] [Fig.2] is a schematic representation of a turbomachine blade obtained by a process described above.

[0081] [Fig.3] [Fig.3] is a schematic representation of a turbomachine blade obtained by a process described above.

[0082] [Fig.4] [Fig.4] is a schematic representation of a turbomachine blade also obtained by a process described above but different from the dawn of [Fig.3].

[0083] [Fig.5] [Fig.5] is a schematic representation of a characteristic by particular of the blades obtained in a particular embodiment of the method described above.

[0084] [Fig.6] [Fig.6] is a schematic representation of a step of a process according to an embodiment.

[0085] [Fig.7] [Fig.7] is a schematic representation of a step of a process according to an embodiment.

[0086] [Fig.8] [Fig.8] is a schematic representation of a step of a method according to an embodiment.

[0087] [Fig.9] [Fig.9] is a schematic representation of a step of a process according to an embodiment.

[0088] [Fig. 10] [Fig. 10] is a schematic representation of a step of a method according to one embodiment. Description of the embodiments

[0089] The invention is now described by means of figures, present for descriptive purposes to illustrate certain embodiments of the invention and which should not be interpreted as limiting the latter.

[0090] The invention is now described by means of figures, present for descriptive purposes to illustrate certain embodiments of the invention and which should not be interpreted as limiting the latter.

[0091] The invention applies generally to obtaining different types of propeller blades or vanes used in aircraft engines.

[0092] The invention finds an advantageous but not exclusive application in methods for obtaining large propeller blades or vanes which are intended to be integrated into pivoting or variable pitch systems. Such propeller blades or vanes are generally provided with a so-called "cylindrical" root, that is to say having a shape of revolution, and good resistance to tensile, bending and circumferential compression forces.

[0093] Such a blade may in particular constitute a blade for shrouded moving wheels such as fan blades or a blade for unshrouded moving wheels such as in so-called “open rotor” aeronautical engines, and preferably a moving blade, also called variable pitch.

[0094] In the remainder of the description, the exemplary embodiments are described in relation to blades for turbomachines. However, the exemplary embodiments apply also to propeller blades for aircraft.

[0095] [Fig. 1] represents, in section along a vertical plane passing through its main axis A, a double-flow turbojet 1. It comprises from upstream to downstream according to the circulation of the air flow, a fan 2, a low-pressure compressor 3, a high-pressure compressor 4, a combustion chamber 5, a high-pressure turbine 6, and a low-pressure turbine 7.

[0096] In the present application, the relative terms of positioning, for example “upstream”, “downstream”, “internal” and “external”, will be understood in relation to the horizontal axis A of the casing defining the axial direction, traveled in the direction of flow of the main and secondary air flows of the turbomachine.

[0097] Thus, a so-called “upstream” element will be crossed before a so-called “downstream” element and a so-called “internal” element will be closer to axis A than an “external” element.

[0098] [Fig.2] illustrates a blade 100 obtained by a method described previously.

[0099] Such a blade 100 comprises a spar 10 comprising a metal blade root portion 13 and an internal portion 11, 12 at least partially metallic.

[0100] In the embodiment shown, part 11 is metallic and is located in the extension of the foot 13, while part 12 is made of a composite material.

[0101] In this way the weight of the spar 10 is reduced, but the mechanical resistance characteristics necessary for the proper functioning of the spar 10 are ensured by the foot 13 and the metal part 12 of the internal part.

[0102] The blade 100 of [Fig.2] further comprises a structural portion 20 arranged on the internal part of the spar.

[0103] In one embodiment, such a structural portion 20 may be made of a material chosen from carbon, glass or aramid fiber composite material or a mixture of two or more of these compounds with a thermosetting or thermoplastic resin.

[0104] The structural portion 20 plays, as its name indicates, a structuring role for the blade 100. It is in fact on the structural portion 20 that the aeronautical skin 30 is arranged, which only plays an aerodynamic role.

