Co-curing forming method for large-size circumferential composite material cavity skin of airplane

Abstract

The invention discloses a co-curing forming method for a large-size circumferential composite material cavity skin of an airplane, which comprises the following steps: step (A), mold splitting design; step (B), distributing prepreg and an inner skin, and paving a solid reinforcing area; (C) performing combined positioning and outer skin paving; (D) hot-pressing and co-curing; and (E) demolding and forming: after co-curing is finished, drawing out the fluoropolymer core material from the cavity channel of the solid reinforcing area, and taking down the pressure pad soft mold to obtain the large-size circumferential composite material cavity skin component of the airplane. The large-size circumferential composite material cavity skin of the aircraft is formed by co-curing forming of the inner skin, the solid reinforcing area and the outer skin, the forming process has the advantages of being high in forming strength and good in sealing performance, a thought is provided for an inner and outer skin stiffened wall plate forming process method, meanwhile, riveting is eliminated, and the production efficiency is improved. The quality of the product is reduced; and the strength is improved.

Description

Technical Field

[0001] The invention relates to the technical field of advanced composite material manufacturing, and in particular to a co-curing molding method for a large-size circumferential composite material cavity skin of an aircraft. Background Art

[0002] Advanced composite materials have the advantages of light weight, high strength, high modulus, fatigue resistance, corrosion resistance, designability and good processability. They are especially suitable for large structures and integral structures and are ideal aviation structural materials. At present, advanced composite materials have begun to be used in the main load-bearing components of large passenger aircraft independently developed by my country. In addition, the outer shell of the drone is made of engineering plastics. This type of material has certain disadvantages in terms of impact strength and corrosion resistance. It is gradually being replaced by carbon fiber composite materials that are light in weight, high in specific stiffness, high in strength, and can be made into various shapes and structures in one piece, so as to improve the carrying capacity of the drone and extend its endurance.

[0003] Skin panels are one of the important components of aircraft structures. The reinforced panel structure composed of skin and frame has large bearing capacity and rigidity, but its own weight is very light, which plays the role of bearing and transmitting aerodynamic loads. After bearing the aerodynamic force, the skin panels transfer the force to the connected fuselage and wing frames. The force is complex. In addition, the skin panels are in direct contact with the outside world, so the material is required to have not only high strength and good plasticity, but also a smooth surface and high corrosion resistance. Modern aircraft skin panels widely use high-strength aluminum and magnesium alloys, and some high-performance aircraft use titanium alloys or composite materials. Composite skin panels are widely used in new fighters and aircraft wing structures designed in recent years due to their excellent properties such as light weight, high strength, high modulus, fatigue resistance, and corrosion resistance.

[0004] At present, the forming process of composite skin panels mainly adopts the forming method of riveted composite panels. However, the riveted composite panel method has the following disadvantages: (1) The process is complicated and the overall consistency is poor, which makes the skin easy to become unstable and reduces the aerodynamic characteristics; (2) There are a large number of rivets, and the surface is not smooth enough, which reduces the accuracy of the aerodynamic shape; (3) The large number of nails leads to a large amount of sealing material and increased weight, which is contrary to the original design intention of reducing the weight of UAVs, fighter jets, and aircraft. In particular, for large-size circumferential composite skin panels of aircraft, the disadvantages are more prominent.

[0005] Therefore, how to improve the composite skin panel forming process using the traditional riveting method is a problem that needs to be solved at present. Summary of the invention

[0006] The technical problem solved by the present invention is to overcome the problem that the existing composite blades cannot ensure that the tenon processing matches the actual blade shape. The present invention is based on the actual formed composite blade, scans the digital model and the raw material digital model for fitting, and based on the reverse design of the tooling surface, accurately positions the blade, ensures the positioning accuracy, improves the positioning efficiency, and does not cause clamping deformation during the blade positioning and clamping process, and has good application prospects.

