An aircraft skin structure and its manufacturing method

Abstract

The present invention relates to the technical field of aircraft skins, and specifically discloses an aircraft skin structure and a manufacturing method thereof. The aircraft skin structure includes an aircraft skin made of a carbon fiber reinforced thermoplastic composite material and a reinforcing rib plate. The aircraft skin and the reinforcing rib plate are connected together through a connection interface layer. Conductive particles are dispersed in the connection interface layer. The conductive particles adopt a mixed system of carbon nanotubes and silver nanowires. The weight ratio of carbon nanotubes to silver nanowires is 2.5 to 3.5:1, and a part of the thermoplastic material of the connection interface layer is embedded in the inner cavity of the carbon nanotubes. The aircraft skin and the reinforcing rib plate are welded and clamped, and an ultrasonic welding device is used to weld the aircraft skin and the reinforcing rib plate together. The present invention significantly improves the mechanical properties of the aircraft skin structure, improves the preparation and forming efficiency of the aircraft skin structure, is convenient to operate, and is suitable for engineering applications.

Description

Technical Field

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

[0002] At present, thermoplastic composites have been used as main components such as wing and fuselage skins on aircraft at home and abroad. They have advantages such as high specific strength, good corrosion resistance, strong energy absorption and impact resistance, and easy recyclability. The corresponding connection structure requirements are increasing. High-performance thermoplastic composites have been applied to aircraft structures, such as components of helicopter fuselages, wings, tails, spoilers, etc. Such thermoplastic composites can gradually replace traditional metal components as structural or functional components.

[0003] The method for manufacturing an aircraft skin structure using thermoplastic composites is mainly the hot pressing process, and structures such as stiffeners are connected to the skin in a certain way. The connection methods include adhesive bonding structures or mechanical connection structures. The adhesive bonding structure has problems such as long curing time and easy environmental influence on the adhesive layer of the finished product in actual production. The mechanical connection structure is prone to weight gain of the structure, which does not meet the lightweight purpose because rivets, bolts and other structural parts need to be introduced, and stress concentration is likely to occur at the connection. Therefore, it is necessary to develop a new type of aircraft skin structure and manufacturing method.

[0004] The existing literature CN119550651A discloses a fiber continuously longitudinally and transversely stiffened integral composite skin structure, in which the skin and the longitudinal and transverse stiffening structures are co-bonded and formed to achieve lightweight and strength improvement. However, this method is complex and time-consuming in processing, and has weak applicability to complex structures; the literature CN108973089A discloses a preparation method for a composite special-shaped skin structure. According to the digital model of the composite special-shaped skin structure, a combined process cover plate matching the forming surface is prepared, and the blank of the composite special-shaped skin structure, the composite process thin layer and the combined process cover plate are stacked on the mold tooling from bottom to top in sequence to form a process combination. However, the preparation process of this hot pressing and curing method is complex; the literature CN111823601 discloses a method for reducing the ultrasonic welding difficulty of engineering plastics by using chemical time. It uses a welding auxiliary reagent to be pre-coated on the welding surface of the engineering plastics, so that the surface of the engineering plastics is pre-dissolved and activated before ultrasonic welding. However, this method has the problem that the selection of the auxiliary reagent is relatively complex, and the surface treatment has limited improvement on the mechanical properties of the joint. In short, the preparation of the thermoplastic composite aircraft skin connection structure currently has problems such as complex preparation, low connection efficiency, and poor connection stability and safety between structural components. Therefore, one of the main problems to be solved is to improve the interface safety of the aircraft skin connection structure while simplifying the preparation process. Summary of the Invention

[0005] Aiming at the problems mentioned in the background art, the object of the present invention is to provide an aircraft skin structure and a manufacturing method thereof.

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

[0007] An aircraft skin structure includes an aircraft skin made of carbon fiber reinforced thermoplastic composite material and a stiffening rib plate. The aircraft skin and the stiffening rib plate are connected together through a connection interface layer. The connection interface layer is dispersed with conductive particles. The conductive particles adopt a mixed system of carbon nanotubes and silver nanowires. The weight ratio of carbon nanotubes to silver nanowires is 2.5 - 3.5:1, and a part of the thermoplastic material of the connection 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 100 nm, the diameter of the silver nanowires is not more than 55 nm, and a part of the silver nanowires extend into or are inserted into the inner cavity of the carbon nanotubes.

