Blade de-icing device

Abstract

The present invention relates to an aircraft part comprising: a composite material structure comprising a fibrous reinforcement embedded in a matrix; a heating element (9) configured to heat the part, the heating element (9) being embedded in the matrix of the composite material structure.

Description

[0001] DAWN DE-ICING DEVICE

[0002] FIELD OF THE INVENTION

[0003] This application relates to the field of aeronautics. More specifically, this application relates to the deicing of aircraft parts, typically parts made of composite material, and more particularly aircraft engine blades.

[0004] STATE OF THE ART

[0005] Certain parts of an aircraft, such as an engine blade, are exposed to a flow of cold air during aircraft operation. Such exposure may result in the formation and subsequent accretion of ice on a surface of these parts that is exposed to the flow of cold air, which may jeopardize the operation of the aircraft.

[0006] STATEMENT OF THE INVENTION

[0007] An aim of the invention is to prevent the formation and / or accretion of ice on a surface of an aircraft part in a simple, inexpensive and easily industrializable manner.

[0008] For this purpose, according to one aspect of the disclosure, there is provided an aircraft part comprising: a composite material structure comprising a fibrous reinforcement embedded in a matrix; at least one heating element configured to heat the part, the heating element being embedded in the matrix of the composite material structure.

[0009] Advantageously, but optionally, the part according to the disclosure may comprise at least one of the following characteristics, taken alone or in combination:

[0010] - the heating element is separate from the fibrous reinforcement;

[0011] - the heating element comprises a matrix-permeable support and a heating member configured to heat the part;

[0012] - the heating member comprises an electrically conductive portion, the support being configured to electrically insulate the electrically conductive portion from the fibrous reinforcement;

[0013] - it further comprises an electrical connection element configured to electrically connect the electrically conductive portion to an electrical power source; and an electrically insulating sheath receiving the electrical connection element so as to electrically insulate the electrical connection element from the fibrous reinforcement; - the support comprises a first layer and a second layer, the heating member being positioned between the first layer and the second layer;

[0014] - at least one of the first layer and the second layer comprises a woven portion and / or a knitted portion;

[0015] - at least one of the first layer and the second layer comprises a warp knit fabric;

[0016] - it comprises a plurality of heating elements distributed with a different density depending on their position within the room; and

[0017] - the part is a blade for an aircraft engine, the blade preferably comprising a plurality of heating elements distributed with a different density depending on their position within the blade, with a higher density at the level of the root of the blade than at the level of the head of the blade.

[0018] According to another aspect of the disclosure, there is provided a fan comprising a hub and a plurality of blades as previously described extending radially from the hub.

[0019] According to another aspect of the disclosure, there is provided a method of manufacturing a part as previously described, comprising the steps of: producing the fiber reinforcement; fixing the heating element on the fiber reinforcement; then solidifying the matrix.

[0020] Advantageously, the method may comprise a step of impregnating the heating element with the matrix. Where appropriate, the fixing of the heating element to the fibrous reinforcement may be carried out before or after the step of impregnating the heating element with the matrix. Furthermore, the step of impregnating the heating element with the matrix may be carried out at the same time as a step of impregnating the fibrous reinforcement with the matrix or, alternatively, before or after the step of impregnating the fibrous reinforcement with the matrix.

[0021] DESCRIPTION OF FIGURES

[0022] Other features, purposes and advantages of the disclosure will become apparent from the following description, which is purely illustrative and not limiting, and which should be read in conjunction with the accompanying drawings in which:

[0023] Figure 1 is a schematic sectional view of an aircraft propulsion assembly.

[0024] Figure 2 is a schematic sectional view of another aircraft propulsion assembly. Figure 3 illustrates various components of an aircraft engine blade.

[0025] Figure 4 illustrates a portion of a blade according to one embodiment.

[0026] Figure 5 is a sectional view of Figure 4.

[0027] Figure 6 illustrates a blade according to one embodiment.

[0028] Figure 7 is a flowchart showing one method of implementing a blade manufacturing process.