[0007] In order to solve the above technical problems and achieve the above objectives, the technical solution adopted by the present invention is: A co-curing molding method for a large-size circumferential composite cavity skin of an aircraft comprises the following steps: Step (A), mold splitting design: (A1), based on the numerical model of large-sized circumferential composite cavity skin components of aircraft, its structure is divided into inner skin, outer skin and solid reinforcement area; (A2) According to the structural forms of the inner skin, outer skin and solid reinforcement area, the corresponding metal forming tooling and pressure pad forming tooling are manufactured, and the ply thickness and ply structure of each part are determined; Step (B), distribution of prepreg and laying of inner skin and solid reinforcement area: (B1), the prepreg used to prepare the cavity skin composite component is divided into three corresponding parts: inner skin, outer skin and solid reinforcement area according to the split mold design; (B2), prepreg laying, inner skin laying on metal forming tooling, laser positioning device is used to assist sheet positioning during laying to ensure the angle and position of prepreg sheet; (B3), laying the prepreg of the solid reinforcement area on the universal flat tooling, and inserting the fluoropolymer core material into the cavity channel of the solid reinforcement area; Step (C), assembly positioning and outer skin laying: (C1), after the solid reinforcement area and the fluoropolymer core material are assembled, a laser projector is used to locate the solid reinforcement area; (C2), adjust the placement of the fluoropolymer core material according to the auxiliary positioning of the margin area engraved lines at both ends of the universal flat tooling; (C3), after adjusting the placement position of the fluoropolymer core material, the triangular gap formed by the solid reinforcement area and the fluoropolymer core material is filled with carbon twisted wire; (C4), after the carbon twisted wire is filled, it is pre-compacted and the outer skin is laid to obtain a cavity skin preform; Step (D), hot pressing co-curing: (D1), positioning the cavity skin preform with a pressure pad soft mold by a pressure pad molding tool, and vacuum packaging with a vacuum bag; (D2), after vacuum packaging, it is sent into the autoclave for hot pressing and co-curing; Step (E), demoulding: After the co-curing is completed, the fluoropolymer core material is pulled out from the cavity channel of the solid reinforcement area, and the pressure pad soft mold is removed to obtain a large-sized circumferential composite cavity skin component of the aircraft.

[0008] Preferably, (B1), the prepreg is in the form of a combination of carbon fiber fabric, carbon fiber unidirectional tape, and antistatic adhesive film to ensure the mechanical strength and conductivity requirements of the cavity skin.

[0009] Preferably, (B2) and (B3) can be performed simultaneously.

[0010] Preferably, (B2), prepreg paving is performed, and the inner skin is paved on a metal forming tool. During paving, a laser locator is used to assist in positioning the sheet to ensure the angle and position of the prepreg sheet. The specific implementation process is as follows: the surface curvature of the parts of the large-size circumferential composite cavity skin of the aircraft is large, and a detachable projector can be placed on an adjustment vehicle to adjust the angle in real time to assist in positioning.

[0011] Preferably, (B3), the prepreg of the solid reinforcement area is laid on a universal flat tooling, and the fluoropolymer core material is inserted into the cavity channel of the solid reinforcement area. If the cavity channel of the solid reinforcement area has conductivity requirements, the anti-static film is wrapped and laid on the fluoropolymer core material, and the surface film laying angles of the fluoropolymer core materials of all solid reinforcement areas are required to be consistent, and the thickness of a single solid reinforcement area is consistent with that of a single fluoropolymer core material after being wrapped with anti-static film.

[0012] Preferably, (C3), the width of the filling twisted wire is calculated according to L = S / d, where L is the twisted wire width, in mm; S is the cross-sectional area of ​​the triangular gap, in mm 2 ; d is the thickness of a single layer of twisted silk, in mm.

[0013] Preferably, (C4), pre-compaction is performed by using a roller or a hot air blower to assist compaction, specifically by making a vacuum bag on the surface of the paved product and applying pressure to the paved prepreg by vacuum pressure, the pre-compaction pressure range is -34Kpa~-98Kpa, and the pressure time is 5 minutes~15 minutes; after the paving is completed, the product and the mold can be sent to the autoclave for low-temperature hot compaction, the temperature range is 45℃~55℃, the pressure range is in the range of 340Kpa~680Kpa, and the pressure time is 30 minutes~60 minutes.

[0014] Preferably, (D1), the pressure pad is made of flexible high-temperature resistant rubber material with a minimum temperature resistance of 190°C and an elongation of ≥300%.

[0015] Preferably, (D2), after vacuum packaging, the mixture is sent into an autoclave for hot pressing co-curing, and the curing process parameters are a pressure of 0.6-0.8 MPa, a curing temperature of 180±6°C, and a curing time of 240-300 min. And when the pressure is applied, it gradually increases from 0.1MPa to the applied pressure, which is used to prevent the pressure pad from being unevenly stressed due to direct pressure application; When heating and cooling, the rate should be no more than 1°C / min to prevent the fluoropolymer core material from expanding too quickly and causing damage between layers.

