A composite material main load-bearing beam and integral forming method

Abstract

The present invention relates to a composite material main load-bearing beam and an integral forming method. The composite material main load-bearing beam comprises a main joint, a wing box tap, a wing box reinforcement sheet, a weight-reducing core layer and a skin, wherein the main joint is fixedly connected to the wing box tap and the wing box reinforcement sheet, and the weight-reducing core layer is fixedly connected to the wing box tap; the skin is prepared by coating the main joint, the wing box tap, the wing box reinforcement sheet and the weight-reducing core layer with fiber prepreg according to the load condition of the wing surface, and adopting an integral compression molding process, wherein the main joint, the wing box tap and the wing box reinforcement sheet are metal parts. The present invention can fill carbon fiber inside the metal joint, and the filled carbon fiber and the outer skin are combined to form a plurality of T-beam structures, which effectively improves the rigidity of the main load-bearing beam; the present invention positions and combines the return beam composed of a plurality of weight-reducing core layers and carbon fibers with metal parts in various forms such as docking, gluing, and screwing, and the overall curing molding further improves the reliability of the structure.

Description

Technical Field

[0001] The invention relates to a composite material main bearing beam and an integral forming method thereof, belonging to the technical field of composite materials. Background Art

[0002] Aircraft weight is closely related to performance and economy, and reducing the weight of aircraft structures is one of the main goals in research and development. Composite wing surfaces have excellent mechanical properties such as high specific stiffness and specific strength, good fatigue resistance and corrosion resistance, and have been widely used in aviation and aerospace structures. Their application range has expanded from the initial secondary load-bearing structure to the current main load-bearing structure.

[0003] The main function of the main load-bearing wing surface is to generate lift. The main load-bearing beam inside the wing surface plays a decisive role in the rigidity of the wing surface. The quality of the main load-bearing beam directly affects the stability and reliability of the aircraft during flight. How to prepare structurally reliable composite main load-bearing beam components has become a hot research topic. Summary of the invention

[0004] The purpose of the present invention is to overcome the shortcomings of the existing process technology and provide a lightweight and high-strength composite main load-bearing beam and an integral forming method of the load-bearing beam. The load-bearing beam is composed of a main joint, a wing box branch joint, a wing box reinforcement sheet, a weight-reducing core layer and a composite material skin, and adopts an integral compression molding process.

[0005] The technical solution adopted by the present invention is as follows:

[0006] In a first aspect, the present invention provides a composite main load-bearing beam, comprising a main joint, a wing box tap, a wing box reinforcement sheet, a weight-reducing core layer and a skin, wherein the main joint is fixedly connected to the wing box tap and the wing box reinforcement sheet, and the weight-reducing core layer is fixedly connected to the wing box tap; the skin is coated on the outside of the main joint, the wing box tap, the wing box reinforcement sheet and the weight-reducing core layer.

[0007] Preferably, the skin is formed by coating the main joint, the wing box branch joint, the wing box reinforcement sheet and the weight-reducing core layer with fiber prepreg according to the load conditions of the wing surface, and then molding and curing the whole.

[0008] Preferably, the fiber prepreg is a carbon fiber prepreg, which comprises carbon fiber and resin, wherein the carbon fiber is unidirectional T300, T700, T800 or T1000 carbon fiber, and the resin is medium-temperature or high-temperature curing epoxy resin. The fiber may also be other applicable fibers, such as glass fiber or quartz fiber.

[0009] Preferably, the main joint, the wing box branch joint and the wing box reinforcement sheet are metal parts; and the weight-reducing core layer is rigid polyurethane foam, phenolic resin foam or PMI foam.

[0010] Preferably, the wing box tap is installed in the corresponding groove of the main joint, the wing box reinforcement sheet is installed in the groove reserved for the wing box tap and the weight reduction core layer, and one end of the weight reduction core layer is installed in the groove reserved for the wing box tap.