[0105] The structural portion 20 must allow, when the root 13 of the blade 100 is moved, the entire blade 100 and in particular the aeronautical skin 30 to take the desired position.

[0106] The blade 100 is preferably a variable-pitch blade. Indeed, the latter are subject to greater constraints in operation and this comes in particular from the fact that it is necessary to be able to reposition them if necessary and that it is desirable that the position of the blades and in particular of their external surfaces is then as precise as possible. Thus, for variable-pitch blades it is necessary that a movement of the root 13 is transmitted as precisely as possible to the rest of dawn 100 and in particular to the aeronautical skin 30 of the dawn.

[0107] As illustrated in [Fig.2], the blade comprises a portion of aeronautical skin 30, arranged at least partially on the structural portion 20, said aeronautical skin 30 forming an external surface of the blade 100.

[0108] By “forming an external surface of the blade” it is understood that the aeronautical skin 30 is in contact with the air encountered by the blade 100.

[0109] [Fig.2] is thus a section of a blade 100, and the portions of spar 10, structure 20 and even the connecting skin 50 are covered on either side by the aeronautical skin 30, which alone forms the surface of the intrados 103 and the extrados 104 of the blade 100.

[0110] This will be better understood from the view of a blade similar to that shown in [Fig.3],

[0111] It will be noted that the leading edge 101 and the trailing edge 102 are identified in [Fig.2].

[0112] As can be seen in [Fig. 3], the blade 100 is characterized in that it comprises a connecting skin 50, arranged between the surface of the internal part 11, 12 of the spar 10 and a surface of the structural portion 20.

[0113] Where appropriate, a regular undulation may be present between the spar 10 and the structural portion 20.

[0114] For the purposes of understanding, the regular undulation of the connecting skin and the surfaces in contact with it is not shown in [Fig.2] and will be presented more schematically in connection with [Fig.5].

[0115] [Fig.3] and [Fig.4] illustrate another view of two blades 100 obtained by two distinct embodiments.

[0116] In [Fig. 3] is shown a blade 100 which comprises a portion of structure 20 around the entire surface of the internal part 11, 12 of the spar 10.

[0117] In the embodiment of [Fig.3], the connecting skin 50 is then present over the entire surface of the internal part 11, 12 of the spar 10, between the spar 10 and the structural portion 20.

[0118] This embodiment ensures excellent continuity of forces between the spar 10 and the rest of the blade 100.

[0119] In such an embodiment and as illustrated in [Fig. 3], the structural portion 20 is then itself completely covered with the aeronautical skin 30.

[0120] [Fig.3] also marks, by rectangle V, the view adopted for [Fig.5].

[0121] [Fig.4] illustrates an embodiment in which, unlike the embodiment of lisation shown in [Fig.3], the structural portion 20 is not arranged around the entire surface of the internal portion 11, 12 of the spar 10.

[0122] As illustrated, in this embodiment, the structural portion 20 is only present for the portions of the leading edge 101 and the trailing edge 102 of the blade.

[0123] More precisely, for the faces of the spar 10 facing the intrados 103 and the extrados 104 of the blade 100, the connecting skin 50 is arranged directly between the internal part 11, 12 of the spar 10 and the aeronautical skin 30.

[0124] In this embodiment, the absence of the structural portion 20 facing the aeronautical skin 30 intended to form the intrados 103 and the extrados 104 allows a weight saving, while ensuring that the presence of the connecting skin, then directly arranged between the internal part 11, 12 of the spar 10 and the aeronautical skin 30 nevertheless allows acceptable continuity of forces.

[0125] In such a case, the surface of the internal part 11, 12 of the spar 10 and that of the aeronautical part 30 directly in contact with the connecting skin 50 may have a corrugated surface as described.

[0126] Furthermore, the connecting skin 50 directly arranged between the surface of the internal part 11, 12 of the spar 10 and the surface of the structural portion 20, may have a corrugation which allows the technical effects already described to be obtained.

