A co-curing forming method of large-size circumferential composite material cavity skin of an airplane

Abstract

The application discloses a co-curing forming method of a large-size circumferential composite cavity skin of an airplane, and comprises the following steps: step (A), mold splitting design; step (B), distribution of prepreg and inner skin and solid stiffening area laying; step (C), combination positioning and outer skin laying; step (D), hot pressing and co-curing; and step (E), demolding forming. After the co-curing is completed, a fluoropolymer core material is drawn out from a cavity channel of the solid stiffening area, a pressure pad soft mold is removed, and a large-size circumferential composite cavity skin component of the airplane is obtained. The large-size circumferential composite cavity skin of the airplane is formed by co-curing of the inner skin, the solid stiffening area and the outer skin, the forming process has the advantages of high forming strength and good sealing performance, not only provides a train of thought for an inner-outer skin stiffened wall plate forming process method, but also eliminates riveting, reduces the weight of the product and improves the strength of the product.

Description

Technical Field

[0001] This invention relates to the field of advanced composite material manufacturing technology, specifically to a co-curing molding method for large-size circumferential composite material cavity skin of aircraft. Background Technology

[0002] Advanced composite materials possess advantages such as lightweight, high strength, high modulus, fatigue resistance, corrosion resistance, good designability, and processability, making them particularly suitable for large and integral structures, and ideal for aerospace structural materials. Currently, advanced composite materials have begun to be used in the main load-bearing components of my country's independently developed large passenger aircraft. Furthermore, the shells of drones are currently made of engineering plastics, which have certain shortcomings in impact resistance and corrosion resistance. These materials are gradually being replaced by carbon fiber composite materials, which are lightweight, have high specific stiffness, high strength, and can be integrally manufactured into various shapes and structures, thereby improving the load-bearing capacity and extending the range of drones.

[0003] Skin panels are a crucial component of aircraft structures. The stiffened panel structure, composed of skin and frame, possesses significant load-bearing capacity and rigidity while remaining lightweight, effectively bearing and transmitting aerodynamic loads. After bearing aerodynamic forces, the skin panels transfer these forces to the connected fuselage and wing frames, resulting in complex stress distribution. Furthermore, since the skin panels are in direct contact with the external environment, the materials required not only must possess high strength and good plasticity, but also a smooth surface and high corrosion resistance. Modern aircraft widely utilize high-strength aluminum and magnesium alloys for their skin panels, while some high-performance aircraft employ titanium alloys or composite materials. Composite material skin panels, due to their superior properties such as lightweight, high strength, high modulus, fatigue resistance, and corrosion resistance, are widely used in new fighter jets and the wing structures of aircraft designed in recent years.

[0004] At present, the main molding process of composite material skin panels adopts the riveting assembly method. However, the riveting assembly method has the following disadvantages: (1) The process is complicated and the overall consistency is poor, which makes the skin unstable and reduces the aerodynamic characteristics; (2) There are many rivets and the surface is not smooth enough, which reduces the accuracy of the aerodynamic shape; (3) There are many rivets, which leads to a large amount of sealing material and increased weight, which contradicts the original intention of reducing the weight of drones, fighters and aircraft. The disadvantages are even more prominent for the large-size circumferential composite material skin panels of aircraft.

[0005] Therefore, how to improve the molding process of composite material skin panels using traditional riveting methods is a problem that needs to be solved. Summary of the Invention

[0006] The technical problem solved by this invention is to overcome the limitations of traditional riveting methods in the molding process of composite material skin panels. The co-curing molding method for large-size circumferential composite cavity skins for aircraft provided by this invention involves the co-curing of an inner skin, a solid reinforcing zone, and an outer skin. This molding process offers advantages such as high molding strength and good sealing performance. It not only provides a new approach to the molding process of reinforced inner and outer skin panels but also eliminates riveting, reduces product weight, and increases its strength, demonstrating promising application prospects.

