Aircraft skin structure and manufacturing method thereof

Abstract

The invention relates to the technical field of aircraft skin, and particularly discloses an aircraft skin structure and a manufacturing method thereof. The aircraft skin structure comprises an aircraft skin and a reinforcing rib plate which are made of a carbon fiber reinforced thermoplastic composite material, the aircraft skin and the reinforcing rib plate are connected together through a connecting interface layer, conductive particles are dispersed in the connecting interface layer and adopt a mixed system of carbon nanotubes and silver nanowires, the weight ratio of the carbon nanotubes to the silver nanowires is (2.5-3.5): 1, and the weight ratio of the carbon nanotubes to the silver nanowires is (2.5-3.5): 1. Part of the thermoplastic material connected with the interface layer is embedded into the inner cavity of the carbon nano tube; and the aircraft skin and the reinforcing rib plate are welded and clamped, and the aircraft skin and the reinforcing rib plate are welded together through ultrasonic welding equipment. The mechanical property of the aircraft skin structure is remarkably improved, the preparation forming efficiency of the aircraft skin structure is improved, operation is convenient, and the method is suitable for engineering application.

Description

Technical Field

[0001] The invention belongs to the technical field of aircraft skin manufacturing, and in particular relates to an aircraft skin structure and a manufacturing method thereof. Background Art

[0002] At present, thermoplastic composite materials have been used as main components such as wings and fuselage skins on domestic and foreign aircraft. They have the advantages of high specific strength, good corrosion resistance, strong energy absorption and impact resistance, and easy recycling. The demand for corresponding connection structures is increasing. High-performance thermoplastic composite materials have been applied to aircraft structures, such as helicopter fuselages, wings, tails, spoilers and other parts. This kind of thermoplastic composite material can be used as structural parts or functional parts to gradually replace traditional metal components.

[0003] The method of using thermoplastic composite materials to manufacture aircraft skin structures is mainly a hot pressing process, and the ribs and other structures are connected to the skin in a certain way. The connection methods include adhesive bonding structure or mechanical connection structure. In actual production, the adhesive bonding structure has problems such as long curing time and the adhesive layer of the finished product is easily affected by the environment. The mechanical connection structure needs to introduce rivets, bolts and other structural parts, which easily causes the structure to increase in weight and does not meet the purpose of lightweighting, and the connection is prone to stress concentration. Therefore, it is necessary to develop a new aircraft skin structure and manufacturing method.

[0004] Existing document CN119550651A discloses a fiber continuous longitudinal and transverse reinforced integral composite skin structure, in which the skin and the longitudinal and transverse reinforcement structure are co-bonded to achieve lightweight and strength improvement, but the method is complex and time-consuming to process, and the applicability of complex structures is weak; document CN108973089A discloses a method for preparing a composite material special-shaped skin structure, according to the digital model of the composite material special-shaped skin structure, a combined process cover plate matching its forming surface is prepared, and the composite material special-shaped skin structure blank, the composite process thin layer and the combined process cover plate are sequentially stacked on the mold tooling from bottom to top to form a process combination, but the preparation process of the hot pressing curing method is complicated; document CN111823601 discloses a method for reducing the difficulty of ultrasonic welding of engineering plastics by using chemical time, which adopts the method of pre-coating welding auxiliary reagents on the welding surface of the engineering plastics, so that the surface of the engineering plastics is pre-dissolved and activated before ultrasonic welding, but this method has the problem of complex selection of auxiliary reagents, and its surface treatment has limited improvement on the mechanical properties of the joint. In summary, the preparation of thermoplastic composite aircraft skin connection structures currently has the problems of complex preparation, low connection efficiency, and poor connection stability and safety between structural components. Therefore, improving the interface safety of aircraft skin connection structures while simplifying the preparation process is one of the main issues that need to be solved. Summary of the invention

[0005] In view of the problems mentioned in the background technology, the present invention aims to provide an aircraft skin structure and a manufacturing method thereof.

[0006] The present invention adopts the following technical solution.

[0007] An aircraft skin structure comprises an aircraft skin and a reinforcing rib plate made of carbon fiber reinforced thermoplastic composite materials. The aircraft skin and the reinforcing rib plate are connected together through a connecting interface layer. The connecting interface layer is dispersed with conductive particles. The conductive particles adopt a mixed system of carbon nanotubes and silver nanowires. The weight ratio of the carbon nanotubes to the silver nanowires is 2.5-3.5:1, and part of the thermoplastic material of the connecting interface layer is embedded in the inner cavity of the carbon nanotubes.