[0127] In one embodiment, the thickness of the connecting skin 50 may be between 0.8 mm and 10 mm.

[0128] The connecting skin does not in fact need to be particularly thick provided that it allows good consistency between the two portions that it connects.

[0129] The connecting skin 50 in fact allows excellent adhesion between the spar 10 and the structural portion 20, or where appropriate between the spar 10 and the aeronautical skin portion 30.

[0130] This excellent adhesion gives the blade 100 mechanical properties at least as good as those of the prior art while also ensuring better continuity of forces between the spar 10 and the rest of the blade 100.

[0131] [Fig.5] illustrates more precisely what is meant by the regular undulations of the connecting skin 50, and also of the surfaces of the internal part 11, 12 of the spar 10 and of the connecting portion 20 in direct contact with the connecting skin.

[0132] The view shown in [Fig.5] is marked by frame V in [Fig.3].

[0133] As has been described, the connecting skin 50 between the internal part of the spar 10 and the structural portion 20 is arranged on surfaces having regular undulations. Such undulations allow excellent continuity of forces between the spar 10 and the structural portion 20.

[0134] The surface undulations preferably have a period T of between 20 mm and 70 mm.

[0135] In one embodiment, the surface undulations have an amplitude A of between 2.0 mm and 8.0 mm.

[0136] As described, these parameters allow for the use of automatic fiber draping methods, while at the same time ensuring that the undulations allow to obtain the expected effects.

[0137] In one embodiment, the surface undulation is preferably present only for the surfaces directly in contact with the connecting skin.

[0138] The thickness of the structural portion 20 must be sufficient to allow the external surface of the structural portion 20, i.e. its surface which is not in contact with the connecting skin, to be smooth.

[0139] Indeed, for aerodynamic constraints, it is important that the external surface of the blade 100 is smooth.

[0140] By "smooth", it is understood that the external surface of the blade is smoother than the surface undulations. However, it is not excluded that this surface nevertheless has a certain roughness, for example since it comprises fibers woven according to a particular arrangement. Such roughness must nevertheless be compatible with an application for an aeronautical blade and be in particular satisfactory knowing the aerodynamic constraints that such an application imposes.

[0141] In one embodiment, the thickness of the spar is greater than or equal to 1.0 mm, for example between 1.0 mm and 50 mm.

[0142] In one embodiment, the thickness of the structure portion is greater than or equal to 2.0 mm, for example between 2.0 mm and 50 mm.

[0143] In one embodiment, the thickness of the aeronautical skin is greater than or equal to 0.8 mm, for example between 0.8 mm and 10 mm.

[0144] It will be noted if clarification were necessary that the undulations of the two surfaces of the connecting skin 50 in [Fig.5] are similar and complementary.

[0145] Finally, Figures 6 to 10 illustrate successive steps of a manufacturing method according to one embodiment of the invention.

[0146] The method described is not the only method according to the invention, and other alternatives are possible whether for a given step or for a set of steps.

[0147] [Fig.6] illustrates a first step of a process in which the internal surface 201 of a mold 200 has been draped with fibers making it possible to form the first portion 31 of an aeronautical skin 30.

[0148] This draping is carried out by an automated draping process. This results in a more reproducible and less expensive process than manual draping.

[0149] It is the internal surface 201 of the mold 200 which allows the shape of the aeronautical skin to be obtained, which makes it possible to increase the reproducibility of the process.

[0150] We then move on to the step represented by [Fig.7], where a first portion 21 of the structural portion 20, also called first structural portion 21, is deposited on the internal surface 31s of the first portion 31 of the aeronautical skin 30.

[0151] The first portion of structure 21 here has a surface 21s which has surface undulations without this being strictly necessary.

[0152] These undulations are represented in the figures in an extremely schematic manner and these figures should not be interpreted in a limiting manner.