[0007] To solve the above-mentioned technical problems and achieve the above-mentioned objectives, the technical solution adopted by the present invention is as follows:

[0008] A co-curing molding method for large-size circumferential composite cavity skin of aircraft includes the following steps:

[0009] Step (A), Mold Design:

[0010] (A1) Based on the digital model of the large-size circumferential composite cavity skin component of the aircraft, its structure is divided into inner skin, outer skin and solid reinforcement area;

[0011] (A2) Based on the structural morphology of the inner skin, outer skin and solid reinforcement area, corresponding metal forming fixtures and pressure pad forming fixtures are made, and the ply thickness and ply structure of each part are determined.

[0012] Step (B): Distribute the prepreg and inner skin, and lay the solid reinforcement area:

[0013] (B1) The prepreg used to prepare the cavity skin composite component is divided into three corresponding parts according to the mold design: inner skin, outer skin and solid reinforcement.

[0014] (B2) Prepreg laying is carried out. The inner skin is laid on the metal forming fixture. During the laying, a laser positioning instrument is used to assist in the positioning of the material sheet to ensure the angle and position of the prepreg sheet.

[0015] (B3) The prepreg is laid on the solid reinforcement zone on a general-purpose flat tooling, and the fluoropolymer core material is inserted into the cavity channel of the solid reinforcement zone.

[0016] Step (C), assembly, positioning, and outer skin installation:

[0017] (C1) After the solid reinforcement area and the fluoropolymer core material are assembled, the solid reinforcement area is positioned using a laser projector;

[0018] (C2) Adjust the placement position of the fluoropolymer core material according to the auxiliary positioning of the margin area markings at both ends of the universal flat tooling;

[0019] (C3) After adjusting the placement of the fluoropolymer core material, fill the triangular gap formed by the solid reinforcing area and the fluoropolymer core material with carbon twisted wire.

[0020] (C4) After filling with carbon twisted wire, pre-compact the material and lay the outer skin to obtain a cavity skin preform.

[0021] Step (D), hot pressing co-curing:

[0022] (D1) The cavity skin preform is positioned by pressure pad soft mold using pressure pad forming tooling, and vacuum sealed using vacuum bag.

[0023] (D2) After vacuum sealing, it is sent to an autoclave for hot pressing and co-curing;

[0024] Step (E), Demolding and Molding: After co-curing, the fluoropolymer core material is extracted from the cavity channel of the solid reinforcement area, and the pressure pad soft mold is removed to obtain the large-size circumferential composite cavity skin component for aircraft.

[0025] Preferably, (B1), the prepreg is 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.

[0026] Preferably, (B2) and (B3) can be performed simultaneously.

[0027] Preferably, (B2) prepreg is laid, and the inner skin is laid on a metal forming fixture. During the laying, a laser positioning instrument is used to assist in the positioning of the material to ensure the angle and position of the prepreg. The specific implementation process is as follows: The surface curvature of the large-size circumferential composite cavity skin of the aircraft is large. A projector can be detached and placed on the adjustment vehicle to adjust the angle in real time to assist in positioning.

[0028] Preferably, (B3), prepreg is laid on a general-purpose flat tooling for solid reinforcement areas, and 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, antistatic adhesive film is wrapped and laid on the fluoropolymer core material. The layup angle of the surface adhesive film of all solid reinforcement areas of fluoropolymer core material is required to be consistent, and the thickness of a single solid reinforcement area and a single fluoropolymer core material after wrapping with antistatic adhesive film is consistent.

[0029] Preferably, (C3), the width of the filling twist is calculated according to L = S / d, where L is the twist width in mm; S is the cross-sectional area of ​​the triangular gap region in mm. 2 ;d represents the thickness of a single layer of twisted yarn, in mm.

[0030] Preferably, (C4), pre-compaction is assisted by using rollers or a hot air blower. Specifically, a vacuum bag is made on the surface of the laid product, and the prepreg is pressed by vacuum pressure. The pre-compaction pressure range is -34Kpa to -98Kpa, and the pressure time is 5 minutes to 15 minutes. After laying, the product and mold can be sent into a hot autoclave for low-temperature hot compaction. The temperature range is 45℃ to 55℃, the pressure range is 340Kpa to 680Kpa, and the pressure time is 30 minutes to 60 minutes.