[0011] In a second aspect, the present invention provides a method for integrally forming a composite material main load-bearing beam, comprising the following steps:

[0012] Assemble the wing box tap and the weight-reducing core layer, and install one end of the weight-reducing core layer in the groove reserved in the wing box tap;

[0013] Assemble the wing box reinforcement sheet with the wing box tap and the weight-reducing core layer, and install the wing box reinforcement sheet in the grooves reserved in the wing box tap and the weight-reducing core layer;

[0014] The wing box tap, wing box reinforcement sheet and weight-reducing core layer assembly are coated with fiber prepreg and pre-cured;

[0015] Assembling the pre-cured wing box tap joint, wing box reinforcement sheet and weight-reducing core layer assembly with the main joint;

[0016] For the assemblies of main joints, wing box branch joints, wing box reinforcement sheets and weight-reducing core layers, fiber prepregs of different thicknesses are coated at different positions according to the load conditions of the wing surface, and the whole is molded and cured.

[0017] Preferably, before the main joint, the wing box tap and the wing box reinforcement sheet are assembled, the main joint, the wing box tap and the wing box reinforcement sheet are sandblasted; the sandblasting uses quartz sand of 20-40 mesh and the air pressure is 0.4MPa-0.7MPa.

[0018] Preferably, before the weight-reducing core layer is assembled with the wing box tap, the foam of the weight-reducing core layer is dried; the drying process is carried out in an oven at 130°C±5°C for 3 hours.

[0019] Preferably, the pre-curing treatment includes: wrapping the assembly coated with prepreg with release cloth and breathable felt in turn, and placing it in a vacuum bag; placing the vacuum bag horizontally in an oven, tearing off the surface protective film of the vacuum bag sealing strip, inserting a vacuum joint, and sealing the vacuum bag; turning on the vacuum pump to ensure that the vacuum degree in the vacuum bag is less than -0.095MPa, setting the oven temperature to 35°C, starting the timing when the oven reaches 35°C, keeping the temperature for 30 minutes, turning off the oven power, and taking out the assembly.

[0020] Optionally, coating fiber prepregs with different thicknesses at different positions according to the load conditions of the airfoil includes:

[0021] Overall cladding of the 7mm area on the upper and lower wing surfaces: with the leading edge of the main load-bearing beam as the 0° direction, cladding prepreg in the directions of [+45 / 0 / 0 / -45], [0 / 0 / 0 / 90], [90 / 0 / 0 / 0], and [-45 / 0 / 0 / +45] in sequence. The thickness of the prepreg is 7mm.

[0022] Overall cladding of the 5mm area on the upper and lower wing surfaces: with the leading edge of the main load-bearing beam as the 0° direction, cladding prepreg in the directions of [+45 / 0 / 0 / -45] and [-45 / 0 / 0 / +45] in sequence, with a thickness of 5mm;

[0023] Overall cladding of the 4mm area of ​​the upper and lower wing surfaces and the joint area: with the leading edge of the main load-bearing beam as the 0° direction, cladding of prepreg in the [0 / 0 / 0 / 90] and [90 / 0 / 0 / 0] directions in sequence, with a prepreg thickness of 4mm;

[0024] Overall covering the 3mm area of ​​the upper and lower wing surfaces and the joint area: taking the leading edge of the main load-bearing beam as the 0° direction, cover with prepreg in the directions of [+45 / 0 / 0 / -45][0 / 0 / 0 / 90][+45 / 0 / 0 / -45][-45 / 0 / 0 / +45][90 / 0 / 0 / 0][-45 / 0 / 0 / +45] in sequence. The thickness of the prepreg is 3mm.

[0025] The beneficial effects of the present invention compared with the prior art are as follows:

[0026] 1) A composite wing surface laying method with a main joint and a wing box branch joint structure is designed. The laying scheme can fill carbon fiber inside the metal joint. The filled carbon fiber and the outer skin are combined to form multiple T-beam structures (the internal carbon fiber is in the Z direction, i.e., the vertical direction, and the skin is in the XY direction, i.e., the horizontal direction, and the two are stacked together to form a T-shaped structure), which effectively improves the rigidity of the main load-bearing beam;

[0027] 2) The present invention positions and combines the return beam composed of multiple weight-reducing core layers and carbon fibers with metal parts through various forms such as butt joint, adhesive joint, and screw joint, and solidifies and forms the entire beam, thereby further improving the reliability of the structure. BRIEF DESCRIPTION OF THE DRAWINGS

[0028] Figure 1 It is a top view of the composite main load-bearing beam.