[0008] In order to further improve the stability and strength of the aircraft skin structure, the diameter of the carbon nanotubes is not less than 100nm, the diameter of the silver nanowires is not more than 55nm, and some of the silver nanowires extend into or are inserted into the inner cavity of the carbon nanotubes.

[0009] As a preferred embodiment, the length of the silver nanowire is 0.8 to 3 times the length of the carbon nanotube.

[0010] Furthermore, the steps of the method for manufacturing the aircraft skin structure of the present invention include: Step 1, preparing a hot pressing mold, an aircraft skin and a stiffening rib plate to be connected, wherein the working surface of the hot pressing mold has a forming structure adapted to the grid portion; the aircraft skin and the stiffening rib plate are both made by a hot pressing forming process; Step 2, laying conductive particles on the surface of the connection area of ​​the aircraft skin; Step 3, placing the lower die working surface of the hot pressing die in step 1 facing the connection area surface of the aircraft skin; Step 4, setting the hot pressing process parameters and starting the hot pressing equipment for hot pressing, so that the contact surface between the aircraft skin and the hot pressing mold is softened under the action of heat and forms a grid structure. In this process, the conductive particles are embedded in the grid structure and form energy conducting ribs; Step 5, welding and clamping the aircraft skin and the stiffener plate, wherein the connection area of ​​the aircraft skin is located below the stiffener plate, and the energy guiding ribs on the aircraft skin are attached to the lower surface of the stiffener plate; Step 6: Use ultrasonic welding equipment to weld the aircraft skin and the reinforcing rib plate together. During the welding process, the energy guiding ribs melt and form a connecting interface layer with the reinforcing rib plate, and then cool and solidify into shape.

[0011] As a preferred solution, the pre-laying thickness of the conductive particles in step 2 is 10-30 μm.

[0012] As a preferred solution, during the welding process, the downward pressure height of the welding tool head is not greater than half the thickness of the reinforcing rib plate, the welding amplitude is 25μm~45μm, the welding pressure is 0.1MPa~0.7MPa, and the welding time is 0.5s~3s.

[0013] Furthermore, in step 5, the aircraft skin is mounted on a cutting board in which an induction coil is embedded; the welding current during the welding process is 25-32A.

[0014] Furthermore, the energy conducting rib is provided with a plurality of vertically arranged grooves, the grooves are close to the edges of the energy conducting ribs, the depth of the grooves is equal to the height of the energy conducting ribs, and the cross-section of the grooves is a rectangle.

[0015] Furthermore, during the welding process, a magnetic field is synchronously applied to activate the eddy current effect of the conductive particles to assist in heating.

[0016] In the present invention, the aircraft skin and the stiffener plate are both made by a hot pressing process, and the steps include: Step 11, preparing carbon fiber woven cloth, thermoplastic resin granular material and a hot pressing mold, wherein the hot pressing mold is designed to have a shape that is adapted to the skin structure; Step 12, the carbon fiber woven cloth is laid in a hot pressing mold; the resin particles are heated and melted by an extruder, and the viscosity is reduced by screw shearing before entering the hot pressing mold, and the hot pressing mold is pressurized to fully infiltrate the resin and the carbon fiber, and a continuous carbon fiber reinforced thermoplastic composite layer is formed after cooling; Step 13, assembling the multi-layer continuous carbon fiber reinforced thermoplastic composite stack in a hot pressing forming mold, setting the hot pressing process parameters, applying heat and pressure through the mold to complete the connection between the continuous carbon fiber reinforced thermoplastic composite layers; Step 14, using the continuous carbon fiber reinforced thermoplastic composite layer obtained in step 13 as the bottom layer, and repeating step 13 multiple times until an aircraft skin of a target size is formed, and the target size is determined by a person skilled in the art according to an actual product model; Referring to the aforementioned steps 11 to 14, a mold suitable for the reinforcing rib plate is selected to prepare the reinforcing rib plate.

