Blade de-icing device

Abstract

This disclosure relates to an aircraft part comprising: a composite material structure including a fiber 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. Figure for the abstract: Fig. 6

Description

Description Title of the invention: Blade de-icing device FIELD OF THE INVENTION This application relates to the field of aeronautics. More specifically, this application concerns the de-icing of aircraft parts, typically parts made of composite materials, and more particularly aircraft engine blades. STATE OF THE ART Certain aircraft parts, such as an engine blade, are exposed to a flow of cold air during aircraft operation. Such exposure can lead to the formation and subsequent accretion of ice on any surface of these parts exposed to the cold airflow, which can jeopardize aircraft operation. Description of the invention One 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. For this purpose, one aspect of the disclosure proposes 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 room, the heating element being embedded in the matrix of the composite material structure. Advantageously, but optionally, the part according to the disclosure may include at least one of the following features, taken alone or in combination: - the heating element is separate from the fibrous reinforcement; - the heating element includes a matrix-permeable support and a heating element configured to heat the part; - the heating element includes an electrically conductive portion, the support being configured to electrically insulate the electrically conductive portion from the fibrous reinforcement: - 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 element being positioned between the first layer and the second layer; - at least one of the first and second layers comprises a woven portion and / or a knitted portion; - at least one of the first and second layers includes a cast-on mesh fabric; - it comprises a plurality of heating elements distributed with varying densities depending on their position within the room; and - the part is an aircraft engine blade, the blade preferably comprising a plurality of heating elements distributed with a different density according to their position within the blade, with a higher density at the base of the blade than at the tip of the blade. According to another aspect of the disclosure, a blower is proposed comprising a hub and a plurality of blades as previously described extending radially from the hub. According to another aspect of the disclosure, a manufacturing process for a part as previously described is proposed, comprising the steps of: Implementation of the fibrous reinforcement: fixing the heating element to the fibrous reinforcement; then matrix solidification. Advantageously, the process may include a step of impregnating the heating element with the matrix. If necessary, the heating element may be attached to the fibrous reinforcement before or after this step. Furthermore, the step of impregnating the heating element with the matrix may be carried out concurrently with 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. DESCRIPTION OF FIGURES Other features, purposes, and benefits of disclosure will become apparent from the following description, which is purely illustrative and not exhaustive, and should be read in conjunction with the accompanying drawings on which: Fig. 1 is a schematic cross-sectional view of an aircraft propulsion assembly. Fig. 2 is a schematic cross-sectional view of another aircraft propulsion assembly. Figure 3 illustrates various components of an aircraft engine blade. Figure 4 illustrates part of a vane according to one embodiment. [Fig.5] is a cross-sectional view of [Fig.4]. Figure 6 illustrates a dawn according to one embodiment. Figure 7 is a flowchart showing a method of implementing a process for manufacturing a blade. Throughout the figures, similar elements bear identical reference numerals. DETAILED DESCRIPTION OF THE INVENTION Propulsion system Figure [Fig.1] illustrates a propulsion assembly 1 having a longitudinal axis XX, and comprising an engine 2 (or turbomachine) and a nacelle 3 surrounding the engine 2. The propulsion unit 1 is intended to be mounted on an aircraft (not shown), such as an airplane or helicopter, for example under the aircraft's wing, on the wing, or at the rear of the aircraft's fuselage. In this respect, the propulsion unit 1 may include a mast (not shown) for connecting the propulsion unit 1 to a part of the aircraft. Engine 2, illustrated in [Fig. 1], is a twin-spool, twin-flow, direct-drive turbojet. However, this is not a limiting factor, as engine 2 may have a different number of spools and / or flows, and / or be another type of turbojet, such as a geared turbojet or a turboprop. Unless otherwise specified, the terms "upstream" and "downstream" refer to the overall direction of airflow through the operating propulsion unit 1. 