[0029] Figure 2 It is a bottom view of the composite main load-bearing beam.

[0030] Figure 3 This is a top view of the main joint.

[0031] Figure 4 This is a bottom view of the main joint.

[0032] Figure 5 is a top view of the wing box tap.

[0033] Figure 6 is a side view of the wing box tap.

[0034] Figure 7 It is a schematic diagram of the wing box reinforcement.

[0035] Figure 8 It is a top view of the metal parts consisting of the main joint, wing box branch joint, and wing box reinforcement plate.

[0036] Fig. 9 It is a bottom view of the metal parts consisting of the main joint, wing box branch joint, and wing box reinforcement plate.

[0037] Fig.10 It is a top view of the “positive mold” of the layup.

[0038] Fig.11 This is a top view of the laminate "male mold".

[0039] Fig.12 It is a top view of the weight-reducing core layer.

[0040] Fig.13 It is a bottom view of the weight-reducing core layer.

[0041] 1-main joint; 2-wing box tap; 3-wing box reinforcement plate; 4-leading edge web; 5-trailing edge web; 6-foam; 7-web; 8-joint area; 9-7mm area; 10-5mm area; 11-4mm area; 12-3mm area. DETAILED DESCRIPTION

[0042] In order to make the above-mentioned objects, features and advantages of the present invention more obvious and easy to understand, the present invention is further described in detail below through specific embodiments and drawings.

[0043] The composite material main load-bearing beam of the present invention is as follows Figure 1 , Figure 2 As shown, it includes main joints, wing box taps, wing box reinforcement sheets and weight-reducing core layers, and is covered with skin ( Figure 1 , Figure 2 The skin is not shown in the figure). The main joint is fixedly connected to the wing box branch joint and the wing box reinforcement sheet, and the weight-reducing core layer is fixedly connected to the wing box branch joint. The skin is formed by coating the fiber prepreg on the outside of the main joint, the wing box branch joint, the wing box reinforcement sheet and the weight-reducing core layer according to the load condition of the wing surface, and the whole is molded and cured.

[0044] The weight-reducing core layer is made of foam material, such as rigid polyurethane foam, phenolic resin foam or PMI foam.

[0045] The composite main load-bearing beam may also include a leading edge web and a trailing edge web. Figure 1 , Figure 2 As shown, it is located at the leading and trailing edges of the main load-bearing beam and is used to transfer the Z-direction (vertical direction) load. Its material is carbon fiber resin-based composite materials and other materials. A web can also be set inside the main load-bearing beam, that is, a panel located between the foam of the weight-reducing core layer that connects the upper and lower wing surfaces of the composite material (such as Fig.10 shown).

[0046] The main joint, wing box branch joint and wing box reinforcement sheet are metal parts, and their materials are metal materials such as titanium alloy or stainless steel. The main joint is used to connect the main load-bearing beam and the folding and unfolding mechanism of the aircraft.

[0047] In one embodiment of the present invention, the main connector is as follows Figure 3 , Figure 4 As shown, the wing box tap is as Figure 5 , Figure 6 As shown, the wing box reinforcement is as follows Figure 7 shown. Figure 8 , Fig. 9 The structural diagram of a metal part composed of a main joint, a wing box tap and a wing box reinforcement sheet. The wing box tap is installed in the corresponding groove of the main joint and is fixedly connected to the main joint by screws or the like. The wing box reinforcement sheet is arranged in the groove reserved in the wing box tap. The wing box tap is provided with a groove for inserting one end of the weight-reducing core layer. In this embodiment, there are 4 wing box taps, and other numbers of wing box taps may also be arranged in other embodiments.

[0048] The integral forming method of the composite material main load-bearing beam of the present invention comprises firstly combining the wing box branch joint, the wing box reinforcement sheet and the weight-reducing core layer, and then coating the carbon fiber prepreg, and then combining the main joint into a laminate "positive mold" after pre-curing treatment. Fig.10 , Fig.11 As shown, carbon fiber prepregs of different thicknesses are then coated at different positions of the "male mold" according to the load conditions of the wing surface, and finally the entire mold is compression molded and cured.