[0009] As a preferred solution, the length of the silver nanowires is 0.8 - 3 times the length of the carbon nanotubes.

[0010] Furthermore, the steps of the manufacturing method of the aircraft skin structure of the present invention include:

[0011] Step 1: Prepare a hot pressing mold, the aircraft skin and the stiffening rib plate to be connected. The working surface of the hot pressing mold has a forming structure adapted to the grid part; both the aircraft skin and the stiffening rib plate are made by a hot pressing forming process;

[0012] Step 2: Lay conductive particles on the surface of the connection area of the aircraft skin;

[0013] Step 3: Align the working surface of the lower mold of the hot pressing mold in Step 1 with the surface of the connection area of the aircraft skin;

[0014] Step 4: Set the hot pressing process parameters and start the hot pressing equipment for hot pressing, so that the contact surface between the aircraft skin and the hot pressing mold softens under the action of heat and forms a grid part structure. During this process, the conductive particles are embedded in the grid part structure and form an energy - conducting rib;

[0015] Step 5: Clamp the aircraft skin and the stiffening rib plate for welding. The connection area of the aircraft skin is located below the stiffening rib plate, and the energy - conducting rib on the aircraft skin abuts against the lower surface of the stiffening rib plate;

[0016] Step 6: Use an ultrasonic welding device to weld the aircraft skin and the stiffening rib plate together. During the welding process, the energy - conducting rib melts and jointly forms a connection interface layer with the stiffening rib plate, and then cools and solidifies to form a shape.

[0017] As a preferred solution, the pre - laid thickness of the conductive particles in Step 2 is 10 - 30 μm.

[0018] As a preferred solution, during the welding process, the pressing height of the welding tool head is not greater than half of the thickness of the stiffening rib plate, the welding amplitude is 25μm - 45μm, the welding pressure is 0.1MPa - 0.7MPa, and the welding time is 0.5s - 3s.

[0019] Further, in step 5, the aircraft skin is installed on the anvil, and an induction coil is buried in the anvil; the welding current during the welding process is 25 - 32A.

[0020] Further, there are several vertically arranged grooves on the energy guiding rib, the grooves are close to the edge of the energy guiding rib, the depth of the grooves is equal to the height of the energy guiding rib, and the cross-section of the grooves is rectangular.

[0021] Further, during the welding process, a magnetic field is synchronously applied to activate the eddy current effect of the conductive particles for auxiliary heating.

[0022] In the present invention, both the aircraft skin and the stiffening rib plate are made by a hot pressing forming process, and the steps include:

[0023] Step 11, prepare carbon fiber woven fabric, thermoplastic resin particle material and a hot pressing forming mold, and the hot pressing forming mold is designed to have a shape adapted to the skin structure;

[0024] Step 12, lay the carbon fiber woven fabric in the hot pressing forming mold; the resin particles are heated and melted by an extruder, enter the hot pressing forming mold after the viscosity is reduced by screw shearing, and the resin and the carbon fiber are fully infiltrated by applying pressure through the hot pressing forming mold, and a continuous carbon fiber reinforced thermoplastic composite layer is formed after cooling.

[0025] Step 13, stack and assemble multiple layers of continuous carbon fiber reinforced thermoplastic composite layers in the hot pressing forming mold, set the hot pressing process parameters, and apply heat and pressure through the mold to complete the connection between the continuous carbon fiber reinforced thermoplastic composite layers.

[0026] Step 14, using the continuous carbon fiber reinforced thermoplastic composite layer obtained in step 13 as the bottom layer, repeat step 13 multiple times until an aircraft skin with the target size is formed, and the target size is determined by those skilled in the art according to the actual product model.

[0027] Referring to the aforementioned steps 11 to 14, select a mold adapted to the stiffening rib plate to prepare the stiffening rib plate.