[0031] Preferably, (D1), the pressure pad is made of flexible high-temperature resistant rubber material with a minimum temperature resistance of 190℃ and an elongation of ≥300%.

[0032] Preferably, (D2) after vacuum sealing, it is placed in an autoclave for thermosetting and co-curing. The curing process parameters are: pressure 0.6-0.8 MPa, curing temperature 180±6℃, and curing time 240-300 min.

[0033] Furthermore, when applying pressure, the pressure is gradually increased from 0.1 MPa to the applied pressure to prevent the pressure pad from being soft and unevenly stressed due to direct pressure application.

[0034] Furthermore, the heating and cooling rate should not exceed 1℃ / min to prevent the fluoropolymer core material from expanding too quickly and causing damage between layers.

[0035] Preferably, in step (E), when extracting the fluoropolymer core material, the direction of the applied force should be consistent with the normal direction of the cavity channel.

[0036] The beneficial effects of this invention are as follows: Compared with the riveting assembly process, the co-curing molding method for large-size circumferential composite cavity skin of aircraft of this invention has the following advantages:

[0037] (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;

[0038] (2) The number of rivets has been reduced, the surface is smoother, and the accuracy of the aerodynamic shape has been improved;

[0039] (3) The amount of rivets and sealing materials used is reduced, and the weight is reduced by 10% to 15% compared with the general riveted combination wall panel.

[0040] Furthermore, this invention completes the manufacturing process of large-size circumferential composite material cavity skin for aircraft by employing a co-curing molding method using a "soft mold pressure pad + fluoropolymer core material + metal tooling" autoclave, and achieves the following characteristics:

[0041] (1) The inner skin, solid reinforcement area and outer skin are laid and cured on the corresponding tooling. This molding process improves the molding quality and increases the laying efficiency.

[0042] (2) After the inner skin, solid reinforcement area and outer skin are co-cured and molded, unlike co-bonding and secondary bonding which require multiple injections into the tank, the manufacturing cost is reduced.

[0043] (3) After the inner skin, solid reinforcement area and outer skin are co-cured and formed, no standard parts are required for riveting and fixing, which reduces the weight of the parts;

[0044] (4) Fluoropolymer core material is used instead of traditional airbag soft mold to reduce airbag tooling costs and improve manufacturing efficiency;

[0045] (5) This molding method can achieve uniform internal pressure on large-size, high-curvature complex composite materials, avoiding the problem of internal delamination, and is suitable for the manufacturing and molding of large-size aircraft skin panels;

[0046] (6) During the molding process of the aircraft fuel tank structure area, an antistatic adhesive film can be applied to meet the safety performance requirements and reduce the amount of sealing material used in subsequent assembly processes.

[0047] (7) Solve the problem of assembly tolerance distribution by using a large-size, large-curvature skin soft mold pressure pad co-curing molding process. Attached Figure Description

[0048] Figure 1 This is a schematic diagram of the co-curing molding method for large-size circumferential composite cavity skin of aircraft.

[0049] Figure 2 This is a schematic cross-sectional view of the structure of a large-size circumferential composite material cavity skin for an aircraft, according to an embodiment of the present invention.

[0050] Figure 3 yes Figure 2 A schematic diagram of the specific process flow of the embodiment;

[0051] Figure 4 yes Figure 2 Three views of the embodiment;

[0052] Figure 5 yes Figure 2 A schematic diagram of the cavity channel in the solid reinforcement region of the embodiment. Detailed Implementation

[0053] The present invention will now be further described with reference to the accompanying drawings.

[0054] like Figure 1As shown, the co-curing molding method for large-size circumferential composite cavity skin of aircraft of the present invention adopts a "pressure pad soft mold + fluoropolymer core material + metal tooling" autoclave molding method to complete the manufacturing process of large-size circumferential composite cavity skin co-curing composite material, specifically including the following steps.