[0016] Preferably, in step (E), when the fluoropolymer core material is extracted, the direction of the force applied should be consistent with the normal direction of the cavity channel.

[0017] The beneficial effects of the present invention are as follows: the co-curing molding method of the large-size circumferential composite cavity skin of an aircraft of the present invention has the following advantages compared with the riveting combined process: (1) The overall and local stiffness of the structure is good, the skin is not easy to become unstable, and the aerodynamic characteristics are improved; (2) The number of rivets is reduced, the surface is smoother, and the accuracy of the aerodynamic shape is improved; (3) The amount of rivets and sealing materials used is reduced, and the weight is reduced by 10%~15% compared to general riveted modular wall panels.

[0018] Moreover, the present invention completes the manufacturing process of large-size circumferential composite cavity skin of aircraft by adopting the autoclave co-curing molding method of "soft mold pressure pad + fluoropolymer core material + metal tooling", and achieves the following characteristics: (1) The inner skin, solid reinforcement area, and outer skin are laid and co-cured on the corresponding tooling. This molding process improves the molding quality and the laying efficiency; (2) After the inner skin, solid reinforcement area, and outer skin are co-cured and formed, unlike co-bonding and secondary bonding, which require multiple tank fillings, the manufacturing cost is reduced; (3) After the inner skin, solid reinforcement area, and outer skin are co-cured, no standard parts need to be riveted and fixed, which reduces the weight of the parts; (4) Use fluoropolymer core materials instead of traditional airbag soft molds to reduce airbag tooling costs and improve manufacturing efficiency; (5) This molding method can achieve uniform internal pressurization of large-size, large-curvature complex structure composite materials, avoiding the problem of internal stratification, and is suitable for the production and molding of large-size aircraft skin panels; (6) During the molding process of the aircraft's fuel tank structure area, a layer of anti-static film can be covered to meet the safety performance and reduce the amount of sealing materials used in the subsequent assembly process; (7) The problem of assembly tolerance allocation is solved by using a co-curing molding process for a large-size, large-curvature skin soft mold pressure pad. BRIEF DESCRIPTION OF THE DRAWINGS

[0019] Figure 1 It is a process schematic diagram of the co-curing molding method of large-size circumferential composite cavity skin of aircraft.

[0020] Figure 2 It is a schematic structural cross-sectional diagram of a large-sized circumferential composite cavity skin of an aircraft according to an embodiment of the present invention; Figure 3 yes Figure 2 Specific process flow diagram of the embodiment; Figure 4 yes Figure 2 Three views of the embodiment; Figure 5 yes Figure 2 Schematic diagram of the cavity channel of the solid reinforcement area of ​​an embodiment. DETAILED DESCRIPTION

[0021] The present invention will be further described below in conjunction with the accompanying drawings.