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 defined as a circle lying on a radial plane and whose center lies on 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 "interior" (or "internal") and "exterior" (or "external") are used with reference to a radial direction, so that the interior part of an element is, along a radial direction, closer to the longitudinal axis XX than the exterior part of the same element. As shown in [Fig. 1], the engine 2 comprises, from upstream to downstream, a blower 20, a compression section 22 including a low-pressure compressor 220 and a high-pressure compressor 222, a combustion chamber 24, and an expansion section 26 including a high-pressure turbine 262 and a low-pressure turbine 260. The blower 20, the rotor portion of the low-pressure compressor 220, and the rotor portion of the low-pressure turbine 260 are connected by a shaft The low-pressure section 280 extends along the longitudinal axis XX, with the blower 20, the low-pressure compressor 220, and the low-pressure turbine 260 forming a single low-pressure unit. The rotor portion of the high-pressure compressor 222 and the rotor portion of the high-pressure turbine 262 are connected by a high-pressure shaft 282 extending along the longitudinal axis XX, thus the high-pressure compressor 222 and the high-pressure turbine 262 forming a single high-pressure unit. As shown in [Fig. 1], the compression section 22, the combustion chamber 24, and the expansion section 26 are enclosed by an engine casing 23, while the blower 20 is enclosed by a blower casing 25.The engine casing 23 and the blower casing 25 are connected to each other by profiled structural arms 27 forming straighteners (or OGV for "Outlet Guide Vanes" in Anglo-Saxon terminology) distributed circumferentially all around the longitudinal axis XX. The longitudinal axis X-X forms an axis of rotation for the blower 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 around the longitudinal axis XX relative to the engine casing 23 and the blower casing 25. The nacelle 3 extends radially outside the engine 2, all around the longitudinal axis XX, so as to surround both the fan housing 25 and the engine housing 23, and to define, with a downstream part of the engine housing 23, a downstream part of a secondary flow B, the upstream part of the secondary flow B being defined by the fan housing 25 and an upstream part of the engine housing 23. The upstream part of the nacelle 3 further defines an air inlet 29 through which the fan 20 draws in the airflow circulating through the propulsion assembly 1. During operation, the fan 20 draws in an airflow, a portion of which, circulating within a primary stream 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 stream A passes completely through the engine casing 23. Another portion of the airflow circulates within the secondary stream B, which takes an elongated annular form surrounding the engine casing 23. The air drawn in by the fan 20 is 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 to power the aircraft on which the propulsion unit 1 is mounted. Fig. 2 illustrates another propulsion assembly 1, also presenting a longitudinal axis XX, and also including an engine 2, as well as a nacelle 3. Unlike engine 2 of propulsion unit 1 shown in [Fig. 1], engine 2 of propulsion unit 1 shown in [Fig. 2] does not include a shrouded fan 20, but an unshrouded fan 20 (or propeller). Nacelle 3, on the other hand, intended to be fixed to the aircraft, in the same way as for the propulsion assembly 1 illustrated in [Fig. 1], and also defines an air inlet 29. The propulsion assembly 1 illustrated in [Fig. 2] is of the "Open-Rotor" type, more specifically in a configuration called "pusher", that is to say, in which the unfaired fan 20 is positioned downstream of the engine 2 and downstream of the air inlet 29. This is not, however, limiting, since a propulsion assembly 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 fan 20, or downstream of the fan 20. In [Fig. 2], the blower 20 comprises two counter-rotating rotor stages 200, 202, meaning that, during operation, the rotor stages 200, 202 are driven in rotation around the longitudinal axis XX in opposite directions. This is not, however, a limitation, since the blower 20 can also comprise a rotor stage, driven in rotation around the longitudinal axis, and a stator stage, fixed in rotation. The stator stage is positioned downstream of the rotor stage and acts as a rectifier to straighten the airflow drawn in by the rotor stage. Generally, the rotor stages 200, 202 of an open-rotor blower rotate more slowly than those of a shrouded blower 20. Furthermore, the length of the fan blades 2000 of fan 20 is greater for a propulsion assembly 1 as illustrated in [Fig.2] than for a propulsion assembly 1 as illustrated in [Fig.1]. The fan blades 2000 of a propulsion assembly 1 as illustrated in [Fig.2] are therefore particularly sensitive to the phenomenon of ice accretion. In addition, 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. As seen in [Fig.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 fixed in rotation with the rotor stages 200, 202 of the blower 20. This is not limiting however since, when the blower 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 blower 20, while a stator part of the low pressure turbine 260 is connected to the stator stage of the blower 20.The compression section 22, the combustion chamber 24 and the expansion section 26 are surrounded by the nacelle 3. In operation, each of the blower 20 and the compression section 22 An airflow is drawn in. 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. 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 to benefit the aircraft on which the propulsion unit 1 is mounted. The motor 2 of each of the propulsion assemblies illustrated in [Fig. 1] and [Fig. 2] comprises at least one rotor, typically the blower 20, and a stator, typically the rectifier 27, each comprising a hub 2001, 2701, centered on the longitudinal axis XX, and from which a plurality of blades 2000, 2700 extend radially. Blade for motor As shown in [Fig. 3], at least one of the blades 2000, 2700 of engine 2, typically all the blades 2000, 2700 of the fan 20 and the stator 27, comprises a blade 4 and a foot 5, the foot 5 enabling the blade 2000, 2700 to be attached to the hub 2001, 2701. The blade 2000, 2700, and more specifically the blade 4, may include a composite material structure comprising a fibrous reinforcement embedded in a matrix. This effectively optimizes the mass of the propulsion assembly 1 and improves its performance. Fiber reinforcement can be