[0049] The carbon fiber prepreg comprises carbon fiber and resin. The carbon fiber adopts unidirectional T300, T700, T800 or T1000 carbon fiber, and the resin adopts medium-temperature or high-temperature curing epoxy resin.

[0050] In one embodiment of the present invention, a method for integrally forming a composite material main load-bearing beam is provided, comprising the following steps:

[0051] Assemble the wing box tap and the weight-reducing core layer, and install one end of the weight-reducing core layer in the groove reserved in the wing box tap;

[0052] Assemble the wing box reinforcement sheet with the wing box tap and the weight-reducing core layer, and install the wing box reinforcement sheet in the grooves reserved in the wing box tap and the weight-reducing core layer;

[0053] The wing box tap, wing box reinforcement sheet and weight-reducing core layer assembly are coated with fiber prepreg and pre-cured;

[0054] Assembling the pre-cured wing box tap joint, wing box reinforcement sheet and weight-reducing core layer assembly with the main joint;

[0055] For the assemblies of main joints, wing box branch joints, wing box reinforcement sheets and weight-reducing core layers, fiber prepregs of different thicknesses are coated at different positions according to the load conditions of the wing surface, and finally the whole is compression molded and cured.

[0056] In one embodiment of the present invention, before the main joint, the wing box tap and the wing box reinforcement are assembled, the main joint, the wing box tap and the wing box reinforcement are sandblasted to increase the roughness of the metal surface and improve the bonding strength of the metal. Further preferably, the sandblasting uses quartz sand of 20-40 mesh and an air pressure of 0.4MPa-0.7MPa.

[0057] In one embodiment of the present invention, before the weight-reducing core layer is assembled with the wing box tap, the foam of the weight-reducing core layer is dried. Further preferably, the drying process is carried out in an oven at 130°C ± 5°C for 3 hours.

[0058] In one embodiment of the present invention, the pre-curing treatment is to place the prepreg-coated assembly into a vacuum bag and then place it into an oven for pre-curing treatment. More preferably, the prepreg-coated assembly is wrapped with a release cloth and a breathable felt in turn and placed into a vacuum bag; the vacuum bag is placed into the oven in a horizontal state, the surface protective film of the vacuum bag sealing strip is torn off, the vacuum joint is inserted, and the vacuum bag is sealed; the vacuum pump is turned on to ensure that the vacuum degree in the vacuum bag is less than -0.095MPa, and the vacuum is drawn at room temperature for 30 minutes; if the room temperature is lower than 25°C, it is necessary to draw a vacuum in the oven for pre-compaction, set the oven temperature to 35°C, start timing when the oven reaches 35°C, turn off the oven power after keeping warm for 30 minutes, and take out the assembly.

[0059] In one embodiment of the present invention, the fiber prepreg with different thicknesses is coated at different positions according to the load condition of the airfoil, including:

[0060] Overall cladding of the 7mm area on the upper and lower wing surfaces: with the leading edge of the main load-bearing beam as the 0° direction, cladding prepreg in the directions of [+45 / 0 / 0 / -45], [0 / 0 / 0 / 90], [90 / 0 / 0 / 0], and [-45 / 0 / 0 / +45] in sequence. The thickness of the prepreg is 7mm.

[0061] Overall cladding of the 5mm area on the upper and lower wing surfaces: with the leading edge of the main load-bearing beam as the 0° direction, cladding prepreg in the directions of [+45 / 0 / 0 / -45] and [-45 / 0 / 0 / +45] in sequence, with a thickness of 5mm;

[0062] Overall cladding of the 4mm area of ​​the upper and lower wing surfaces and the joint area: with the leading edge of the main load-bearing beam as the 0° direction, cladding of prepreg in the [0 / 0 / 0 / 90] and [90 / 0 / 0 / 0] directions in sequence, with a prepreg thickness of 4mm;

[0063] Overall covering the 3mm area of ​​the upper and lower wing surfaces and the joint area: taking the leading edge of the main load-bearing beam as the 0° direction, cover with prepreg in the directions of [+45 / 0 / 0 / -45][0 / 0 / 0 / 90][+45 / 0 / 0 / -45][-45 / 0 / 0 / +45][90 / 0 / 0 / 0][-45 / 0 / 0 / +45] in sequence. The thickness of the prepreg is 3mm.