[0055] Step (A), Mold Design:

[0056] (A1) Based on the digital model of the large-size circumferential composite cavity skin component of the aircraft, its structure is divided into inner skin, outer skin and solid reinforcement area;

[0057] (A2) Based on the structural morphology of the inner skin, outer skin and solid reinforcement area, corresponding metal forming fixtures and pressure pad forming fixtures are made, and the ply thickness and ply structure of each part are determined.

[0058] Step (B): Distribute the prepreg and inner skin, and lay the solid reinforcement area:

[0059] (B1) The prepreg used to prepare the cavity skin composite material component is divided into three corresponding parts according to the mold design: inner skin, outer skin and solid reinforcement. The prepreg adopts 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.

[0060] (B2) Prepreg laying is carried out. The inner skin is laid on the metal forming tooling. During the laying, a laser positioning instrument is used to assist in the positioning of the material to ensure the angle and position of the prepreg. The specific implementation process is as follows: The surface curvature of the large-size circumferential composite cavity skin of the aircraft is large. A projector can be detached and placed on the adjustment vehicle to adjust the angle in real time to assist in positioning.

[0061] (B3) The prepreg of the solid reinforcement area is laid on the general-purpose 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 antistatic adhesive film is wrapped and laid on the fluoropolymer core material. The layup angle of the surface adhesive film of the fluoropolymer core material of all solid reinforcement areas is required to be consistent. Preferably, the thickness of a single solid reinforcement area and a single fluoropolymer core material after wrapping with antistatic adhesive film is consistent. (B2) and (B3) can be carried out simultaneously.

[0062] Step (C), assembly, positioning, and outer skin installation:

[0063] (C1) After the solid reinforcement area and the fluoropolymer core material are assembled, the solid reinforcement area is positioned using a laser projector;

[0064] (C2) Adjust the placement position of the fluoropolymer core material according to the auxiliary positioning of the margin area markings at both ends of the universal flat tooling;

[0065] (C3) After adjusting the placement of the fluoropolymer core material, fill the triangular gap formed by the solid reinforcing area and the fluoropolymer core material with carbon twisted wire. The width of the twisted wire is calculated according to L = S / d, where L is the width of the twisted wire in mm and S is the cross-sectional area of ​​the triangular gap in mm². 2 ;d represents the thickness of a single layer of twisted yarn, in mm;

[0066] (C4) After filling with carbon twisted yarn, pre-compactment is performed, and the outer skin is laid to obtain a cavity skin preform. Pre-compactment is assisted by using rollers or a hot air blower. Specifically, a vacuum bag is made on the surface of the laid product, and the prepreg is pressed by vacuum pressure. The pressure range of pre-compactment is -34Kpa to -98Kpa, and the pressure time is 5 minutes to 15 minutes. After laying, the product and mold can be sent into a hot autoclave for low-temperature hot compaction. The temperature range is 45℃ to 55℃, the pressure range is 340Kpa to 680Kpa, and the pressure time is 30 minutes to 60 minutes. This temperature can soften materials such as carbon fiber prepreg and antistatic film without reaching the temperature at which the materials undergo glass transition reaction. This results in a tighter bond under high pressure, reduces the content of internal air bubbles, and reduces the possibility of internal quality problems after the product is cured.

[0067] Step (D), hot pressing co-curing:

[0068] (D1) The cavity skin preform is positioned by pressure pad soft mold using pressure pad forming tooling, and vacuum sealed using vacuum bag.

[0069] (D2) After vacuum sealing, the product is placed in an autoclave for thermosetting and co-curing. The curing process parameters are: pressure 0.6-0.8 MPa, curing temperature 180±6℃, and curing time 240-300 min.

[0070] Furthermore, when applying pressure, the pressure is gradually increased from 0.1 MPa to the applied pressure to prevent the pressure pad from being soft and unevenly stressed due to direct pressure application.

[0071] Furthermore, the heating and cooling rate should not exceed 1℃ / min to prevent the fluoropolymer core material from expanding too quickly and causing damage between layers.

[0072] Step (E), Demolding and Molding: After co-curing, the fluoropolymer core material is extracted from the cavity channel of the solid reinforcement area, and the pressure pad soft mold is removed to obtain the large-size circumferential composite cavity skin component of the aircraft. When extracting the fluoropolymer core material, the direction of force should be consistent with the normal direction of the cavity channel.