[0022] like Figure 1 As shown, the co-curing molding method of the large-size circumferential composite cavity skin of the aircraft of the present invention adopts the "pressure pad soft mold + fluoropolymer core material + metal tooling" autoclave molding method to complete the manufacturing process of the large-size circumferential composite cavity skin co-curing composite material, which specifically includes the following steps: Step (A), mold splitting design: (A1), based on the numerical model of large-sized circumferential composite cavity skin components of aircraft, its structure is divided into inner skin, outer skin and solid reinforcement area; (A2) According to the structural forms of the inner skin, outer skin and solid reinforcement area, the corresponding metal forming tooling and pressure pad forming tooling are manufactured, and the ply thickness and ply structure of each part are determined; Step (B), distribution of prepreg and laying of inner skin and solid reinforcement area: (B1), the prepreg used to prepare the cavity skin composite material component is divided into three corresponding parts: inner skin, outer skin and solid reinforcement area according to the split mold design. The prepreg is in the form of a combination of carbon fiber fabric, carbon fiber unidirectional tape and antistatic film to ensure the mechanical strength and conductivity requirements of the cavity skin; (B2), prepreg laying is performed, and the inner skin is laid on the metal forming tooling. During laying, a laser positioning device is used to assist the positioning of the sheet to ensure the angle and position of the prepreg sheet. The specific implementation process is as follows: the surface curvature of the parts of the large-sized circumferential composite cavity skin of the aircraft is large, and a detachable projector can be placed on the adjustment vehicle to adjust the angle in real time to assist in positioning; (B3), prepreg of solid reinforcement area is laid on universal flat tooling, and fluoropolymer core material is inserted into the cavity channel of solid reinforcement area. If the cavity channel of solid reinforcement area has conductivity requirement, antistatic film is wrapped and laid on fluoropolymer core material, and the surface film laying angle of fluoropolymer core material of all solid reinforcement areas is required to be consistent, and the thickness of single solid reinforcement area and single fluoropolymer core material after antistatic film is wrapped is consistent. Preferably, (B2) and (B3) can be carried out simultaneously; Step (C), assembly positioning and outer skin laying: (C1), after the solid reinforcement area and the fluoropolymer core material are assembled, a laser projector is used to locate the solid reinforcement area; (C2), adjust the placement of the fluoropolymer core material according to the auxiliary positioning of the margin area engraved lines at both ends of the universal flat tooling; (C3), after adjusting the placement of the fluoropolymer core material, fill the triangular gap formed by the solid reinforcement area and the fluoropolymer core material with carbon twisted wire. The width of the filled twisted wire is calculated according to L = S / d, where L is the twisted wire width, in mm; S is the cross-sectional area of ​​the triangular gap, in mm 2 ; d is the thickness of the twisted yarn single layer, in mm; (C4), after the filling of carbon twisted filaments is completed, pre-compacting is performed, and the outer skin is laid to obtain a cavity skin preform. The pre-compacting is performed by using a roller or a hot air blower to assist the compaction. In addition, the pre-compacting is performed by using a roller or a hot air blower to assist the compaction. Specifically, a vacuum bag is made on the surface of the laid product, and the laid prepreg is pressed by vacuum pressure. The pre-compacting pressure range is -34Kpa~-98Kpa, and the pressure time is 5 minutes~15 minutes. After the laying is completed, the product and the mold can be sent to the autoclave for low-temperature hot compaction. The temperature range is 45℃~55℃, the pressure range is 340Kpa~680Kpa, and the pressure time is 30 minutes~60 minutes. This temperature can soften materials such as carbon fiber prepreg and antistatic film, and will not reach the temperature at which the material undergoes a glass transition reaction, so that it is more tightly combined under high pressure, reducing the internal bubble content and reducing the possibility of internal quality problems after the product is cured and formed; Step (D), hot pressing co-curing: (D1), positioning the cavity skin preform with a pressure pad soft mold by a pressure pad molding tool, and vacuum packaging with a vacuum bag; (D2), after vacuum packaging, it is sent into the autoclave for hot pressing co-curing. The curing process parameters are pressure of 0.6-0.8MPa, curing temperature of 180±6℃, and curing time of 240-300min. And when the pressure is applied, it gradually increases from 0.1MPa to the applied pressure, which is used to prevent the pressure pad from being unevenly stressed due to direct pressure application; When heating and cooling, the rate should be no more than 1°C / min to prevent the fluoropolymer core material from expanding too quickly and causing damage between layers; Step (E), demolding: After the co-curing is completed, the fluoropolymer core material is pulled out from the cavity channel of the solid reinforcement area, and the pressure pad soft mold is removed to obtain a large-scale circumferential composite cavity skin component of the aircraft. When the fluoropolymer core material is pulled out, the force direction should be consistent with the normal direction of the cavity channel.

[0023] The co-curing molding method of the large-sized circumferential composite cavity skin of the aircraft of the present invention uses auxiliary tooling such as metal molding tooling, pressure pad molding tooling, universal flat tooling, fluoropolymer core material, and pressure pad soft mold. The structural shape suitable for the design can be selected according to the shape requirements of the large-sized circumferential composite cavity skin of the aircraft. The metal molding tooling is used to mold the inner skin, the fluoropolymer core material occupies the position spacing of the cavity channel, the pressure pad molding tooling is provided with positioning guide pins, the fluoropolymer core material is provided with position markings for mold closing positioning, and there are lifting holes to facilitate the transportation of the above-mentioned tooling.

[0024] The co-curing molding method of the large-size circumferential composite cavity skin of the aircraft of the present invention has a better effect, especially in the fuel tank area of ​​the aircraft wing structure. The molding process has a high sealing performance, which provides a favorable factor for the overall fuel tank design. At the same time, it provides ideas for the integrated molding process method of complex composite structural parts, reduces the manufacturing cost of components, and expands the application of composite materials in the civilian field. Figure 2-Figure 5 The specific implementation process of Figure 2 It is a schematic structural cross-sectional view of a large-sized circumferential composite cavity skin of an aircraft according to an embodiment; Figure 3 yes Figure 2 Specific process flow diagram of the embodiment; Figure 4 yes Figure 2 Three views of the embodiment; Figure 5 yes Figure 2 Schematic diagram of the cavity channel of the solid reinforcement area of ​​an embodiment.