[0028] 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; during the forming process of the continuous carbon fiber reinforced thermoplastic composite layer (i.e., in step 13), the hot pressing temperature is the resin melting temperature ±10 °C, and the pressurization stage is divided into multiple levels of pressurization (initially 0.5-2 MPa, and then increased to 2.0-4 MPa), and the holding time is 0.5-1 h; in step 6, the solder joint spacing is 1±0.3 cm, and the boundary distance is 5 mm~10 mm.

[0029] After step 6, the obtained aircraft skin connection structure (which can also be called an aircraft skin connection component) is integrally placed in a heat treatment furnace for stress relief, and then machining treatment and surface roughness grinding are carried out to form an aircraft skin connection structure of thermoplastic composite material.

[0030] Beneficial effects: The present invention comprehensively utilizes the manufacturing methods of energy guiding rib hot pressing forming and ultrasonic welding, realizes the dual optimization of the aircraft skin structure and materials and the modification of the base material, significantly improves the mechanical properties of the aircraft skin structure, greatly improves the energy utilization rate and forming efficiency of the connection interface, improves the preparation and forming efficiency of the aircraft skin structure, is easy to operate, and is suitable for engineering applications. Description of the Drawings

[0031] Figure 1 It is a schematic diagram (side view) of the energy guiding rib forming process in Example 1;

[0032] Figure 2 It is a schematic diagram (three-dimensional view) of the energy guiding rib forming process in Example 1;

[0033] Figure 3 It is a schematic diagram (side view) of the ultrasonic welding of the aircraft skin structure in Example 1;

[0034] Figure 4 It is a schematic diagram (three-dimensional view) of the ultrasonic welding of the aircraft skin structure in Example 1;

[0035] Figure 5 It is a schematic diagram of the energy guiding rib forming process in Example 2;

[0036] Figure 6 It is a temperature-time curve diagram of the hot pressing die in Example 1.

[0037] Description of the drawing numbers: 1 - upper die of the hot pressing die, 2 - aircraft skin, 3 - lower die, 4 - temperature control heating element, 5 - welding tool head, 6 - stiffening rib plate, 7 - ultrasonic solder joint, 8 - anvil, 9 - induction coil, 10 - energy guiding rib, 11 - groove, 12 - first groove, 13 - second groove. Detailed Embodiments

[0038] The technical solution of the present invention will be clearly and completely described below in conjunction with the accompanying drawings. Obviously, the described embodiments are only a part of the embodiments of the present invention, rather than all of the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by those of ordinary skill in the art without creative efforts shall fall within the protection scope of the present invention. In the embodiments, the reference numerals in the drawings only show partial schematics of the corresponding components. Embodiment 1

[0039] Combined with Figures 1 to 4 As shown, an aircraft skin structure includes an aircraft skin 2 made of carbon fiber reinforced thermoplastic composite material and a stiffening rib plate 6. The aircraft skin 2 and the stiffening rib plate 6 are connected together through a connection interface layer, and conductive particles are dispersed in the connection interface layer. The conductive particles adopt a mixed system of carbon nanotubes and silver nanowires, and a part of the thermoplastic material of the connection interface layer is embedded in the inner cavity of the carbon nanotubes. The manufacturing method of the aircraft skin structure includes the following steps:

[0040] Step 1: Prepare a hot pressing mold, the aircraft skin 2 and the stiffening rib plate 6 to be connected (welded). The working surface of the hot pressing mold has a forming structure adapted to the grid part (the grid part in the present invention refers to a grid-like structure composed of a number of protrusions and grooves. After the conductive particles are embedded in the protrusions, the energy-conducting ribs 10 are formed). Both the aircraft skin 2 and the stiffening rib plate 6 are made by the hot pressing forming process. Among them, the thicknesses of the aircraft skin 2 and the stiffening rib plate 6 to be welded are both 3 mm.

[0041] Step 2: Lay conductive particles on the surface of the connection area of the aircraft skin 2 to be welded, and the pre-laying thickness is 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, and the diameter of the silver nanowires is 55 nm. A 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.

[0042] Step 3: Place the lower mold 3 of the hot pressing mold in place. A temperature control heating element 4 is arranged on the lower mold 3, 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 faces the surface of the connection area of the aircraft skin 2 to be welded.