[0017] As a preferred solution, in the continuous carbon fiber reinforced thermoplastic composite layer, the fiber volume content needs to be controlled at 30%-50%, and the single layer thickness is 0.2-0.4 mm; the hot pressing temperature during the forming process of the continuous carbon fiber reinforced thermoplastic composite layer (i.e., in step 13) is the resin melting temperature ±10 ° C, the pressurization stage is divided into multiple stages of pressurization (initial 0.5-2 MPa, then increased to 2.0-4 MPa), and the holding time is 0.5-1 h; in step 6, the welding point spacing is 1±0.3 cm, and the boundary distance is 5 mm~10 mm.

[0018] After step 6 is completed, the obtained aircraft skin connection structure (also referred to as an aircraft skin connection component) is placed in a heat treatment furnace as a whole for stress relief, and then machined and surface roughness polished to form an aircraft skin connection structure of a thermoplastic composite material.

[0019] Beneficial effects: The present invention comprehensively utilizes the manufacturing methods of hot pressing and ultrasonic welding of energy guiding ribs to achieve dual optimization of aircraft skin structure and materials and modification of parent materials, significantly improves the mechanical properties of aircraft skin structure, greatly improves the energy utilization rate and forming efficiency of the connection interface, improves the preparation and forming efficiency of aircraft skin structure, is easy to operate, and is suitable for engineering applications. BRIEF DESCRIPTION OF THE DRAWINGS

[0020] Figure 1 Schematic diagram of the energy director rib forming process (lateral direction) in Example 1; Figure 2 It is a schematic diagram of the energy director rib forming process (stereoscopic) in Example 1; Figure 3 This is a (side) schematic diagram of the ultrasonically welded aircraft skin structure in Example 1; Figure 4 This is a (stereo) schematic diagram of the ultrasonically welded aircraft skin structure in Example 1; Figure 5 Schematic diagram of the energy guiding rib forming process in Example 2; Figure 6 This is a temperature-time curve of the hot pressing mold in Example 1.

[0021] Explanation of the accompanying figure numbers: 1—upper mold of the hot pressing mold, 2—aircraft skin, 3—lower mold, 4—temperature control heating element, 5—welding tool head, 6—reinforcement rib plate, 7—ultrasonic welding point, 8—anvil, 9—induction coil, 10—energy guiding rib, 11—groove, 12—first groove, 13—second groove. DETAILED DESCRIPTION

[0022] The technical solution of the present invention is described clearly and completely below in conjunction with the accompanying drawings. Obviously, the described embodiments are only part of the embodiments of the present invention, not all of the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by ordinary technicians in this field without creative work are within the scope of protection of the present invention. In the embodiments, the numbers in the accompanying drawings are partial schematic diagrams of the corresponding components. Example 1