[0025] In summary, the co-curing molding method of the large-sized circumferential composite cavity skin of the aircraft of the present invention is formed by co-curing the inner skin, the solid reinforcement area and the outer skin, and the maximum length of the skin is about 4m. Figure 2-5A specific example of a large-sized circumferential composite cavity skin of an aircraft is described. The solid reinforcement area is a "I"-shaped structure, and the local reinforcement area is a "U"-shaped structure. The inner skin is laid on the metal forming tooling according to requirements, and the solid reinforcement area can be laid on the universal flat tooling at the same time. A laser projector is used to position the solid reinforcement area and the fluoropolymer core material. After the positioning is completed, the outer skin is laid and the pressure pad is positioned on the surface, and then co-curing is performed. This molding process has the advantages of high molding strength and good sealing. It not only provides ideas for the inner and outer skin reinforced wall panel molding process method, but also eliminates riveting, reduces the quality of the product and improves its strength. The present invention completes the manufacturing process of large-sized circumferential composite cavity skin of an aircraft by adopting the "soft mold pressure pad + fluoropolymer core material + metal tooling" autoclave co-curing molding method, and achieves the following characteristics: (1) The inner skin, solid reinforcement area, and outer skin are laid and co-cured on the corresponding tooling. This molding process improves the molding quality and the laying efficiency; (2) After the inner skin, solid reinforcement area, and outer skin are co-cured and formed, unlike co-bonding and secondary bonding, which require multiple tank fillings, the manufacturing cost is reduced; (3) After the inner skin, solid reinforcement area, and outer skin are co-cured, no standard parts need to be riveted and fixed, which reduces the weight of the parts; (4) Use fluoropolymer core materials instead of traditional airbag soft molds to reduce airbag tooling costs and improve manufacturing efficiency; (5) This molding method can achieve uniform internal pressurization of large-size, large-curvature complex structure composite materials, avoiding the problem of internal stratification, and is suitable for the production and molding of large-size aircraft skin panels; (6) During the molding process of the aircraft's fuel tank structure area, a layer of anti-static film can be covered to meet the safety performance and reduce the amount of sealing materials used in the subsequent assembly process; (7) The problem of assembly tolerance allocation is solved by using a co-curing molding process for a large-size, large-curvature skin soft mold pressure pad.

[0026] The above shows and describes the basic principles and main features of the present invention and the advantages of the present invention. It should be understood by those skilled in the art that the present invention is not limited to the above embodiments. The above embodiments and descriptions are only for explaining the principles of the present invention. Without departing from the spirit and scope of the present invention, the present invention may have various changes and improvements, which fall within the scope of the present invention to be protected. The scope of protection of the present invention is defined by the attached claims and their equivalents.

Claims

1. A co-curing molding method for large-size circumferential composite cavity skin of aircraft, characterized in that: The following steps are included: Step (A), mold splitting design: (A1), based on the numerical model of large-sized circumferential composite cavity skin components of aircraft, its structure is divided into inner skin, outer skin and solid reinforcement area; (A2) According to the structural forms of the inner skin, outer skin and solid reinforcement area, the corresponding metal forming tooling and pressure pad forming tooling are manufactured, and the ply thickness and ply structure of each part are determined; Step (B), distribution of prepreg and inner skin, solid reinforcement area laying: (B1), the prepreg used to prepare the cavity skin composite component is divided into three corresponding parts: inner skin, outer skin and solid reinforcement area according to the split mold design; (B2), prepreg laying, inner skin laying on metal forming tooling, laser positioning device is used to assist sheet positioning during laying to ensure the angle and position of prepreg sheet; (B3), laying the prepreg of the solid reinforcement area on the universal flat tooling, and inserting the fluoropolymer core material into the cavity channel of the solid reinforcement area; Step (C), assembly positioning and outer skin laying: (C1), after the solid reinforcement area and the fluoropolymer core material are assembled, a laser projector is used to locate the solid reinforcement area; (C2), adjust the placement of the fluoropolymer core material according to the auxiliary positioning of the margin area engraved lines at both ends of the universal flat tooling; (C3), after adjusting the placement position of the fluoropolymer core material, the triangular gap formed by the solid reinforcement area and the fluoropolymer core material is filled with carbon twisted wire; (C4), after the carbon twisted wire is filled, it is pre-compacted and the outer skin is laid to obtain a cavity skin preform; Step (D), hot pressing co-curing: (D1), positioning the cavity skin preform with a pressure pad soft mold by a pressure pad molding tool, and vacuum packaging with a vacuum bag; (D2), after vacuum packaging, it is sent into the autoclave for hot pressing and co-curing; Step (E), demoulding: After the co-curing is completed, the fluoropolymer core material is pulled out from the cavity channel of the solid reinforcement area, and the pressure pad soft mold is removed to obtain a large-sized circumferential composite cavity skin component of the aircraft.