formed from a single-piece fiber (or textile) preform, obtained through three-dimensional or multi-layer weaving with varying thickness. It can include warp and weft strands. Three-dimensional weaving generally indicates that the warp strands follow sinuous paths to link together weft strands belonging to different weft layers, with the exception of unlinking. It should be noted that a three-dimensional weave, particularly an interlock weave, can include 2D surface weaves. Various three-dimensional weaves can be used, such as interlock, multi-satin, or multi-layer weaves. Fiber reinforcement can thus include woven (two-dimensional or three-dimensional), braided, knitted, or laminated fiber arrays.The fibers of the fibrous reinforcement may include one of the following materials: carbon, glass, basalt, aramid, polypropylene and / or ceramic. The matrix typically comprises an organic material (thermoset, thermoplastic, or elastomer) or a carbon matrix. For example, the matrix may include a plastic material, typically a polymer, such as epoxy, bismaleimide, or polyimide. Blade 4 has, at least in part, an aerodynamic profile designed to be placed in a flow when the propulsion unit 1 is operating, in order to generate lift. The aerodynamic profile comprises an intrados 40, an extrados 42, a leading edge 44 and a trailing edge 46. The leading edge 44 is configured to extend in relation to the airflow within the propulsion assembly 1, and corresponds to the forward part of an airfoil that faces the airflow and divides the airflow into an underside flow 40 and an upper side flow 42. The trailing edge 46, for its part, corresponds to the rear part of the airfoil, where the underside flows 40 and 42 meet. The underside 40 and even the upper side 42 of the blade 4 can be covered with a polyurethane film for erosion protection. In one embodiment (not shown), the blade may comprise two skins, which are joined together and extend generally opposite each other. The skins are shaped to define the aerodynamic profile. The skins are made of a composite material comprising fiber reinforcement densified by a matrix. They are therefore monolithic and are made in one piece according to a non-limiting embodiment. Alternatively, it is possible to consider one fiber reinforcement for the lower surface (intrados) and another for the upper surface (extrados). In an embodiment illustrated in [Fig.3], the 2000, 2700 vane further comprises a spar 6, a filler piece 7 and a shield 8. The spar 6 may comprise, as illustrated in [Fig. 3], a blade root portion 5 extending outside the blade 4 and a blade portion arranged inside the blade 4 to form a web. The blade root portion 5 is configured to be inserted into the hub 2001, 2701. The spar 6 may be made of metal and be 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: steel, titanium, titanium alloy (in particular TA6V, comprising titanium, aluminum, vanadium, and traces of carbon, iron, oxygen, and nitrogen), nickel-based superalloy such as Inconel, or aluminum alloy. Alternatively, the spar 6 may comprise a composite material comprising a matrix-densified fiber reinforcement.Similar to the composite material structure of blade 4, the matrix of spar 6 typically comprises an organic material (thermoset, thermoplastic, or elastomer) or a carbon matrix. For example, the matrix comprises a plastic material, typically a polymer, such as epoxy, bismaleimide, or polyimide. The fibers of the spar 6's free-floating reinforcement comprise at least one of the following materials: carbon, glass, basalt, aramid, polypropylene, and / or ceramic. The fibrous reinforcement of spar 6 may include woven (two-dimensional or three-dimensional), braided, knitted, or laminated fiber arrangements. The matrix of spar 6 and the matrix of the composite material structure of blade 2000, 2700 may, where applicable, be identical. The fibers of the fibrous reinforcement of spar 6... These can be made from a material identical or different from the fibers of the fibrous reinforcement of the composite 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 radially oriented and the weft predominantly oriented along the chord of the blade 4 at the aerodynamic rib height. However, the spar 6 can also be a more mechanically advantageous assembly of different organic matrix composite materials (thermoset, thermoplastic, or elastomer) reinforced with long fibers (carbon, glass, aramid, polypropylene) in various fiber arrangements (woven, braided, knitted, unidirectional). The filler piece 7 is positioned within the airfoil structure of the blade 4 and surrounds the spar 6. The filler piece 7 can be made of a material with internal cavities, such as an organic foam (polyethacrylimide, polyethylene terephthalate (PET), polyvinyl chloride (PVC), polyetherimide (PEI), polyvinyl, carbon, polyisocyanurate, polyurethane, etc.) or a metallic foam (particularly aluminum alloy), or even a Nomex®-type honeycomb, Kevlar, fiberglass, or aluminum. Advantageously, the filler piece 7 is covered with a skin of organic matrix composite material 400 to increase the impact resistance of the blade 2000, 2700. Finally, the leading edge 44 of the airfoil can be reinforced by an added and fixed shield 8, for example by bonding. The shield 8 can be made of titanium or titanium alloy, stainless steel, steel, aluminum, nickel, etc. Heating element Figures 4 and 5 illustrate a heating element 9 configured to heat blades 2000 and 2700, while Figure 6 illustrates that the heating element 9 is attached to and fixed onto blades 2000 and 2700 by being embedded in the matrix of the composite structure. Fan blades 2000 and stator blades 27 are particularly susceptible to ice and / or frost buildup, and it is therefore particularly advantageous for the heating element 