[0064] In one embodiment of the present invention, a method for integrally forming a composite material main load-bearing beam is provided, comprising the following steps:

[0065] The first step is to sandblast the main joint, wing box branch joint and wing box reinforcement plate. Sandblasting is to increase the roughness of the metal surface and improve the bonding strength of the metal parts. The quartz sand used for sandblasting is 20-40 mesh and the air pressure is 0.4MPa-0.7MPa.

[0066] Step 2: Dry the foam of the weight-reducing core layer of the main load-bearing beam in an oven at 130°C ± 5°C for 3 hours, and then use a hair dryer to clean the surface debris. The dried main beam foam should be immediately sealed and stored, and it is forbidden to store it in the air. It can only be taken out when it is used. In this embodiment, the weight-reducing core layer is foam such as Fig.12 , 13 As shown, it comprises 4 long strips of foam, and a strip-shaped notch is provided between the 4 long strips of foam at one end to facilitate insertion into the groove reserved on the wing box tap.

[0067] Step 3: Fit each foam piece to the corresponding wing box joint for trial installation, and insert the foam into the groove reserved for the wing box joint. Trim the foam according to the situation, and use 120-grit sandpaper to grind it to ensure that the foam and the wing box joint insertion end face are fully matched, the gap is less than 1mm, and there is enough gap between the webs of the foam.

[0068] Step 4: Evenly apply a layer of J-47A adhesive film to the wing box joints and foam bonding areas, and perform docking assembly.

[0069] Step 5: Assemble the wing box reinforcement piece with the wing box joint and foam for trial. Install the wing box reinforcement piece in the groove reserved for the wing box joint and foam machine. Locally trim the wing box reinforcement piece to ensure that the joint end surfaces of the wing box reinforcement piece, the wing box joint and the foam are flush.

[0070] Step 6: Evenly apply a layer of J-47A adhesive film to the bonding area of ​​the wing box reinforcement sheet, and assemble it with the wing box joint and foam.

[0071] Step 7: Coat the assemblies of the four wing box joints, wing box reinforcement sheets and foams with a layer of prepreg in the direction of [90 / -45 / 0 / +45] respectively, and butt them together in sequence at the web corners of the upper and lower wing surfaces of each foam.

[0072] Step 8: Wrap the prepreg-covered components with release cloth and breathable felt in turn, and put them into a vacuum bag prepared in advance.

[0073] Place the vacuum bag horizontally in the oven, tear off the protective film on the surface of the vacuum bag sealing strip, insert the vacuum joint, and seal the vacuum bag.

[0074] Turn on the vacuum pump to ensure that the vacuum degree in the vacuum bag is less than -0.095MPa, and evacuate the bag at room temperature for 30 minutes. If the room temperature is lower than 25°C, it is necessary to evacuate the bag in an oven for pre-compaction, set the oven temperature to 35°C, start timing when the oven reaches 35°C, keep the temperature for 30 minutes, turn off the oven power supply, and take out the assembly.

[0075] Step 9: Evenly stick a layer of J-47A adhesive film on the bonding area of ​​the main joint, and place the wing box sub-joint into the corresponding position of the main joint, ensuring that it fits perfectly with the contact surface of the main joint with a gap of less than 1mm.

[0076] Step 10: Remove the prepreg from the screw installation positions of the wing box tap and main joint, apply a layer of anaerobic glue evenly on the screw thread area, put each screw into the corresponding installation position, tighten the screw, and ensure that the screw is flush with the surface of the wing box tap. Fill the screw hole with an appropriate amount of prepreg to make the screw surface flush.

[0077] Step 11: Overall clad the 7mm area of ​​the upper and lower wing surfaces. With the leading edge of the main load-bearing beam as the 0° direction, clad the prepreg in the directions of [+45 / 0 / 0 / -45], [0 / 0 / 0 / 90], [90 / 0 / 0 / 0], and [-45 / 0 / 0 / +45] in sequence, and butt them in sequence at the trailing edge of the upper wing surface web corner. The thickness of the prepreg in this step is 7mm, and the position of the 7mm area is as follows: Fig.11 shown.