[0023] Combination Figures 1 to 4As shown, an aircraft skin structure includes an aircraft skin 2 and a reinforcing rib plate 6 made of carbon fiber reinforced thermoplastic composite material, the aircraft skin 2 and the reinforcing rib plate 6 are connected together through a connecting interface layer, the connecting interface layer is dispersed with conductive particles, the conductive particles are a mixed system of carbon nanotubes and silver nanowires, and part of the thermoplastic material of the connecting interface layer is embedded in the carbon nanotube cavity. The manufacturing method of the aircraft skin structure includes the following steps: Step 1, preparing a hot pressing mold, an aircraft skin 2 to be connected (welded) and a reinforcing rib plate 6, the working surface of the hot pressing mold having a forming structure adapted to a grid portion (the grid portion in the present invention refers to a grid-like structure composed of a plurality of protrusions and grooves, and the protrusions are embedded with conductive particles to form energy conducting ribs 10), the aircraft skin 2 and the reinforcing rib plate 6 are both made by a hot pressing molding process; wherein the thickness of the aircraft skin 2 to be welded and the reinforcing rib plate 6 are both 3 mm; Step 2, laying conductive particles on the surface of the connection area of ​​the aircraft skin 2 to be welded, with a pre-laying thickness of 20-30 μm; the conductive particles adopt a mixed system of carbon nanotubes and silver nanowires, the weight ratio of carbon nanotubes to silver nanowires is 3:1, the diameter of the carbon nanotubes is 100-150 nm, the diameter of the silver nanowires is 55 nm, part of the silver nanowires in the mixed system extends into the inner cavity of the carbon nanotubes, and the length of the silver nanowires is 0.8-3 times the length of the carbon nanotubes; Step 3, put the lower mold 3 of the hot pressing mold in place, the lower mold 3 is provided with a temperature control heating element 4, and the aircraft skin 2 is placed on the lower mold 3; the working surface of the lower mold 3 of the hot pressing mold in step 1 is aligned with the surface of the connection area of ​​the aircraft skin 2 to be welded; Step 4, setting the hot pressing process parameters and starting the hot pressing equipment to perform hot pressing, so that the aircraft skin 2 to be welded is softened under the action of heat and forms a grid structure, and the conductive particles in this process are embedded in the grid structure and form energy conducting ribs 10, and the aircraft skin 2 with energy conducting ribs 10 is obtained; combined with Figure 6 As shown, the hot pressing heating rate is 20°C / min, and the temperature is increased to 10°C above the melting temperature Tm of the thermoplastic composite plate and then kept warm for 10 minutes. The cooling rate is used as a control variable, and the original value is 30°C / min. On this basis, each increase of 10°C / min is used as a gradient to control the crystallinity of the material; the lower pressure parameter of the upper mold 1 of the hot pressing mold is controlled to be 0.4MPa, and no pressure is applied when the temperature starts to drop. After the temperature drop is completed, the energy guiding rib 10 is obtained, and the side length of the energy guiding rib 10 is 1mm, the height is 0.2mm, and the spacing between adjacent energy guiding ribs 10 is 0.2mm; Step 5, Combine Figure 3 and Figure 4As shown, the obtained aircraft skin 2 with energy guiding ribs 10 and the stiffening rib plate 6 are welded and clamped, the aircraft skin 2 with energy guiding ribs 10 is located below the stiffening rib plate 6, the energy guiding ribs 10 on the aircraft skin 2 are attached to the lower surface of the stiffening rib plate 6, and the aircraft skin 2 is installed on the anvil 8, in which the induction coil 9 is embedded; Step 6, use ultrasonic welding equipment to weld the aircraft skin 2 with energy conducting ribs 10 and the reinforcing rib plate 6 together, the embedded induction coil 9 is started synchronously with the ultrasonic welding equipment, and during the welding process, a magnetic field is applied synchronously to activate the eddy current effect of the conductive particles for auxiliary heating, the energy conducting ribs 10 are melted and form a connection interface layer with the reinforcing rib plate 6, and then cooled and solidified (the cooling and solidification temperature reduction gradient is 10°C / min, and the air humidity at room temperature is 20-30%) to obtain an aircraft skin structure; during the welding process, the downward pressing height of the welding tool head 5 is not greater than half the thickness of the reinforcing rib plate 6, the welding amplitude is 35μm, the welding pressure is 0.2MPa, the welding time is 1s, and the welding current is 30A; the spacing between adjacent ultrasonic welding points 7 is 1cm, and the boundary distance is 6mm.

[0024] After step 6 is completed, the obtained aircraft skin connection structure is placed in a heat treatment furnace as a whole for stress elimination, and then machined and surface roughness polished to form an aircraft skin connection structure of a thermoplastic composite material.

[0025] During the welding process, under the action of longitudinal ultrasonic vibration, the energy conducting ribs 10 at the interface produce stress concentration and preferentially deform and heat up; under different cooling rates, the material near the energy conducting ribs 10 has low crystallinity and is easy to deform and heat up; while the conductive particles increase the friction heat generated in the early stage of welding, the heat generated by the eddy current effect of the induced magnetic field causes the interface to heat up, and after the interface materials are heated to the melting temperature, they diffuse and spread with each other, and the upper and middle parts of the energy conducting ribs 10 and the melted part of the material on the lower surface of the reinforcing rib plate 6 form a weld point, and the melted part of the material on the lower surface of the reinforcing rib plate 6 spreads to the grid groove corresponding to the unmelted part of the energy conducting rib 10 to form a mechanical interlocking structure. In this scheme, the dual regulation of the energy conducting rib structure and the material is cleverly achieved, so that the ultrasonic welding energy is concentrated at the interface, which improves the connection strength and stability of the aircraft skin connection structure, as well as the welding efficiency and quality, and also greatly improves the energy utilization rate of the ultrasonic welding interface. Example 2