9 to be attached to this type of blade 2000 and 2700. It should be noted, however, that the heating element 9 is separate from the fibrous reinforcement of the composite material structure to which it is attached and fixed. The heating element 9 can be sewn onto the fibrous reinforcement, preferably with stitches at the ends of the heating element 9, or glued, or even fixed to the fibrous reinforcement by means of inserts, for example, of the Parisian tack type. As can be seen in [Fig.4] and [Fig.5], the heating element 9 comprises a support 90 and a heating element 92. The heating element 92 is configured to heat the blade 2000, 2700. Preferably, as illustrated in [Fig.4] and [Fig.5], the heating element 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 element 92 can also be of the chemical type, and provide heat by chemical reaction of internal components of the heating element 92, or hydraulic, and provide heat by thermal conduction of a heat transfer fluid circulating through the heating element 92. In any case, [Fig.4] and [Fig.5] illustrate that the heating element 92 includes an electrically conductive portion 920, which is configured to heat the blade 2000, 2700. The electrically conductive portion 920 can take the form of a coil, as illustrated in [Fig.[4], in order to optimize the distribution of heat produced by the heating element 92. The coil comprises a number of sections of electrical wire that are bent, identical or not, and connected to one another, for example, by being joined together, that is, by being made from a single piece. Preferably, the heating element 92 comprises a metal, such as copper, as it is a material that conducts heat well. In any case, the patterns, material, and cross-section of the electrically conductive elements in the electrically conductive portion 920 are parameters that can be adjusted according to the heating requirement. It is nevertheless necessary that the heating element, and therefore the electrical wires that compose it if applicable, be sufficiently flexible to conform to the shape of the blade 2000, 2700. As illustrated in [Fig. 4] and [Fig. 5], the support 90 is permeable to the matrix, meaning it is configured to be impregnated by the matrix. Furthermore, particularly when the heating element 92 is electrically powered and includes 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 prevents the formation of current loops in the blade 2000, 2700, which could damage it. Advantageously, as seen in [Fig. 5], the support 90 comprises a first layer 901 and a second layer 902, with the heating element 92 positioned between the first layer 901 and the second layer 902. This improves the electrical insulation of the heating element 92 from the fibrous reinforcement. Back to the [Fig.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 provide the required permeability to the matrix, as it exhibits 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 cast-iron mesh, which is illustrated in [Fig. 4], of which the . Permeability to the matrix is ​​optimal, meaning that the mesh is sufficiently large to prevent disruption of matrix flow and sufficiently narrow to ensure electrical insulation of the electrically conductive portion 920. Woven knit fabrics are made from warp loops formed lengthwise and interlaced widthwise. Such knits have the advantage of not unraveling. Generally, the backing 90 can comprise any polymer that is not electrically conductive and is also sufficiently heat-resistant, meaning that it does not deteriorate at temperatures ranging from 50°C to 100°C. Alternatively, or in addition, both of the first layer 901 and the second layer 902 are as described above, or even identical. When the heating element 92 is of the electric type, as shown in [Fig. 4] and [Fig. 5], an electrical connection element 94 can be provided, configured to electrically connect the electrically conductive portion 920 of the heating element 92 to an electrical power supply 96. This electrical connection element 94 can extend beyond the support 90, typically along the foot 5 of the blade 2000, 2700. Therefore, to prevent the electrical connection element 94 from making electrical contact with the fibrous reinforcement, an electrically insulating sheath 98 can be provided to house the electrical connection element 94, thus electrically isolating the electrical connection element 94 from the fibrous reinforcement. Thus, as seen in [Fig. 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 not limiting, however, since it is quite possible to consider different patches, distributed over the entire surface of the blade 2000, 2700. Depending on the areas 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 whose electrically conductive portion 920 carries a greater electrical power, which it dissipates by Joule effect, for example whose electrically conductive portion 920 includes thicker electrical wires.If necessary, it is possible to provide an electrical network (with a power supply and an electrical connection element 94) for each heating element 92 or, conversely, a single electrical network connected to all the heating elements, but where the surface density of the electrical wires of the different heating elements varies according to their positioning on the blade 2000, 2700. Typically, ice accretion is generally more frequent and greater at the level of a portion of the blade 4 which is close to the foot 5 of the blade 2000, 2700. The electrical network of the heating element 92 is therefore denser there. than at the top 50 of the 2000, 2700 blade where accretion is less frequent, due to the speeds reached during the rotation of the 2000, 2700 blade around the longitudinal axis XX. Different heating elements 9 can therefore be provided on the surface of the 2000, 2700 