[0078] Step 12: Overall cladding of the 5mm area on the upper and lower wing surfaces. With the leading edge of the main load-bearing beam as the 0° direction, clad the prepreg in the directions of [+45 / 0 / 0 / -45] and [-45 / 0 / 0 / +45] in sequence, and butt them in sequence at the trailing edge of the lower wing surface web corner. The thickness of the prepreg in this step is 5mm, and the position of the 5mm area is as follows: Fig.11 shown.

[0079] Step 13: Overall cladding of the 4mm area and joint area of ​​the upper and lower wing surfaces. With the leading edge of the main load-bearing beam as the 0° direction, clad the prepreg in the [0 / 0 / 0 / 90] and [90 / 0 / 0 / 0] directions in sequence, and butt them in sequence at the trailing edge of the upper wing surface web corner. The thickness of the prepreg in this step is 4mm, and the position of the 4mm area and the joint area is as follows: Fig.11 shown.

[0080] Step 14: Wrap the prepreg-covered components with release cloth and breathable felt in turn, and put them into the vacuum bag prepared in advance.

[0081] Place the vacuum bag horizontally in the oven, tear off the protective film on the surface of the vacuum bag sealing strip, insert the vacuum joint, and seal the vacuum bag.

[0082] Turn on the vacuum pump to ensure that the vacuum degree in the vacuum bag is less than -0.095MPa, and evacuate the bag at room temperature for 30 minutes. If the room temperature is lower than 25°C, it is necessary to evacuate the bag in an oven for pre-compaction, set the oven temperature to 35°C, start timing when the oven reaches 35°C, keep the temperature for 30 minutes, turn off the oven power, and take out the assembly.

[0083] Step 15: Overall clad the 3mm area and joint area of ​​the upper and lower wing surfaces. From the wing root to the wing tip, with the leading edge of the main load-bearing beam as the 0° direction, clad the prepreg in the directions of [+45 / 0 / 0 / -45][0 / 0 / 0 / 90][+45 / 0 / 0 / -45][-45 / 0 / 0 / +45][90 / 0 / 0 / 0][-45 / 0 / 0 / +45] in sequence, and butt them in sequence at the trailing edges of the upper and lower wing surface web corners. The thickness of the prepreg in this step is 3mm, and the positions of the 3mm area and the joint area are as follows: Fig.11 shown.

[0084] Step 16: Place the wrapped airfoil preform into a compression molding mold and place the mold on a hot press.

[0085] Step 17: Heat the hot press from room temperature → to 130℃±5℃ at a rate of (2-5)℃ / min → keep warm for 2h → apply a pressure of 0.3MPa → continue to heat to 185℃±5℃ → keep warm for 2h, and turn off the hot press.

[0086] Step 18: Naturally cool down the mold temperature to below 50°C and demould the product.

[0087] The position and direction of laying the prepreg in the above steps can be adjusted according to actual conditions and are not limited to the implementation methods in the above steps.

[0088] The specific embodiments of the present invention disclosed above are intended to help understand the content of the present invention and implement it accordingly. It can be understood by those skilled in the art that various replacements, changes and modifications are possible without departing from the spirit and scope of the present invention. The present invention should not be limited to the contents disclosed in the embodiments of this specification, and the scope of protection of the present invention shall be subject to the scope defined in the claims.

Claims

1. A composite material main bearing beam, characterized in that: It includes a main joint, a wing box tap, a wing box reinforcement sheet, a weight-reducing core layer and a skin, wherein the main joint is fixedly connected to the wing box tap and the wing box reinforcement sheet, and the weight-reducing core layer is fixedly connected to the wing box tap; the skin is covered on the outside of the main joint, the wing box tap, the wing box reinforcement sheet and the weight-reducing core layer; wherein the wing box tap is installed in the corresponding groove of the main joint, the wing box reinforcement sheet is installed in the groove reserved for the wing box tap and the weight-reducing core layer, and one end of the weight-reducing core layer is installed in the groove reserved for the wing box tap; the main joint, the wing box tap and the wing box reinforcement sheet are metal parts.

2. The composite main load-bearing beam according to claim 1, characterized in that: The skin is formed by coating the main joint, the wing box branch joint, the wing box reinforcement sheet and the weight-reducing core layer with fiber prepreg according to the load conditions of the wing surface, and then molding and curing the whole.

