Forming method of ultralight composite material wing structure

Abstract

The invention discloses a forming method of an ultralight composite material wing structure. The forming method comprises the following steps: respectively curing and forming the central wing box, the wingtip and the I-shaped beam, and then paving the skin for secondary curing to form the ultralight composite material wing structure. Wherein the I-shaped beam is formed by paving carbon fiber unidirectional prepreg, vertical ribs of the I-shaped beam are paved with forward and reverse C-shaped prepreg, then the upper surface and the lower surface of the I-shaped beam are respectively paved, and finally, the I-shaped beam is solidified under the temperature condition of 90-135 DEG C and the pressure condition of 0.3-0.6 MPa to form the composite material I-shaped beam structure. The manufacturing method has the advantages that the composite material wing which is good in surface quality, meets the requirement on appearance size and is light in weight is manufactured, particularly, the structural scheme of adopting the I-shaped beam and the thin skin is reasonable, local stress concentration is avoided, and the weight is light; the technical scheme that the chemical beam, the central wing box and the wingtip are fixed step by step and then overall co-curing forming is conducted is adopted, and compared with integrated forming, the method has the advantages that the precision of formed parts is higher, and feasibility is better.

Description

Technical Field

[0001] The present invention relates to a wing structure of an aircraft, and more particularly to a forming method for an ultra-light composite material wing structure. Background Art

[0002] Composite materials mainly rely on reinforcing materials to provide mechanical properties, and the reinforcing materials mainly appear in the form of fibers. They not only have low density, high strength, high modulus, high temperature resistance, chemical corrosion resistance, and low thermal expansion coefficient, but also have the flexibility and plaitability of fibers. Therefore, composite material structures have excellent characteristics such as high specific strength and specific modulus, light weight and high strength, high damage tolerance, and moisture resistance, and are widely used in the aerospace field. The outstanding feature of the airframe structure of the new generation of large airliners is the extensive use of composite materials, which can not only reduce the weight of the aircraft, increase the service life of the aircraft, and reduce the maintenance cost of the aircraft, but also increase the cabin pressure and air humidity, improving the economy and environmental protection of the aircraft.

[0003] With the application of composite materials in wings, their structural design and forming process have always been the focus of research at home and abroad. The structural design of composite material wings mainly includes overall design and internal structure design, and the forming process mainly includes forming method selection, forming process parameter design, etc. Among them, for the structure of the wing, it usually consists of five parts: left and right wing bodies with tapered cross-sections, a center wing box, and left and right wing tips. The wing tips are symmetrically distributed at both ends in opposite directions. The upper and lower surfaces of the wing body and the center wing box are both flat, and the lower surface is required to be a horizontal plane after combination. However, since the structural form of the wing mainly depends on the load transfer path on the wing, in order to make the wing have high compressive strength, high specific strength, and as light a weight as possible, it is necessary to design according to the internal structure form of the wing and reasonable material distribution. Therefore, what form of structure can meet the requirement of ultra-light weight to the greatest extent is a technical problem to be solved urgently. In addition, on this basis, what specific manufacturing process can meet the external dimension requirements of the above specific structural form and achieve the corresponding mechanical property requirements is an even more urgent technical problem to be solved. Summary of the Invention

[0004] The technical problem to be solved by the present invention is to provide a forming method for an ultra-light composite material wing structure with high strength, light weight, and external dimensions that can meet the requirements.

[0005] To solve the above technical problems, a forming method of an ultra-light composite wing structure of the present invention is provided. The ultra-light composite wing structure includes a central wing box, wing bodies on both left and right sides of the central wing box, and wing tips on both left and right sides of the wing bodies. The wing body has an I-beam therein and the I-beam extends and is embedded into the wing tips. The forming method is as follows: First, the central wing box, wing tips, and I-beam are respectively cured and formed, and then a skin is laid and secondarily cured to form an ultra-light composite wing structure; wherein, the I-beam is made by laying carbon fiber unidirectional prepreg, and at the vertical ribs of the I-beam, forward and reverse C-shaped prepreg are laid, then the upper and lower surfaces are respectively laid, and finally it is cured at a temperature condition of 90 - 135 °C and a pressure condition of 0.3 - 0.6 MPa to form a composite I-beam structure.