[0029] Throughout the figures, similar elements have identical references.

[0030] DETAILED DESCRIPTION OF THE INVENTION

[0031] Propulsion unit

[0032] Figure 1 illustrates a propulsion unit 1 having a longitudinal axis XX, and comprising an engine 2 (or turbomachine) and a nacelle 3 surrounding the engine 2.

[0033] The propulsion assembly 1 is intended to be mounted on an aircraft (not shown), such as an airplane or a helicopter, for example under the wing of the aircraft, on the wing or at the rear of the fuselage of the aircraft. In this regard, the propulsion assembly 1 may comprise a mast (not shown) intended to connect the propulsion assembly 1 to a part of the aircraft.

[0034] The engine 2 illustrated in Figure 1 is a twin-spool, twin-flow, direct-drive turbojet engine. This is not, however, limiting since the engine 2 may have a different number of spools and / or flows, and / or be another type of turbojet engine, such as a geared turbojet engine or a turboprop engine.

[0035] Unless otherwise specified, the terms “upstream” and “downstream” are used with reference to the overall direction of airflow through the propulsion unit 1 in operation. Similarly, an axial direction corresponds to the direction of the longitudinal axis XX and a radial direction is a direction perpendicular to the longitudinal axis XX and intersecting the longitudinal axis XX. Furthermore, an axial plane is a plane containing the longitudinal axis XX and a radial plane is a plane perpendicular to the longitudinal axis XX. A circumference is understood to be a circle belonging to a radial plane and whose center belongs to the longitudinal axis XX. A tangential or circumferential direction is a direction tangent to a circumference: it is orthogonal to the longitudinal axis XX but does not pass through the longitudinal axis XX.Finally, the adjectives "inner" (or "internal") and "outer" (or "external") are used in reference to a radial direction so that the inner part of an element is, in a radial direction, closer to the longitudinal axis XX than the outer part of the same element. As seen in Figure 1, the engine 2 comprises, from upstream to downstream, a fan 20, a compression section 22 comprising a low pressure compressor 220 and a high pressure compressor 222, a combustion chamber 24 and an expansion section 26 comprising a high pressure turbine 262 and a low pressure turbine 260. The fan 20, the rotor part of the low pressure compressor 220, and the rotor part of the low pressure turbine 260 are connected to each other by a low pressure shaft 280 extending along the longitudinal axis XX, the fan 20, the low pressure compressor 220 and the low pressure turbine 260 then forming a low pressure body.The rotor portion of the high-pressure compressor 222 and the rotor portion of the high-pressure turbine 262 are connected to each other by a high-pressure shaft 282 extending along the longitudinal axis XX, the high-pressure compressor 222 and the high-pressure turbine 262 then forming a high-pressure body. As seen in FIG. 1, the compression section 22, the combustion chamber 24 and the expansion section 26 are surrounded by a motor casing 23, while the fan 20 is surrounded by a fan casing 25. The motor casing 23 and the fan casing 25 are connected to each other by profiled structural arms 27 forming rectifiers (or OGV for "Outlet Guide Vanes" in English terminology) distributed circumferentially all around the longitudinal axis XX.The longitudinal axis XX forms an axis of rotation for the fan 20, the rotor part of the compression section 22 and the rotor part of the expansion section 26, which are capable of being driven in rotation about the longitudinal axis XX relative to the engine casing 23 and the fan casing 25.

[0036] The nacelle 3 extends radially outside the engine 2, all around the longitudinal axis XX, so as to surround both the fan casing 25 and the engine casing 23, and to define, with a downstream portion of the engine casing 23, a downstream portion of a secondary flow path B, the upstream portion of the secondary flow path B being defined by the fan casing 25 and an upstream portion of the engine casing 23. The upstream portion of the nacelle 3 further defines an air inlet 29 through which the fan 20 draws in the air flow circulating through the propulsion unit 1.