[0153] Once the arrangement of the first portion of structure 21 has been completed, the spar 10, the surface of which is covered with the connecting skin 50, can be arranged.

[0154] Preferably, the spar 10 draped by the connecting skin 50 can be obtained by draping a corrugated surface of the internal part of a spar 10 by fibers of a first nature, the spar comprising a metal blade root part and an internal part at least partially metallic.

[0155] In the section of [Fig.8] and for representation purposes, the foot portion 13 is not shown.

[0156] [Fig.8] shows how the spar 10 is arranged in the method of the invention.

[0157] The connecting skin 50 and the surface of the spar 10, or more precisely of the internal part of the spar 10, have regular undulations complementary to the undulations provided on the surface 21s of the first structural portion 21.

[0158] A portion 51 of the connecting skin 50 is arranged directly in contact with the surface 21s of the first portion of structure 21.

[0159] Another part 52 of the connecting skin 50 remains unopposed at the end of the step shown in [Fig.8].

[0160] This part 52 will however be covered in the following step by the second portion of structure 22.

[0161] [Fig.9] represents the step following the arrangement of the spar 10 on the first portion of structure 21.

[0162] On the uncovered connecting skin 52, the second portion of structure 22 can be arranged. This arrangement can be manual or automated.

[0163] Preferably, the second portion of structure 22 is prepared to the desired dimensions, compatible with the undulations of the connecting skin 50, independently of the method currently described.

[0164] In one embodiment, the preparation of the second portion of structure 22 may comprise a particular adjustment step carried out after the production of the structure described in [Fig.9]. This makes it possible to ensure exact compatibility between the second portion of structure 22 and the stack produced up to now.

[0165] Once the second portion of structure 22 has been deposited on the spar 10 and on the rest of the construction described up to that point, a complete portion of structure 20 is obtained which extends all around the spar 10, and which is separated from the latter by the surface skin 50.

[0166] In one embodiment, moreover shown in [Fig. 10], a second mold part 300, already draped with a second portion 32 of aeronautical skin 30 can then be placed facing the first mold portion 200 to form a mold comprising all the components of a blade 100.

[0167] Once the mold is formed as shown in [Fig. 10], a resin can be injected into the mold formed from the parts 200 and 300, and the assembly can be placed at a temperature and pressure sufficient to produce densification of the resin to allow the formation of a composite material blade.

[0168] These temperature parameters will be chosen according to the fibers chosen, the known thermal resistances of the structure portions 20 or even according to the chemical nature of the resin introduced into the mold to ensure the formation of a blade ultimately having the desired properties.

[0169] It will be noted that the method described in FIGS. 6 to 10 makes it possible to obtain a blade in which the structural portion 20 is present all around the spar 10.

[0170] Nevertheless, it will appear that alternatives can be derived from the description just given.

[0171] For example, it is possible to produce a blade 100 according to the embodiment of [Fig. 4] and in which the structural portion 20 is not present all around the spar. For this, it is necessary to have not a single first structural portion 21 as shown in [Fig. 7], but two complementary structural portions arranged for one close to the leading edge 101 and for the other to the trailing edge 102, thus leaving a part of the spar 10 not covered by the structural portion 20.

[0172] In other embodiments, the method may comprise after the step shown in [Fig.9] a step different from that shown in [Fig. 10].

[0173] For example, the second portion of aeronautical skin 32 can be draped directly over an assembly similar to that of [Fig.9].

[0174] A holding plate can then be placed on a second portion of aeronautical skin 32 thus draped, and the assembly is placed under temperature and pressure conditions allowing the formation of a blade made of composite material.

[0175] This mode of representation is not illustrated, but should not be considered as outside the invention.

[0176] Furthermore, embodiments of a method have just been described in which the surfaces of the internal part of the spar (10) and of the structural portion (20) in contact with the connecting skin (50) have a regular undulation, in order to present the most complex embodiment.