[0006] The thickness of the I-beam is 1.2 - 2.4 mm, the thickness of the carbon fiber / epoxy single-layer prepreg is 0.1 - 0.2 mm, and 6 layers are laid on each of the upper, lower surfaces and the vertical ribs of the I-beam.

[0007] After the vertical ribs of the I-beam are laid with forward and reverse C-shaped prepreg, the ply angles are symmetrically distributed.

[0008] In the length direction of the I-beam, the proportion of the 0° ply is not less than 50%.

[0009] In the I-beam, the same ply angles are differentially arranged and evenly dispersed throughout the ply.

[0010] The difference in ply angles between adjacent plies in the I-beam should be < 60°.

[0011] The I-beam includes 0° ply and ±45° ply.

[0012] The skin is made of carbon fiber / epoxy resin prepreg, and 3 layers of the skin are laid, and the ply angles of each layer are designed as 0° and 90°, and the designed ply laying sequence is: [90 / 0 / 90].

[0013] The advantages of the present invention are as follows:

[0014] By strengthening the design of the main load-bearing section of the wing, and since the force on the middle part of the wing is much greater than that on the wing tip, an internal I-beam structure form is designed for the left and right wing bodies. Both the left and right wing tips are made of carbon fiber / epoxy resin solid structures to increase the load-bearing capacity. To ensure the strength at the connection between the wing body and the wing tip, an embedded connection structure is designed. Through the design of the I-beam distribution method and the ply design, and the overall co-curing forming scheme design of the wing, a composite material wing with good surface quality, meeting the requirements of external dimensions, and light weight is finally obtained, verifying the feasibility of the design concept and the forming process. In particular, the structural scheme of using an I-beam plus a thin skin is reasonable, avoiding local stress concentration and being light in weight; adopting the process scheme of first step-by-step curing the I-beam, the central wing box, and the wing tip, and then overall co-curing forming, compared with integral forming, this method has higher precision and better feasibility for the formed parts. Brief Description of the Drawings

[0015] Figure 1 It is a schematic diagram of the external structure of the ultra-light composite material wing structure in the present invention;

[0016] Figure 2 It is a schematic diagram of the distribution state of the I-beam structure in the present invention;

[0017] Figure 3 It is a schematic diagram of the cross-sectional structure of the I-beam structure in the present invention;

[0018] Figure 4 It is a schematic diagram of the laying state structure of the I-beam structure in the present invention;

[0019] Figure 5 It is a curing curve graph of the wing. Detailed Description of the Invention

[0020] The following further elaborates in detail on the forming method of the ultra-light composite material wing structure of the present invention in conjunction with the drawings and the detailed description of the invention.