[0037] In operation, the fan 20 draws in an air flow, a portion of which, circulating within a primary vein A, is successively compressed within the compression section 22, ignited within the combustion chamber 24 and expanded within the expansion section 26 before being ejected from the engine 2. The primary vein A passes through the engine casing 23 from one side to the other. Another portion of the air flow circulates within the secondary vein B which takes an elongated annular shape surrounding the engine casing 23, the air drawn in by the fan 20 being straightened by the straighteners 27 and then ejected from the propulsion unit 1. In this way, the propulsion unit 1 generates thrust. This thrust can, for example, be used for the benefit of the aircraft on which the propulsion unit 1 is attached and fixed.

[0038] Figure 2 illustrates another propulsion unit 1, also having a longitudinal axis XX, and also comprising an engine 2, as well as a nacelle 3.

[0039] Unlike the engine 2 of the propulsion unit 1 illustrated in FIG. 1, the engine 2 of the propulsion unit 1 illustrated in FIG. 2 does not comprise a ducted fan 20, but an unducted fan 20 (or propeller). The nacelle 3 is, for its part, intended to be fixed to the aircraft, in the same way as for the propulsion unit 1 illustrated in FIG. 1, and also defines an air inlet 29. The propulsion unit 1 illustrated in FIG. 2 is of the “Open-Rotor” type, more particularly in a configuration called “pusher”, that is to say in which the unducted fan 20 is positioned downstream of the engine 2 and downstream of the air inlet 29.This is not, however, limiting, since a propulsion unit 1 of the “Open-Rotor” type can also be in a configuration called “puller”, in which the fan 20 is positioned upstream of the engine 2, the air inlet 29 being positioned upstream of the fan 20, between the two rotor stages 200, 202 of the fan 20, or downstream of the fan 20.

[0040] In Figure 2, the fan 20 comprises two counter-rotating rotor stages 200, 202, that is to say that, in operation, the rotor stages 200, 202 are driven in rotation about the longitudinal axis XX in opposite directions. This is not, however, limiting, since the fan 20 can also comprise a rotor stage, driven in rotation about the longitudinal axis, and a stator stage, fixed in rotation, the stator stage being positioned downstream of the rotor stage, and behaves like a rectifier in order to straighten the air flow sucked in by the rotor stage. Generally, the rotor stages 200, 202 of an open-rotor rotate less quickly than a ducted fan 20. Furthermore, the length of the fan blades 2000 is greater for a propulsion unit 1 as illustrated in FIG. 2 than for a propulsion unit 1 as illustrated in FIG. 1.The blades 2000 of the fan 20 of a propulsion unit 1 as illustrated in FIG. 2 are therefore particularly sensitive to the phenomenon of ice accretion. Furthermore, unlike a shrouded fan 20, the entire surface of the blades 2000 of the rotor stages 200, 202 of the fan 20 of an open-rotor can be the site of icing.

[0041] As seen in Figure 2, the engine 2 comprises, from upstream to downstream, a compression section 22, a combustion chamber 24 and an expansion section 26 comprising a high pressure turbine 262 and a low pressure turbine 260. The rotor part of the high pressure turbine 262 is connected to at least a portion of the rotor part of the compression section 22 by a high pressure shaft 282 extending along the longitudinal axis XX. The low-pressure turbine 260 comprises two rotors, each integral in rotation with the rotor stages 200, 202 of the fan 20. This is however not limiting since, when the fan 20 comprises a rotor stage and a stator stage, a rotor part of the low-pressure turbine 260 is connected to the rotor stage of the fan 20, while a stator part of the low-pressure turbine 260 is connected to the stator stage of the fan 20.The compression section 22, the combustion chamber 24 and the expansion section 26 are surrounded by the nacelle 3.

[0042] In operation, each of the fan 20 and the compression section 22 draws in a flow of air. The air A drawn in by the compression section 22 is, successively, compressed within the compression section 22, ignited within the combustion chamber 24 and expanded within the expansion section 26 before being ejected from the engine 2. The air B drawn in by the fan 20 circulates around the nacelle 3 before being ejected downstream of the propulsion unit 1. In this way, the propulsion unit 1 generates thrust. This thrust can, for example, be used for the benefit of the aircraft on which the propulsion unit 1 is attached and fixed.