[0177] These undulations should not be considered essential, and embodiments where they are not present can be obtained in the manner just described.

Claims

Claims

1. Method for manufacturing a turbomachine blade (100) comprising: - draping at least one surface of the internal portion of a spar (10) with fibers of a first nature intended to form a connecting skin (50), the spar comprising a metal blade root portion and an at least partially metal internal portion, the draping being carried out by an automatic draping method; - draping a surface (201) of a mold (200) by fibers (31) of a second nature to form an aeronautical skin (30), the draping being carried out by an automatic draping method; - the arrangement of a first portion of structure (21) on the fibers of a second nature arranged on the surface of the mold (31); - the arrangement of the spar (10) draped over the first portion of structure; - the arrangement of a second portion of structure (22) on the draped spar; - closing the mold; - possibly a resin injection step; and - the application to the closed mold of temperature and pressure conditions allowing the formation of a blade in composite material.

2. A method according to claim 1 comprising: - draping of at least one surface of the internal part of a spar (10) by fibers of a first nature intended to form a connecting skin (50), the spar comprising a metallic blade root part (13) and an internal part at least partially metallic, the draping being carried out by an automatic draping method; - draping a first surface (201) of a mold (200) with fibers of a second nature (31) to form an aeronautical skin (30); - draping a second surface of a mold (300) with fibers of a second nature (32) to form an aeronautical skin, the draping being carried out by an automatic draping method; - the arrangement of a first portion of structure (21) on the aeronautical skin (30) arranged on the first surface of the mold (201); - the arrangement of the draped spar on the first portion of structure (21); - the arrangement of a second portion of structure (22) on the spar draped; - closing the mold by arranging the second surface of the mold (300), draped by fibers of a second nature (32) on the second portion of structure; - possibly a step of injecting a resin; - applying temperature and pressure conditions allowing the formation of a blade in composite material.

3. Method according to claim 1 comprising: - draping at least one surface of the internal part of a spar (10) by fibers of a first nature intended to form a connecting skin (50), the spar comprising a metal blade root part (13) and an internal part at least partially metallic, the draping being carried out by an automatic draping method; - draping a first surface (201) of a mold (200) by fibers of a second nature (32) to form an aeronautical skin (30), the draping being carried out by an automatic draping method; - arranging a first structural portion (21) on the aeronautical skin (30) arranged on the first surface of the mold; - arranging the draped spar (10) on the first structural portion (21); - arranging a second structural portion (22) on the draped spar, the second structural portion;- the draping of the external surface of the second portion of structure (22) by fibers of a second nature (32) to form an aeronautical skin, the draping being carried out by an automatic draping method; - the arrangement of a holding plate to ensure the shape of the draped aeronautical skin is maintained; - possibly a step of injecting a resin into the loading thus obtained; - the application of temperature and pressure conditions allowing the formation of a blade in composite material.;

4. Method according to one of claims 1 to 3, in which the laying angle for at least one automatic draping step is between 10° and 20°.

5. Method according to claim 1 to 4, in which the tension of the fibers deposited during the automatic draping steps is non-zero and less than or equal to 20 daN.

6. Method according to any one of claims 1 to 5, in which the surfaces of the internal part of the spar and of the portion of structure in contact with the connecting skin have a regular undulation.

7. The method of claim 6, wherein the undulation of the surface of the inner portion of the spar and the undulation of the surface of the structural portion comprise a period of between 20 mm and 70 mm.

8. A method according to claim 6 or 7, wherein the undulation of the surface of the inner portion of the spar and the undulation of the surface of the structural portion comprise an amplitude of between 2.0 mm and 8.0 mm.

9. A method according to any one of claims 1 to 8, wherein the fibers of a first nature are carbon fibers.

10. Method according to any one of claims 1 to 9, in which the fibers of a second nature are chosen from glass, carbon, aramid fibers or a mixture of at least two of these elements.

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