blade, preferably being distributed with a different density depending on their position within the 2000, 2700 blade. The density of the heating elements 9 corresponds here to the number of heating elements 9 per unit of space, this unit being either surface or volume.Thus, in the advantageous case where the heating elements 9 are distributed with a higher density at the foot 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 area and / or volume is greater in a region located at the foot 5 of the blade 2000, 2700 than in a region located at the head 50 of the blade 2000, 2700, as for example seen in [Fig.6]. . Manufacturing process With reference to [Fig.7], a process E for manufacturing a 2000, 2700 blade for motor 2 as previously described, generally includes the realization E1 of the fibrous reinforcement, the fixing E2 of the heating element 9 on the fibrous reinforcement and the solidification E3 of the matrix. One way to carry out this manufacturing process E is to use a vacuum resin injection process called RTM (for "Resin Transfer Molding" in English) or VARTM (for "Vacuum Assisted Resin Transfer Molding" in English). This process generally involves preparing a fibrous preform by three-dimensional weaving, then placing this preform in a mold and injecting a polymerizable resin, such as an epoxy resin, which acts as the matrix that will impregnate the preform, possibly maintaining a reduced pressure during impregnation (in the case of VARTM).After polymerization and hardening of the blade 4, and more specifically of the skin made of organic matrix composite material 400 where applicable, i.e., after solidification E3 of the matrix, the leading edge 44 of the blade 4 is reinforced by the shield 8, preferably metallic, which is attached and fixed, for example by bonding. Polymerization is a form of solidification E3 of the matrix, as is thermosetting. Thus, solidification E3 of the matrix may or may not require the application of external heat. If so, the manufacturing process E includes a curing step. Of course, other processes for solidification E3 of the matrix are conceivable, which depend in particular on the composition of the matrix. The heating element 9 can typically be attached and fixed E3 to the fibrous reinforcement within the mold, typically by bonding, before the matrix is injected so that it can impregnate both the fibrous reinforcement and the heating element 9. This is not, however, limiting, since the heating element 9 can also be attached and fixed to the fibrous 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 fibrous reinforcement can be controlled by means of a laser. In the previously described process, the heating element 9 and the fiber reinforcement are impregnated by the matrix within the mold. However, this is not a limiting factor, as it is also possible to pre-impregnate the heating element 9 and / or the fiber reinforcement with the matrix, then attach and fix the heating element 9 to the fiber reinforcement before the matrix solidifies (E3). Once the matrix has solidified, it is possible to plan a machining step of the blade 2000, 2700 to give it the desired aerodynamic profile, before attaching and fixing the shield 8 at the leading edge 44. The manufacture of the spar 6 can, however, involve several specific processes such as machining, forging, forming, casting or even additive manufacturing (3D printing). If necessary, the spar 6 and the filler piece 7 are inserted into the composite material structure before the E3 solidification step of the matrix, typically by baking. Control steps for the 2000, 2700 blade can be planned to verify its mechanical strength before fixing it to the 2001, 2701 hub. Benefits obtained By embedding the heating element 9 in the matrix of the composite material structure of the blade 2000, 2700, various advantages are obtained. First, this eliminates the need for a separate heating mat attached to part of the blade's external surface, typically between the blade and the protective shield attached to the blade's leading edge. Removing such a heating mat results in a blade without excess thickness, which could negatively impact the blade's aerodynamic properties or limit the use of the protective shield, as adhesion at the point of excess thickness would be weaker. Furthermore, a degraded protective shield design would compromise the blade's mechanical strength, especially in the event of a foreign object impact. In addition, the heating element is less susceptible to corrosion than a heating mat. Finally, the blade manufacturing process is simpler and less expensive than attaching a heating mat, as it is more easily replicated on a large scale. Furthermore, this eliminates the need to heat the dawn using hot air drawn from the Engine. Eliminating a hot air extraction de-icing system reduces the complexity and mass of the engine, which no longer includes ducts for carrying hot air from the engine to the parts to be defrosted. Furthermore, it reduces the manufacturing complexity of the cooling fan blades, which no longer need internal ducts to receive the hot air. Finally, this eliminates the need to design the blade to deform under the internal pressure of a fluid to break through ice accumulated on its external surface, as with the leading edge of an aircraft wing. Such geometric modifications to the blade would, in fact, be too detrimental to engine efficiency. Although a heating element attached to an aircraft engine blade has been described, this is not an exhaustive list. An aircraft may also include other parts with a composite 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 components of the aircraft fuselage or nacelle. Furthermore, such composite parts may be manufactured using the manufacturing process described earlier.