[0043] The engine 2 of each of the propulsion units illustrated in FIG. 1 and in FIG. 2 comprises at least one rotor, typically the fan 20, and one stator, typically the rectifier 27, which each comprise a hub 2001, 2701, centered on the longitudinal axis XX, and from which a plurality of blades 2000, 2700 extend radially.

[0044] Engine blade

[0045] As seen in Figure 3, at least one of the blades 2000, 2700 of the engine 2, typically all the blades 2000, 2700 of the fan 20 and the rectifier 27, comprises a blade 4 and a root 5, the root 5 making it possible to fix the blade 2000, 2700 to the hub 2001, 2701. The blade 2000, 2700, and more particularly the blade 4, may comprise a structure made of composite material comprising a fiber reinforcement embedded in a matrix. This in fact makes it possible to optimize the mass of the propulsion assembly 1 and improves its performance.

[0046] The fiber reinforcement may be formed from a single-piece fiber (or textile) preform, obtained by three-dimensional or multi-layer weaving, with varying thickness. It may include warp and weft strands. Three-dimensional weaving generally indicates that the warp strands follow sinuous paths in order to interconnect weft strands belonging to different weft strand layers except for delinkages, it being noted that a three-dimensional weave, in particular an interlock weave, may include 2D surface weaves. Different three-dimensional weave weaves may be used, such as interlock, multi-satin or multi-voile weaves. The fiber reinforcement may thus include woven (two-dimensional or three-dimensional), braided, knitted or laminated fiber arrangements. The fibers of the fiber reinforcement may comprise one of the following materials: carbon, glass, basalt, aramid, polypropylene and / or ceramic.

[0047] The matrix typically comprises an organic material (thermosetting, thermoplastic or elastomeric) or a carbon matrix. For example, the matrix comprises a plastic material, typically a polymer, for example epoxy, bismaleimide or polyimide.

[0048] The blade 4 has, at least over a portion, an aerodynamic profile suitable for being placed in a flow when the propulsion unit 1 is in operation, in order to generate lift. The aerodynamic profile comprises a lower surface 40, an upper surface 42, a leading edge 44 and a trailing edge 46. The leading edge 44 is configured to extend opposite the airflow within the propulsion unit 1, and corresponds to the front part of an aerodynamic profile which faces the airflow and which divides the airflow into a lower surface flow 40 and an upper surface flow 42. The trailing edge 46, for its part, corresponds to the rear part of the aerodynamic profile, where the lower surface 40 and upper surface 42 flows meet. The lower surface 40 or even the upper surface 42 of the blade 4 may be covered with a polyurethane film for protection against erosion.

[0049] In one embodiment (not shown) the blade may comprise two skins, which are connected to each other and extend generally opposite each other. The skins are shaped so as to define together the aerodynamic profile. The skins are made of a composite material comprising the fiber reinforcement densified by a matrix. They are therefore monolithic and are made in a single piece according to a non-limiting embodiment. As a variant, it is possible to consider one fiber reinforcement for the intrados and another for the extrados.

[0050] In an embodiment illustrated in Figure 3, the blade 2000, 2700 further comprises a spar 6, a filler piece 7 and a shield 8.