[0043] Step 4: Set the hot pressing process parameters and start the hot pressing equipment for hot pressing, so that the aircraft skin 2 to be welded softens under the thermal action and forms a grid part structure. During this process, the conductive particles are embedded in the grid part structure and form the energy-conducting ribs 10, and the aircraft skin 2 with the energy-conducting ribs 10 is obtained. Combined with Figure 6As shown, the heating rate during hot pressing is 20 °C / min. It is heated to 10 °C above the melting temperature Tm of the thermoplastic composite board and then held for 10 min. The cooling rate is used as a control variable. The original is 30 °C / min, and on this basis, every increase of 10 °C / min is taken as a gradient to control the crystallinity of the material; the downward pressure parameter of the upper die 1 of the hot pressing die is controlled at 0.4 MPa, and no pressure is applied when starting to cool. After the cooling is completed, the energy guiding rib 10 is obtained. The side length of the energy guiding rib 10 is 1 mm and the height is 0.2 mm, and the distance between adjacent energy guiding ribs 10 is 0.2 mm;

[0044] Step 5, as shown in Figure 3 and Figure 4 After that, the aircraft skin 2 with the energy guiding rib 10 and the stiffening rib plate 6 obtained are welded and clamped. The aircraft skin 2 with the energy guiding rib 10 is located below the stiffening rib plate 6. The energy guiding rib 10 on the aircraft skin 2 abuts against the lower surface of the stiffening rib plate 6. The aircraft skin 2 is installed on the anvil 8, and an induction coil 9 is buried in the anvil 8;

[0045] Step 6, the aircraft skin 2 with the energy guiding rib 10 and the stiffening rib plate 6 are welded together by an ultrasonic welding device. The buried induction coil 9 is started synchronously with the ultrasonic welding device. 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 guiding rib 10 melts and jointly forms a connection interface layer with the stiffening rib plate 6, and then cools and solidifies (the cooling and solidification temperature gradient is 10 °C / min, and the air humidity under this room temperature condition is 20 - 30%), obtaining the aircraft skin structure; during the welding process, the downward pressing height of the welding tool head 5 is not greater than half of the thickness of the stiffening rib plate 6, the welding amplitude is 35 μm, the welding pressure is 0.2 MPa, the welding time is 1 s, and the welding current is 30 A; the distance between adjacent ultrasonic welding points 7 is 1 cm, and the boundary distance is 6 mm.

[0046] After Step 6, the obtained aircraft skin connection structure as a whole is put into a heat treatment furnace for stress relief, and then machining treatment and surface roughness grinding are carried out to form the aircraft skin connection structure of the thermoplastic composite material.

[0047] During the welding process, under the action of longitudinal ultrasonic vibration, the energy guiding rib 10 at the interface of the aircraft skin 2 with the energy guiding rib 10 generates stress concentration and preferentially deforms and heats up; at different cooling rates, the material near the energy guiding rib 10 has a low crystallinity and is easy to deform and heat up; while increasing the frictional heat generation in the early stage of welding, the conductive particles generate heat under the eddy current effect of the induced magnetic field to promote the interface temperature rise. After the interface material is heated to the melting temperature, it diffuses and spreads with each other. The melted part of the material on the lower surface of the upper and middle parts of the energy guiding rib 10 and the strengthening rib plate 6 forms a solder joint, and the melted part of the material on the lower surface of the strengthening rib plate 6 spreads into the grid grooves corresponding to the unfused part of the energy guiding rib 10 to form a mechanical interlocking structure. In this solution, the dual control of the energy guiding rib structure and material is cleverly realized, so that the ultrasonic welding energy is concentrated at the interface, improving the connection strength and stability of the aircraft skin connection structure, as well as the welding efficiency and quality, and greatly improving the energy utilization rate of the ultrasonic welding interface. Example 2