[0073] The co-curing molding method for large-size circumferential composite cavity skin of aircraft of the present invention utilizes auxiliary tooling including metal forming tooling, pressure pad forming tooling, general-purpose flat tooling, fluoropolymer core material, and pressure pad soft mold. The structural shape can be designed to fit the shape requirements of the large-size circumferential composite cavity skin of aircraft. The metal forming tooling is used for inner skin forming, the fluoropolymer core material occupies the position spacing of the cavity channel, the pressure pad forming tooling is equipped with positioning guide pins, the fluoropolymer core material is equipped with position markings for mold closing positioning, and has lifting holes to facilitate the handling of the above tooling.

[0074] The co-curing molding method for large-size circumferential composite cavity skin of aircraft of the present invention is particularly effective in the fuel tank area of ​​aircraft wing structures. This molding process has high sealing performance, providing a favorable factor for the overall fuel tank design. At the same time, it provides ideas for integrated molding processes of complex composite structural components, reduces the manufacturing cost of parts, and expands the application of composite materials in the civilian field. Figures 2-5 The specific implementation process, Figure 2 This is a schematic cross-sectional view of the structure of a large-size circumferential composite material cavity skin for an aircraft, according to one embodiment. Figure 3 yes Figure 2 A schematic diagram of the specific process flow of the embodiment; Figure 4 yes Figure 2 Three views of the embodiment; Figure 5 yes Figure 2 A schematic diagram of the cavity channel in the solid reinforcement region of the embodiment.

[0075] In summary, the co-curing molding method for large-size circumferential composite cavity skin of aircraft of the present invention involves co-curing an inner skin, a solid reinforcing zone, and an outer skin. The maximum length of the skin is approximately 4m. Figure 2-5 This invention provides a specific example of a large-size circumferential composite cavity skin for an aircraft. The solid reinforcement area has an "I"-shaped structure, while the partial reinforcement area has a "U"-shaped structure. The inner skin is laid on a metal forming fixture as required, while the solid reinforcement area can be laid simultaneously on a general-purpose flat fixture. A laser projector is used to position the solid reinforcement area and the fluoropolymer core material. After positioning, the outer skin is laid on the surface, and the pressure pad is positioned, followed by co-curing. This molding process has the advantages of high molding strength and good sealing performance. It not only provides a new approach to the molding process of inner and outer skin reinforced panels but also eliminates riveting, reduces product weight, and improves its strength. This invention completes the manufacturing process of a large-size circumferential composite cavity skin for an aircraft by using a "soft mold pressure pad + fluoropolymer core material + metal fixture" autoclave co-curing molding method, achieving the following characteristics:

[0076] (1) The inner skin, solid reinforcement area and outer skin are laid and cured on the corresponding tooling. This molding process improves the molding quality and increases the laying efficiency.

[0077] (2) After the inner skin, solid reinforcement area and outer skin are co-cured and molded, unlike co-bonding and secondary bonding which require multiple injections into the tank, the manufacturing cost is reduced.

[0078] (3) After the inner skin, solid reinforcement area and outer skin are co-cured and formed, no standard parts are required for riveting and fixing, which reduces the weight of the parts;

[0079] (4) Fluoropolymer core material is used instead of traditional airbag soft mold to reduce airbag tooling costs and improve manufacturing efficiency;

[0080] (5) This molding method can achieve uniform internal pressure on large-size, high-curvature complex composite materials, avoiding the problem of internal delamination, and is suitable for the manufacturing and molding of large-size aircraft skin panels;

[0081] (6) During the molding process of the aircraft fuel tank structure area, an antistatic adhesive film can be applied to meet the safety performance requirements and reduce the amount of sealing material used in subsequent assembly processes.

[0082] (7) Solve the problem of assembly tolerance distribution by using a large-size, large-curvature skin soft mold pressure pad co-curing molding process.