[0051] The spar 6 may comprise, as illustrated in FIG. 3, a blade root portion 5 which extends outside the blade 4 and a blade portion which is arranged inside the blade 4 so as to form a core thereof. The blade root portion 5 is configured to be inserted into the hub 2001, 2701. The spar 6 may be made of metal and in a single piece, in which the blade root portion 5 and the blade portion are monolithic. The metallic material of the spar 6 may comprise at least one of the following materials: steel, titanium, titanium alloy (in particular TÀ6V, comprising titanium, aluminum, vanadium and traces of carbon, iron, oxygen and nitrogen), nickel-based superalloy such as Inconel, aluminum alloy. Alternatively, the spar 6 may comprise a composite material comprising a fibrous reinforcement densified by a matrix.Similar to the composite material structure of the blade 4, the matrix of the spar 6 typically comprises an organic material (thermosetting, thermoplastic or elastomer) or a carbon matrix. For example, the matrix comprises a plastic material, typically a polymer, for example epoxy, bismaleimide or polyimide. The fibers of the fibrous reinforcement of the spar 6 comprise at least one of the following materials: carbon, glass, basalt, aramid, polypropylene and / or ceramic. The fibrous reinforcement of the spar 6 may comprise woven (two-dimensional or three-dimensional), braided, knitted or laminated fibrous arrangements. The matrix of the spar 6 and the matrix of the composite material structure of the blade 2000, 2700 may, if appropriate, be identical. The fibers of the fiber reinforcement of the spar 6 may be made of a material identical to or different from the fibers of the fiber reinforcement of the composite material structure of the blade 2000, 2700.Ultimately, the spar 6 is preferably made of an epoxy organic matrix composite material reinforced with 3D woven carbon fibers with the warp direction predominantly oriented radially and the weft predominantly oriented along the chord of the blade 4 at the aerodynamic vein height. However, the spar 6 may also be a more mechanically advantageous assembly of different organic matrix composite materials (thermosetting, thermoplastic or elastomer) reinforced with long fibers (carbon, glass, aramid, polypropylene) in several fiber arrangements (woven, braided, knitted, unidirectional).

[0052] The filler piece 7 is placed within the aerodynamic profile structure of the blade 4 and surrounds the spar 6. The filler piece 7 may be made of a material having internal cavities, such as a foam of organic origin (polyethacrylimide, polyethylene terephthalate (PET), polyvinyl chloride (PVC), polyetherimide (PEI), polyvinyl, carbon, polyisocyanurate, polyurethane, etc.) or metallic (in particular aluminum alloy), or a honeycomb of the Nomex® type, Kevlar, fiberglass or aluminum. Advantageously, the filler piece 7 is covered with a skin of organic matrix composite material 400 to increase the resistance of the blade 2000, 2700 to impact.

[0053] Finally, the leading edge 44 of the aerodynamic profile can be reinforced by a shield 8 added and fixed, for example by gluing. The shield 8 can be titanium or titanium alloy, stainless steel, steel, aluminum, nickel, etc. Element of

[0054] Figure 4 and Figure 5 illustrate a heating element 9 configured to heat the blade 2000, 2700, and Figure 6 illustrates that the heating element 9 is attached and fixed to the blade 2000, 2700 by being embedded in the matrix of the composite material structure. The fan blades 2000 and the rectifier blades 2700 are particularly sensitive to the phenomenon of ice and / or frost accretion, and it is therefore particularly advantageous for the heating element 9 to be attached and fixed to this type of blade 2000, 2700.

[0055] It should be noted, however, that the heating element 9 is separate from the fiber reinforcement of the composite material structure, to which it is attached and fixed. The heating element 9 may be sewn onto the fiber reinforcement, preferably with stitches at the ends of the heating element 9, or glued, or even be fixed to the fiber reinforcement by means of inserts, for example of the Parisian thumbtack type.

[0056] As seen in Figure 4 and Figure 5, the heating element 9 comprises a support 90 and a heating member 92.