2. The co-curing molding method of large-size circumferential composite cavity skin of aircraft according to claim 1 is characterized in that: (B1), the prepreg is in the form of a combination of carbon fiber fabric, carbon fiber unidirectional tape, and antistatic adhesive film to ensure the mechanical strength and conductivity requirements of the cavity skin.

3. The co-curing molding method of large-size circumferential composite cavity skin of aircraft according to claim 1 is characterized in that: (B2) and (B3) can be performed simultaneously.

4. The co-curing molding method of large-size circumferential composite cavity skin of aircraft according to claim 3 is characterized in that: (B2) Prepreg laying is carried out and the inner skin is laid on the metal forming tooling. During laying, a laser locator is used to assist in the positioning of the sheet to ensure the angle and position of the prepreg sheet. The specific implementation process is as follows: The surface curvature of the parts of the large-size circumferential composite cavity skin of the aircraft is large. A detachable projector can be placed on the adjustment vehicle to adjust the angle in real time to assist in positioning.

5. The co-curing molding method of large-size circumferential composite cavity skin of aircraft according to claim 3, characterized in that: (B3) The prepreg of the solid reinforcement area is laid on the universal flat tooling, and the fluoropolymer core material is inserted into the cavity channel of the solid reinforcement area. If the cavity channel of the solid reinforcement area has conductivity requirements, the anti-static film is wrapped and laid on the fluoropolymer core material, and the surface film laying angles of all solid reinforcement area fluoropolymer core materials are required to be consistent. The thickness of a single solid reinforcement area is consistent with that of a single fluoropolymer core material after being wrapped with anti-static film.

6. The co-curing molding method of large-size circumferential composite cavity skin of aircraft according to claim 1, characterized in that: (C3), the width of the filling twisted wire, is calculated according to L = S / d, where L is the twisted wire width, in mm; S is the cross-sectional area of ​​the triangular gap, in mm2; d is the thickness of the single layer of twisted wire, in mm.

7. The co-curing molding method of large-size circumferential composite cavity skin of aircraft according to claim 1, characterized in that: (C4), pre-compaction is achieved by using rollers or hot air blowers to assist compaction. Specifically, a vacuum bag is made on the surface of the paved product and the prepreg is pressed by vacuum pressure. The pre-compaction pressure ranges from -34Kpa to -98Kpa, and the pressure time is 5 minutes to 15 minutes. After paving is completed, the product and mold can be sent to the autoclave for low-temperature hot compaction. The temperature range is 45℃~55℃, the pressure range is 340Kpa~680Kpa, and the pressure time is 30 minutes to 60 minutes.

8. The co-curing molding method of large-size circumferential composite cavity skin of aircraft according to claim 1, characterized in that: (D1), the pressure pad is made of flexible high temperature resistant rubber material, with a minimum temperature resistance of 190°C and an elongation of ≥300%.

9. The co-curing molding method of large-size circumferential composite cavity skin of aircraft according to claim 1, characterized in that: (D2), after vacuum packaging, it is sent into the autoclave for hot pressing co-curing. The curing process parameters are pressure of 0.6-0.8MPa, curing temperature of 180±6℃, and curing time of 240-300min. And when the pressure is applied, it gradually increases from 0.1MPa to the applied pressure, which is used to prevent the pressure pad from being unevenly stressed due to direct pressure application; When heating and cooling, the rate should be no more than 1°C / min to prevent the fluoropolymer core material from expanding too quickly and causing damage between layers.

10. The co-curing molding method of large-size circumferential composite cavity skin of aircraft according to claim 1, characterized in that: In step (E), when the fluoropolymer core material is extracted, the force applied should be in the same direction as the normal direction of the cavity channel.

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