3. The composite main load-bearing beam according to claim 2, characterized in that: The fiber prepreg is a carbon fiber prepreg, which contains carbon fiber and resin. The carbon fiber adopts unidirectional T300, T700, T800 or T1000 carbon fiber, and the resin adopts medium-temperature or high-temperature curing epoxy resin.

4. The composite main load-bearing beam according to claim 1, characterized in that: The weight-reducing core layer is rigid polyurethane foam, phenolic resin foam or PMI foam.

5. An integral forming method applicable to the composite material main load-bearing beam according to claim 1, characterized in that: The following steps are involved: Assemble the wing box tap and the weight-reducing core layer, and install one end of the weight-reducing core layer in the groove reserved in the wing box tap; Assemble the wing box reinforcement sheet with the wing box tap and the weight-reducing core layer, and install the wing box reinforcement sheet in the grooves reserved in the wing box tap and the weight-reducing core layer; The wing box tap, wing box reinforcement sheet and weight-reducing core layer assembly are coated with fiber prepreg and pre-cured; Assembling the pre-cured wing box tap joint, wing box reinforcement sheet and weight-reducing core layer assembly with the main joint; For the assemblies of main joints, wing box branch joints, wing box reinforcement sheets and weight-reducing core layers, fiber prepregs of different thicknesses are coated at different positions according to the load conditions of the wing surface, and the whole is molded and cured.

6. The method according to claim 5, characterized in that Before the main joint, the wing box branch joint and the wing box reinforcement sheet are assembled, the main joint, the wing box branch joint and the wing box reinforcement sheet are sandblasted; the sandblasting uses quartz sand of 20-40 meshes and the air pressure is 0.4MPa-0.7MPa.

7. The method according to claim 5, characterized in that Before assembling the weight-reducing core layer and the wing box tap, the foam of the weight-reducing core layer is dried; the drying process is carried out in an oven at 130°C±5°C for 3 hours.

8. The method according to claim 5, characterized in that The pre-curing treatment includes: wrapping the assembly coated with prepreg with release cloth and breathable felt in turn, and putting it into a vacuum bag; putting the vacuum bag horizontally into an oven, tearing off the surface protective film of the vacuum bag sealing strip, inserting a vacuum joint, and sealing the vacuum bag; turning on the vacuum pump to ensure that the vacuum degree in the vacuum bag is less than -0.095MPa, setting the oven temperature to 35°C, starting the timing when the oven reaches 35°C, keeping the temperature for 30 minutes, turning off the oven power, and taking out the assembly.

9. The method according to claim 5, characterized in that The fiber prepreg with different thicknesses is coated at different positions according to the load condition of the airfoil, including: Overall cladding of the 7mm area on the upper and lower wing surfaces: with the leading edge of the main load-bearing beam as the 0° direction, cladding prepreg in the directions of [+45 / 0 / 0 / -45], [0 / 0 / 0 / 90], [90 / 0 / 0 / 0], and [-45 / 0 / 0 / +45] in sequence. The thickness of the prepreg is 7mm. Overall cladding of the 5mm area on the upper and lower wing surfaces: with the leading edge of the main load-bearing beam as the 0° direction, cladding prepreg in the directions of [+45 / 0 / 0 / -45] and [-45 / 0 / 0 / +45] in sequence, with a thickness of 5mm; Overall cladding of the 4mm area of ​​the upper and lower wing surfaces and the joint area: with the leading edge of the main load-bearing beam as the 0° direction, cladding of prepreg in the [0 / 0 / 0 / 90] and [90 / 0 / 0 / 0] directions in sequence, with a prepreg thickness of 4mm; Overall covering the 3mm area of ​​the upper and lower wing surfaces and the joint area: taking the leading edge of the main load-bearing beam as the 0° direction, cover with prepreg in the directions of [+45 / 0 / 0 / -45][0 / 0 / 0 / 90][+45 / 0 / 0 / -45][-45 / 0 / 0 / +45][90 / 0 / 0 / 0][-45 / 0 / 0 / +45] in sequence. The thickness of the prepreg is 3mm.

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