[0048] A manufacturing method of an aircraft skin structure, referring to Example 1, the main differences from Example 1 mainly include: Combining Figure 5 As shown, there are several vertically arranged grooves 11 on the energy guiding rib 10. The grooves 11 are close to the edge of the energy guiding rib 10. The height of the grooves 11 is equal to the height of the energy guiding rib 10. The width of the grooves 11 is 0.1 mm, and the distance between adjacent grooves 11 is 0.1 mm. The two grooves 11 facing each other on two adjacent energy guiding ribs 10 are staggered from each other (for example: Figure 5 In the first groove 12 and the second groove 13 on two different energy guiding ribs 10 are staggered from each other. The first groove 12 and the second groove 13 can be collectively referred to as the groove 11); the side length of the energy guiding rib 10 is 2 mm and the height is 0.3 mm, and the distance between adjacent energy guiding ribs 10 is 0.2 mm. In this embodiment, the further advantages include: the melted part of the material on the lower surface of the strengthening rib plate 6 spreads into the grid grooves and the grooves 11 corresponding to the unfused part of the energy guiding rib 10, and grows branch feet from the grooves 11 to form a mechanical structure similar to "chip pins" and "interpenetration (the branch feet grown in the grooves 11 facing each other interpenetrate, such as the branch feet grown in multiple first grooves 12 and multiple second grooves 13 interpenetrate)", and its stability and mechanical properties are further optimized, especially the shear resistance is significantly improved.

[0049] In the embodiment, the steps for both the aircraft skin and the strengthening rib plate to be made by the hot pressing forming process include:

[0050] Step 11, prepare carbon fiber woven fabric, thermoplastic resin particle material and a hot pressing forming mold, and the hot pressing forming mold is designed to have a shape adapted to the skin structure;

[0051] Step 12: Lay the carbon fiber woven fabric in the hot pressing mold. Heat and melt the resin particles by an extruder, reduce the viscosity through screw shearing, then introduce them into the hot pressing mold. Apply pressure through the hot pressing mold to fully infiltrate the resin and the carbon fiber. After cooling, a continuous carbon fiber reinforced thermoplastic composite layer is formed.

[0052] Step 13: Stack and assemble multiple layers of the continuous carbon fiber reinforced thermoplastic composite layer in the hot pressing mold. Set the hot pressing process parameters, apply heat and pressure through the mold to complete the connection between the continuous carbon fiber reinforced thermoplastic composite layers.

[0053] The fiber volume content in the continuous carbon fiber reinforced thermoplastic composite layer accounts for 40%, and the single-layer thickness is 0.4 mm. During the forming process of the continuous carbon fiber reinforced thermoplastic composite layer (i.e., in Step 13), the hot pressing temperature is equal to the resin melting temperature. The pressurization stage is divided into multiple levels of pressurization (initially 1.5 MPa, then increased to 3 MPa), and the pressure holding time is 0.8 h.

[0054] Step 14: Use the continuous carbon fiber reinforced thermoplastic composite layer obtained in Step 13 as the bottom layer, and repeat Step 13 multiple times until an aircraft skin with the target size is formed.

[0055] Referring to the aforementioned Steps 11 to 14, select a mold suitable for the stiffener plate and prepare the stiffener plate.

[0056] In Example 1 and Example 2, conductive particles are dispersed in the connection interface layer between the aircraft skin 2 and the stiffener plate 6 with the energy guiding ribs 10. Moreover, part of the material (resin) of the thermoplastic composite plate is embedded in the inner cavity of the carbon nanotubes, and part of the silver nanowires extend (insert) into the inner cavity of the carbon nanotubes. This structure in which the resin, carbon nanotubes, and silver nanowires are mutually interlocked significantly improves the mechanical properties of the connection interface layer.

[0057] Comparative Example 1: Refer to Example 1. The difference from Example 1 is that the conductive particles and their corresponding hot pressing steps are omitted.

[0058] Comparative Example 2: Refer to Example 2. The difference from Example 2 is that the conductive particles and their corresponding hot pressing steps are omitted.

[0059] The mechanical properties of the aircraft skin structure (connection structure) obtained in the examples and comparative examples were tested. The results showed that: in Example 1, the strength of the connection structure was 29.7 MPa, the joint shear force was 18,540 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 21,140 N; in Comparative Example 1, the strength of the connection structure was 25.3 MPa, and the joint shear force was 15,811 N; in Comparative Example 2, the strength of the connection structure was 26.98 MPa, and the joint shear force was 16,289 N.