[0057] The heating member 92 is configured to heat the blade 2000, 2700. Preferably, as illustrated in FIG. 4 and FIG. 5, the heating member 92 operates on electrical energy and heats the blade 2000, 2700 by Joule effect by dissipating electric current flowing through it. This is not, however, limiting since the heating member 92 can also be of the chemical type, and provide heat by chemical reaction of components internal to the heating member 92, or hydraulic, and provide heat by thermal conduction of a heat transfer fluid flowing through the heating member 92. In any event, FIG. 4 and FIG. 5 illustrate that the heating member 92 comprises an electrically conductive portion 920, which is configured to heat the blade 2000, 2700.The electrically conductive portion 920 may take the form of a coil, as illustrated in FIG. 4, in order to optimize the distribution of the heat produced by the heating member 92. The coil comprises a certain number of portions of electrical wires which are bent, identical or not to each other, and connected to each other, for example by being made of one material with each other, that is to say by being obtained in one piece with each other. Preferably, the heating member 92 comprises a metal, such as copper, because it is a material which conducts heat well. In any event, the patterns, the material and the section of the electrically conductive elements in the electrically conductive portion 920 are parameters which can be adjusted according to the heating requirement.It is nevertheless appropriate that the heating member, and therefore the electrical wires which compose it where applicable, are sufficiently flexible to fit the shape of the blade 2000, 2700.

[0058] As illustrated in Figure 4 and Figure 5, the support 90 is permeable to the matrix, that is to say it is configured to be impregnated by the matrix. In addition, in particular when the heating member 92 operates with electrical energy and comprises an electrically conductive portion 920, the support 90 is configured to electrically insulate the electrically conductive portion 920 from the fibrous reinforcement, which is generally electrically conductive. This makes it possible to avoid the creation of current loops in the blade 2000, 2700, which could damage it. Advantageously, as visible in Figure 5, the support 90 comprises a first layer 901 and a second layer 902, the heating member 92 being positioned between the first layer 901 and the second layer 902. This makes it possible to improve the electrical insulation of the heating member 92 with respect to the fibrous reinforcement.Returning to Figure 4, at least one of the first layer 901 and the second layer 902 may comprise a woven portion and / or a knitted portion, which make it possible to obtain the required matrix permeability, since it has sufficient porosity. In an advantageous variant, at least one of the first layer 901 and the second layer 902 may comprise a marquisette fabric, typically a warp knit, which is illustrated in Figure 4, whose matrix permeability is optimal, that is to say that it has meshes that are wide enough so that the matrix flow is not disturbed, and narrow enough to guarantee the electrical insulation of the electrically conductive portion 920. Warp knit fabrics are made from warp loops formed lengthwise and interlaced across the width of the knit. Such knits have the advantage of not having its stitches come undone.Generally, the support 90 may comprise any polymer which is not electrically conductive and is also sufficiently heat resistant, i.e. it does not deteriorate at temperatures ranging from 50°C to 100°C. Alternatively, or in addition, both the first layer 901 and the second layer 902 are as described above, or even identical.

[0059] When the heating member 92 is of the electrical type, as visible in FIG. 4 and in FIG. 5, it is possible to provide an electrical connection element 94 configured to electrically connect the electrically conductive portion 920 of the heating member 92 to an electrical power source 96. This electrical connection element 94 can extend outside the support 90, typically along the root 5 of the blade 2000, 2700. Therefore, to prevent the electrical connection element 94 from establishing electrical contact with the fiber reinforcement, it is possible to provide an electrically insulating sheath 98 receiving the electrical connection element 94 so as to electrically insulate the electrical connection element 94 from the fiber reinforcement.