[0083] The foregoing has shown and described the basic principles, main features, and advantages of the present invention. Those skilled in the art should understand that the present invention is not limited to the above embodiments. The embodiments and descriptions in the specification are merely illustrative of the principles of the invention. Various changes and modifications can be made to the invention without departing from its spirit and scope, and all such changes and modifications fall within the scope of the present invention as claimed. The scope of protection of this invention is defined by the appended claims and their equivalents.

Claims

1. A co-curing molding method for large-size circumferential composite cavity skin of an aircraft, characterized in that, Includes the following steps, Step (A), Mold Design: (A1) Based on the digital model of the large-size circumferential composite cavity skin component of the aircraft, its structure is divided into inner skin, outer skin and solid reinforcement area; (A2) Based on the structural morphology of the inner skin, outer skin and solid reinforcement area, corresponding metal forming fixtures and pressure pad forming fixtures are made, and the ply thickness and ply structure of each part are determined. Step (B): Distribute the prepreg and inner skin, and lay the solid reinforcement area: (B1) The prepreg used to prepare the cavity skin composite component is divided into three corresponding parts according to the mold design: inner skin, outer skin and solid reinforcement. (B2) Prepreg laying is carried out. The inner skin is laid on the metal forming fixture. During the laying, a laser positioning instrument is used to assist in the positioning of the material sheet to ensure the angle and position of the prepreg sheet. (B3) The prepreg is laid on the solid reinforcement zone on a general-purpose flat tooling, and the fluoropolymer core material is inserted into the cavity channel of the solid reinforcement zone. Step (C), assembly, positioning, and outer skin installation: (C1) After the solid reinforcement area and the fluoropolymer core material are assembled, the solid reinforcement area is positioned using a laser projector; (C2) Adjust the placement position of the fluoropolymer core material according to the auxiliary positioning of the margin area markings 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 reinforcing area and the fluoropolymer core material with carbon twisted wire. (C4) After filling with carbon twisted wire, pre-compact the material and lay the outer skin to obtain a cavity skin preform. Step (D), hot pressing co-curing: (D1) The cavity skin preform is positioned by pressure pad soft mold using pressure pad forming tooling, and vacuum sealed using vacuum bag. (D2) After vacuum sealing, it is sent to an autoclave for hot pressing and co-curing; Step (E), Demolding and Molding: After co-curing, the fluoropolymer core material is extracted from the cavity channel of the solid reinforcement area, and the pressure pad soft mold is removed to obtain the large-size circumferential composite cavity skin component for aircraft.

2. The co-curing molding method for large-size circumferential composite material cavity skin of aircraft according to claim 1, characterized in that, (B1) The prepreg is made 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 for large-size circumferential composite cavity skin of aircraft according to claim 1, characterized in that, (B2) and (B3) are carried out simultaneously.

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

5. The co-curing molding method for large-size circumferential composite cavity skin of aircraft according to claim 3, characterized in that, (B3) The prepreg is laid on the solid reinforcement area on a general-purpose 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 antistatic adhesive film is wrapped and laid on the fluoropolymer core material. The layup angle of the surface adhesive film of the fluoropolymer core material in all solid reinforcement areas is required to be consistent. The thickness of a single solid reinforcement area and a single fluoropolymer core material after wrapping with antistatic adhesive film is consistent.

6. The co-curing molding method for 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 width of the twisted wire in mm; S is the cross-sectional area of ​​the triangular gap in mm2; and d is the thickness of a single layer of twisted wire in mm.

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

8. The co-curing molding method for 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℃ and an elongation of ≥300%.

9. The co-curing molding method for large-size circumferential composite cavity skin of aircraft according to claim 1, characterized in that, (D2) After vacuum sealing, the product is placed in an autoclave for thermosetting and co-curing. The curing process parameters are: pressure 0.6-0.8 MPa, curing temperature 180±6℃, and curing time 240-300 min. Furthermore, when applying pressure, the pressure is gradually increased from 0.1 MPa to the applied pressure to prevent the pressure pad from being soft and unevenly stressed due to direct pressure application. Furthermore, the heating and cooling rate should not exceed 1℃ / min to prevent the fluoropolymer core material from expanding too quickly and causing damage between layers.

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

Images