[0060] From the interface temperature curve during the welding process, it can be seen that the effective connection time (Tm < T < Td) in the examples was increased compared with the comparative examples, the welding interface temperature was mainly distributed near the interface, and the corresponding energy utilization rate in the examples was higher. Defect detection (welding joint detection) was carried out on the connection structure in the examples, and the results showed that there were no defects such as pores.

[0061] Comparative Example 3, the component structure refers to Example 1, and the difference between it and Example 1 is that the connection method between the aircraft skin and the stiffener is adhesive bonding.

[0062] Production evaluation was carried out on the aircraft skin connection structures of the same specifications in Example 1 and Comparative Example 3. The results showed that: the production and processing time of the aircraft skin connection structure in Example 1 was about 3 h; the production and processing time of the aircraft skin connection structure in Comparative Example 3 was about 6.3 h; it can be seen that the solution in the examples significantly improved the production efficiency of the aircraft thermoplastic composite skin connection structure while ensuring the structural stability and safety.

Claims

1. A manufacturing method of an aircraft skin structure, characterized in that: The aircraft skin structure includes an aircraft skin made of carbon fiber reinforced thermoplastic composite material and a reinforcing rib plate. The aircraft skin and the reinforcing rib plate are connected together through a connection interface layer. The connection interface layer is dispersed with conductive particles. The conductive particles adopt a mixed system of carbon nanotubes and silver nanowires. The weight ratio of carbon nanotubes to silver nanowires is 2.5 to 3.5:1, and part of the thermoplastic material of the connection interface layer is embedded in the inner cavity of the carbon nanotubes; The steps of the manufacturing method include: Step 1, prepare a hot pressing mold, the aircraft skin and the reinforcing rib plate to be connected. The working surface of the hot pressing mold has a forming structure adapted to the grid part; both the aircraft skin and the reinforcing rib plate are made by a hot pressing forming process; Step 2, lay conductive particles on the surface of the connection area of the aircraft skin; Step 3, make the working surface of the lower mold of the hot pressing mold in Step 1 face the surface of the connection area of the aircraft skin; Step 4, set the hot pressing process parameters and start the hot pressing equipment for hot pressing, so that the aircraft skin softens under the action of heat and forms a grid part structure. During this process, the conductive particles are embedded in the grid part structure and form an energy conducting rib; Step 5, clamp the aircraft skin and the reinforcing rib plate. The connection area of the aircraft skin is located below the reinforcing rib plate, and the energy conducting rib on the aircraft skin abuts against the lower surface of the reinforcing rib plate; Step 6, use an ultrasonic welding device to weld the aircraft skin and the reinforcing rib plate together. During the welding process, the energy conducting rib melts and jointly forms a connection interface layer with the reinforcing rib plate, and then cools and solidifies to form a shape.

2. The manufacturing method according to claim 1, wherein: The diameter of the carbon nanotubes is not less than 100 nm, the diameter of the silver nanowires is not greater than 55 nm, and part of the silver nanowires extend into the inner cavity of the carbon nanotubes.

3. The manufacturing method according to claim 2, characterized in that: The length of the silver nanowires is 0.8 to 3 times the length of the carbon nanotubes.

4. The manufacturing method according to claim 3, characterized in that: In Step 2, the pre-laying thickness of the conductive particles is 10 to 30 μm.

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

6. The manufacturing method according to claim 5, characterized in that: In Step 5, the aircraft skin is installed on an anvil, and an induction coil is buried in the anvil; the welding current during the welding process is 25 to 32 A.

7. The manufacturing method according to any one of claims 1-6, characterized in that: The energy conducting rib has a plurality of vertically arranged grooves. The grooves are close to the edge of the energy conducting rib, and the groove depth is equal to the height of the energy conducting rib.

8. The manufacturing method according to claim 7, characterized in that: During the welding process, a magnetic field is synchronously applied to activate the eddy current effect of the conductive particles for auxiliary heating.

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