[0060] Thus, as visible in Figure 6, the heating element 9 can take the form of a mat integrated into the blade 2000, 2700, at the level of all or part of an external surface of the blade 2000, 2700. The mat shape is however not limiting, since it is entirely possible to envisage different patches, distributed over the entire surface of the blade 2000, 2700. Depending on the zones of the blade 2000, 2700 exposed to a flow of cold air, it is possible to densify certain parts with a heating element 9 configured to dissipate a greater quantity of heat, typically of which the electrically conductive portion 920 sees a greater electrical power circulate, which it dissipates by Joule effect, for example of which the electrically conductive portion 920 comprises thicker electrical wires.If necessary, it is possible to provide an electrical network (with an electrical power supply and an electrical connection element 94) per heating member 92 or, on the contrary, a single electrical network connected to all of the heating members, but where the surface density of the electrical wires of the different heating members varies according to their positioning on the blade 2000, 2700. Typically, the accretion of ice is generally more frequent and more significant at the level of a portion of the blade 4 which is close to the root 5 of the blade 2000, 2700. The electrical network of the heating member 92 is therefore denser there than at the tip 50 of the blade 2000, 2700 where the accretion is rarer, due to the speeds reached during the rotation of the blade 2000, 2700 around the longitudinal axis XX.Different heating elements 9 can therefore be provided on the surface of the blade 2000, 2700, preferably being distributed with a different density depending on their position within the blade 2000, 2700. The density of the heating elements 9 corresponds here to the number of heating elements 9 per unit of space, this unit being able to be surface or volume. Thus, in the advantageous case where the heating elements 9 are distributed with a higher density at the root 5 of the blade 2000, 2700 than at the head 50 of the blade 2000, 2700, this means that the number of heating elements per unit of surface area and / or volume is higher in a region located at the root 5 of the blade 2000, 2700 than in a region located at the head 50 of the blade 2000, 2700, as for example visible in FIG. 6.

[0061] Manufacturing process

[0062] With reference to Figure 7, a method E for manufacturing a blade 2000, 2700 for engine 2 as previously described, generally comprises the production E1 of the fiber reinforcement, the fixing E2 of the heating element 9 on the fiber reinforcement and the solidification E3 of the matrix. One way of carrying out this manufacturing method E consists of using a vacuum resin injection method called RTM (for “Resin Transfer Molding” in English terminology) or VÀRTM (for “Vacuum Assisted Resin Transfer Molding” in English terminology). This method generally consists of preparing a fiber preform by three-dimensional weaving then placing this preform in a mold and injecting a polymerizable resin such as an epoxy resin, which is the matrix which will impregnate the preform, possibly maintaining a reduced pressure during the impregnation (in the case of VÀRTM).After polymerization and hardening of the blade 4, and more precisely of the skin made of organic matrix composite material 400 where appropriate, that is to say after solidification E3 of the matrix, the leading edge 44 of the blade 4 is reinforced by the shield 8, preferably metallic, which is added and fixed, for example by gluing. Polymerization is a form of solidification E3 of the matrix, as is also thermosetting. Thus, the solidification E3 of the matrix may or may not require the supply of external heat. Where appropriate, the manufacturing method E includes a baking step. Of course, other methods of solidification E3 of the matrix are conceivable, which depend in particular on the composition of the matrix.

[0063] The heating element 9 can typically be attached and fixed E3 to the fiber reinforcement within the mold, typically by gluing, before the matrix is ​​injected, so that the latter can impregnate both the fiber reinforcement and the heating element 9. This is not, however, limiting, since the heating element 9 can also be attached and fixed to the fiber reinforcement after the latter has been impregnated by the matrix, the heating element 9 then being impregnated by the matrix before its solidification E3. Advantageously, the position of the heating element 9 on the fiber reinforcement can be controlled by means of a laser.

[0064] In what has been described previously, the heating element 9 and the fibrous reinforcement are impregnated by the matrix in the mold. This is not, however, limiting, since it is also possible to pre-impregnate the heating element 9 and / or the fibrous reinforcement with the matrix, then to attach and fix the heating element 9 on the fibrous reinforcement, before solidification E3 of the matrix.

[0065] Once the matrix has solidified, it is possible to provide a step of machining the blade 2000, 2700 to give it the desired aerodynamic profile, before attaching and fixing the shield 8 to it at the leading edge 44.

[0066] The manufacture of the spar 6 can, for its part, involve several specific processes such as for example machining, forging, forming, casting or even additive manufacturing (3D printing). If necessary, the spar 6 and the filling part 7 are inserted into the composite material structure before the solidification step E3 of the matrix, typically by baking.

[0067] Steps for checking the blade 2000, 2700 may be provided in order to check its mechanical strength before fixing it to the hub 2001, 2701.

[0068] Benefits obtained

[0069] By embedding the heating element 9 in the matrix of the composite material structure of the blade 2000, 2700, various advantages are obtained.