[0026] A method for manufacturing an aircraft skin structure, referring to Example 1, the difference between the method and Example 1 mainly includes: combining Figure 5As shown, the energy conducting rib 10 has a plurality of vertically arranged grooves 11, the grooves 11 are close to the edge of the energy conducting rib 10, the height of the grooves 11 is equal to the height of the energy conducting rib 10, the width of the grooves 11 is 0.1 mm, the spacing between adjacent grooves 11 is 0.1 mm, and the two facing grooves 11 on two adjacent energy conducting ribs 10 are staggered (for example: Figure 5 The first groove 12 and the second groove 13 on two different energy conducting ribs 10 are staggered, and the first groove 12 and the second groove 13 can be collectively referred to as groove 11); the side length of the energy conducting rib 10 is 2mm, the height is 0.3mm, and the spacing between adjacent energy conducting ribs 10 is 0.2mm. In this embodiment, further advantages include: the melted part of the material on the lower surface of the reinforcing rib plate 6 spreads into the grid grooves and grooves 11 corresponding to the unmelted part of the energy conducting rib 10, and branches grow from the grooves 11 to form a mechanical structure similar to "chip pins" and "interpenetrating (branch pins grown in face-to-face grooves 11 interpenetrate with each other, such as branches grown in multiple first grooves 12 and multiple second grooves 13 interpenetrate with each other)", and its stability and mechanical properties are further optimized, especially the shear resistance is significantly improved.

[0027] In the embodiment, the steps of manufacturing the aircraft skin and the stiffener plate by a hot pressing process include: Step 11, preparing carbon fiber woven cloth, thermoplastic resin granular material and a hot pressing mold, wherein the hot pressing mold is designed to have a shape that is adapted to the skin structure; Step 12, the carbon fiber woven cloth is laid in a hot pressing mold; the resin particles are heated and melted by an extruder, and the viscosity is reduced by screw shearing before entering the hot pressing mold, and the hot pressing mold is pressurized to fully infiltrate the resin and the carbon fiber, and a continuous carbon fiber reinforced thermoplastic composite layer is formed after cooling; Step 13, assembling the multi-layer continuous carbon fiber reinforced thermoplastic composite stack in a hot pressing forming mold, setting the hot pressing process parameters, applying heat and pressure through the mold to complete the connection between the continuous carbon fiber reinforced thermoplastic composite layers; The fiber volume content of the continuous carbon fiber reinforced thermoplastic composite layer is 40%, and the single layer thickness is 0.4 mm; the hot pressing temperature during the forming process of the continuous carbon fiber reinforced thermoplastic composite layer (i.e., step 13) is equal to the melting temperature of the resin, and the pressurization stage is divided into multiple stages (initial 1.5 MPa, then increased to 3 MPa), and the holding time is 0.8 h; Step 14, using the continuous carbon fiber reinforced thermoplastic composite layer obtained in step 13 as the bottom layer, and repeating step 13 multiple times until an aircraft skin of a target size is formed; Referring to the aforementioned steps 11 to 14, a mold suitable for the reinforcing rib plate is selected to prepare the reinforcing rib plate.

[0028] In Examples 1 and 2, conductive particles are dispersed in the connection interface layer between the aircraft skin 2 and the reinforcing rib plate 6 behind the energy conducting ribs 10, and part of the material (resin) of the thermoplastic composite sheet is embedded in the carbon nanotube lumen, and part of the silver nanowires extends into (inserts) the carbon nanotube lumen. This structure in which the resin, carbon nanotubes, and silver nanowires are interlocked significantly improves the mechanical properties of the connection interface layer.

[0029] Comparative Example 1: Referring to Example 1, the difference between it and Example 1 is that the conductive particles and the corresponding hot pressing step are omitted.

[0030] Comparative Example 2: Referring to Example 2, the difference between it and Example 2 is that the conductive particles and the corresponding hot pressing step are omitted.

[0031] The aircraft skin structures (connection structures) obtained in the examples and comparative examples were tested for mechanical properties, and the results showed that in Example 1, the strength of the connection structure was 29.7 MPa, the joint shear force was 18540 N, and the effective connection area was increased by about 43% compared with the joint without energy guiding ribs; in Example 2, the strength of the connection structure was 34.22 MPa, and the joint shear force was 21140 N; in Comparative Example 1, the strength of the connection structure was 25.3 MPa, and the joint shear force was 15811 N; in Comparative Example 2, the strength of the connection structure was 26.98 MPa, and the joint shear force was 16289 N.