Claims

Demands

1. Aircraft part comprising: a composite material structure including embedded fibrous reinforcement in a matrix; at least one heating element (9) configured to heat the room, the heating element (9) being introduced into the matrix of the structure in composite material.

2. Part according to claim 1, wherein the heating element (9) is distinct from the fibrous reinforcement.

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

4. Part according to claim 3, wherein the heating element (92) includes an electrically conductive portion (920), the support (90) being configured to electrically isolate the electrically conductive (920) of the fibrous reinforcement.

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

6. Part according to any one of claims 3 to 5, wherein the support (90) includes 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 the first layer (901) and the second layer (902) includes 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 the second layer (902) includes a thrown knit fabric.

9. A piece according to any one of claims 1 to 8, comprising a plurality of heating elements (9) distributed with a different density according to their position within the room.

10. Part according to any one of claims 1 to 9, wherein the part is a blade (2000, 2700) for aircraft engine (2), the blade (2000, 2700) preferably comprising a plurality of heating elements (9) distributed with varying densities depending on their position within the dawn (2000, 2700), with a higher density at the foot level (5) of the blade (2000, 2700) than at the level of the head (50) of the blade (2000, 2700).

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

12. A method for manufacturing a part according to any one of claims 1 to 10, including the steps of: realization (El) of the fibrous reinforcement; fixing (E2) of the heating element (9) onto the fibrous reinforcement; then solidification (E3) of the matrix.