[0070] First, this eliminates the need to use a heating mat attached and fixed to a portion of the blade's external surface, typically between the blade and the protective shield attached and fixed to the leading edge of the blade. Removing such a heating mat makes it possible to obtain a blade that does not have excess thickness, which would be likely to penalize the aerodynamic properties of the blade or limit the possibility of using the protective shield, since the adhesion at the excess thickness would be weaker. And a degraded arrangement of the protective shield would compromise the mechanical strength of the blade, especially in the event of impact from a foreign body. In addition, the heating element is less sensitive to corrosion than the heating mat would be. Finally, the blade manufacturing process is simpler and less expensive than adding and fixing a heating mat, because it is more easily repeatable on a large scale.

[0071] Second, this eliminates the need to heat the blade using hot air taken from the engine. Removing a hot air defrosting system reduces the complexity and mass of the engine, which then no longer includes the ducts for transporting hot air from the engine to the parts to be defrosted. In addition, this reduces the manufacturing complexity of the blades, which no longer need to include internal transport ducts to receive the hot air.

[0072] Finally, this eliminates the need to provide that the blade is deformable under the internal pressure of a fluid to break the ice accumulated at the level of an external surface of the blade, as for a leading edge of an aircraft wing. Such geometric modifications of the blade would in fact be too penalizing for the efficiency of the engine.

[0073] Although a heating element attached to a blade for an aircraft engine has been described, this is not limiting. Indeed, an aircraft may also comprise other parts comprising a composite material structure, which may also incorporate a heating element, an electrical connection element and / or an electrically insulating sheath as previously described. Such parts may be parts of the aircraft fuselage or the nacelle. Furthermore, such composite material parts may be manufactured according to the manufacturing method previously described.

Claims

DEMANDS 1. Aircraft part comprising: a composite material structure including a fibrous reinforcement embedded in a matrix; a plurality of heating elements (9) distributed with a different density according to their position within the part, each heating element (9) being configured to heat the part and being embedded in the matrix of the composite material structure.

2. Part according to claim 1, in which each heating element (9) is separate from the fibrous reinforcement.

3. Part according to any one of claims 1 and 2, wherein each heating element (9) comprises: a support (90) permeable to the matrix; and a heating member (92) configured to heat the part.

4. Part according to claim 3, in which the heating element (92) comprises an electrically conductive portion (920), the support (90) being configured to electrically isolate the electrically conductive portion (920) from the fibrous reinforcement.

5. Part according to claim 4, further comprising: an electrical connection element (94) configured to electrically connect the electrically conductive portion (920) to an electrical power supply (96); and an electrically insulating sheath (98) receiving the electrical connection element (94) so ​​as to electrically insulate the electrical connection element (94) from the fibrous reinforcement.

6. Part according to any one of claims 3 to 5, wherein the support (90) comprises a first layer (901) and a second layer (902), the heating element (92) being positioned between the first layer (901) and the second layer (902).

7. Part according to claim 6, wherein at least one of the first layer (901) and of the second layer (902) comprises a woven portion and / or a knitted portion.

8. Part according to any one of claims 6 and 7, wherein at least one of the first layer (901) and of the second layer (902) comprises a cast-on mesh fabric.

9. Part according to any one of claims 1 to 8, wherein the part is a blade (2000, 2700) for an aircraft engine (2), the blade (2000, 2700) comprising a plurality of heating elements (9) distributed with a different density according to their position within the blade (2000, 2700), with a higher density at the base (5) of the blade (2000, 2700) than at the tip (50) of the blade (2000, 2700).

10. Blower (20) comprising a hub (2001) and a plurality of blades (2000) according to claim 9 extending radially from the hub (2001).

11. Method for manufacturing a part according to any one of claims 1 to 9, comprising the steps of: production (E1) of the fibrous reinforcement; fixing (E2) of the heating elements (9) on the fibrous reinforcement; then solidification (E3) of the matrix.

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