[0032] It can be seen from the interface temperature curve of the welding process that the effective connection time (Tm<T<Td) in the embodiment is longer than that in the comparative embodiment, the welding interface temperature is mainly distributed near the interface, and the corresponding energy utilization rate in the embodiment is higher. The connection structure in the embodiment is subjected to defect detection (welding joint detection), and the results show that there are no defects such as pores.

[0033] Comparative Example 3, the component structure refers to Example 1, and the difference between Example 3 and Example 1 is that the connection between the aircraft skin and the reinforcement ribs is bonding.

[0034] The aircraft skin connection structures of the same specifications in Example 1 and Comparative Example 3 were evaluated for production, and the results showed that the production processing time of the aircraft skin connection structure in Example 1 was about 3 h; the production processing time of the aircraft skin connection structure in Comparative Example 3 was about 6.3 h. It can be seen that the scheme in the embodiments significantly improves the production efficiency of the aircraft thermoplastic composite skin connection structure while ensuring structural stability and safety.

Claims

1. An aircraft skin structure, characterized in that: The invention comprises an aircraft skin and a reinforcing rib plate of carbon fiber reinforced thermoplastic composite materials, wherein the aircraft skin and the reinforcing rib plate are connected together by a connecting interface layer, wherein the connecting interface layer is dispersed with conductive particles, wherein the conductive particles adopt a mixed system of carbon nanotubes and silver nanowires, wherein the weight ratio of the carbon nanotubes to the silver nanowires is 2.5-3.5:1, and part of the thermoplastic material of the connecting interface layer is embedded in the inner cavity of the carbon nanotubes.

2. The aircraft skin structure according to claim 1, characterized in that: The diameter of the carbon nanotube is not less than 100 nm, the diameter of the silver nanowire is not more than 55 nm, and part of the silver nanowire extends into the inner cavity of the carbon nanotube.

3. The aircraft skin structure according to claim 2, characterized in that: The length of silver nanowires is 0.8 to 3 times that of carbon nanotubes.

4. The method for manufacturing an aircraft skin structure according to any one of claims 1 to 3, characterized in that the steps include: Step 1, preparing a hot pressing mold, an aircraft skin and a stiffening rib plate to be connected, wherein the working surface of the hot pressing mold has a forming structure adapted to the grid portion; the aircraft skin and the stiffening rib plate are both made by a hot pressing forming process; Step 2, laying conductive particles on the surface of the connection area of ​​the aircraft skin; Step 3, placing the lower die working surface of the hot pressing die in step 1 facing the connection area surface of the aircraft skin; Step 4, setting the hot pressing process parameters and starting the hot pressing equipment to perform hot pressing, so that the aircraft skin softens under the action of heat and forms a grid structure. During this process, the conductive particles are embedded in the grid structure and form energy conducting ribs; Step 5, clamping the aircraft skin and the stiffener plate, wherein the connection area of ​​the aircraft skin is located below the stiffener plate, and the energy guiding ribs on the aircraft skin are attached to the lower surface of the stiffener plate; Step 6: Use ultrasonic welding equipment to weld the aircraft skin and the reinforcing rib plate together. During the welding process, the energy guiding ribs melt and form a connecting interface layer with the reinforcing rib plate, and then cool and solidify into shape.

5. The manufacturing method according to claim 4, characterized in that: The pre-laying thickness of the conductive particles in step 2 is 10-30 μm.

6. The manufacturing method according to claim 4, characterized in that: During the welding process, the downward pressing height of the welding tool head is not greater than half the thickness of the reinforcing rib plate, the welding amplitude is 25μm~45μm, the welding pressure is 0.1MPa~0.7MPa, and the welding time is 0.5s~3s.

7. The manufacturing method according to claim 4, characterized in that: In step 5, the aircraft skin is mounted on a cutting board in which an induction coil is embedded; the welding current during the welding process is 25-32A.

8. The manufacturing method according to any one of claims 5 to 7, characterized in that: The energy conducting rib is provided with a plurality of vertically arranged grooves, the grooves are close to the edge of the energy conducting rib, and the depth of the grooves is equal to the height of the energy conducting rib.

9. The manufacturing method according to claim 8, characterized in that: During the welding process, a magnetic field is applied synchronously to activate the eddy current effect of the conductive particles to assist in heating.

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