[0021] Since the structural form of the wing mainly depends on the load transfer path on the wing, in order to make the wing as light as possible while having high compressive strength and high specific strength, the present invention is designed according to the internal structural form of the wing and reasonable material distribution. After verification, the stress concentration areas are mainly at the leading edge and the middle inside the supports at both ends of the wing test, and at the winglet corners. Therefore, the main load-bearing section of the wing is strengthened. And the force on the middle of the wing is much greater than that on the winglets. Based on this, the internal structure of the wing-body is designed: the left and right winglets and the central wing box are made into carbon fiber / epoxy resin solid structures to increase the load-bearing capacity. The designed ultra-light composite material wing structure specifically includes solid structures arranged at the left and right winglets 1 and the central wing box 2, and an I-beam 4 arranged on the wing-body 3 of the wing and embedded in the winglet 1. During specific production, a corresponding placement groove is set at the connection between the winglet and the wing-body 3, and the shape of the placement groove is adapted to the I-beam. The I-beam 4 is embedded in the placement groove. The structure of the I-beam is as Figure 2 shown, and its cross-section presents an I-shape. Due to its unique hollow structure, compared with solid-web members, it has the advantages of reducing weight and saving material costs. The I-beam is made into the full length of the wing-body of the wing, and both sides are embedded in the winglets to strengthen the connection and increase the load-bearing capacity. Through simulation analysis, the internal design of the ultra-light composite material wing is 3 I-beam structures. As can be seen from the above analysis, the ultra-light composite material wing structure includes a central wing box, the wing-bodies on both sides of the central wing box, and the winglets on both sides of the wing-body. The left and right winglets and the central wing box are set as solid structures. The wing-body has I-beams and the I-beams are set to the full length of the wing-body of the wing. At the same time, the I-beams extend and are embedded in the winglets. The specific forming method is: first, the central wing box, the winglets and the I-beams are respectively cured and formed, and then the skin is laid and cured for the second time to form the ultra-light composite material wing structure. That is to say, in the present invention, the central wing box, the winglets and the I-beams are first cured to form independent component structures, and then the skin is laid integrally outside each independent component to form the entire wing mechanism. Among them, the I-beam is made by laying carbon fiber unidirectional prepreg ( Figure 3 ), and 3-6 layers of positive and negative C-shaped prepreg are laid at the vertical ribs of the I-beam (P1 represents the first ply, P2 represents the second ply, and so on in the figure), and then 3-6 layers of prepreg are respectively laid on the upper and lower surfaces (as Figure 4As shown in the figure, finally, it is cured at a temperature of 90 - 135°C and a pressure of 0.3 - 0.6 MPa to form a composite I-beam structure. In this embodiment, the size of the I-beam is 400 mm × 10 mm × h, where h is the height of the vertical rib, and the specific value is determined by the placement position of the I-beam. The thickness of the I-beam is 1.2 mm, and the thickness of the carbon fiber / epoxy single-layer prepreg is about 0.1 mm. Therefore, 6 layers are laid on the upper and lower surfaces and the C-shaped vertical ribs respectively. The ply design should follow the following specific requirements: (1) To avoid the occurrence of tension-bending and tension-shear coupling in the wing after curing, the ply design should consider the symmetric distribution of plies on both sides of the neutral plane; (2) Considering the load-bearing problem, the proportion of plies in the length direction (0°) should be no less than 50%. The so-called length direction refers to the direction from left to right along the wing, and the I-beam is laid along the left-right direction of the wing; (3) To reduce stress concentration, the differential arrangement of the same ply angle (i.e., different angles of 0°, 45°, -45°) should be arranged in a corresponding specific order as much as possible and evenly dispersed throughout the ply; (4) To reduce the interlaminar shear stress between adjacent plies, the ply angle difference between adjacent plies should be < 60°; (5) There should be 0° plies to resist impact loads, and at the same time, a certain amount of ±45° plies should also be provided to quickly spread the impact load. Considering the above requirements comprehensively, the ply sequence of the preform is: [0 / +45 / -45 / -45 / +45 / 0], and the combination method of each preform is as Figure 3 shown.

[0022] In addition, it should also be mentioned that the wing skin in this embodiment is made of carbon fiber / epoxy prepreg. Considering the load-bearing problem and the continuous integrity of the circumferential fibers of the entire skin, the ply angle is designed to be 0° and 90°, and the ply sequence is designed as: [90 / 0 / 90] (0° is still along the left-right direction of the wing). In this embodiment, the carbon fiber / epoxy prepreg used is cured at medium temperature. According to the resin characteristics, gelation occurs at 80°C and linear polymerization occurs. Therefore, it is kept at 80°C for 30 minutes, and after the insulation ends, the temperature is raised to 130°C and kept warm for another 60 minutes. The curing process curve of the complete cured part is as Figure 5 shown. Through the above method, a composite wing with high strength, light weight, and the shape and size meeting the requirements is successfully prepared. Its load-bearing ratio is 2.32 N / g, and finally, the rationality of the composite wing structure design and the forming process plan is verified.

[0023] In addition, in the present invention, the raw materials of the wing composite material adopt the Lianyungang Shenying C100R3724S unidirectional carbon fiber prepreg, and its performance is shown in Table 1.

[0024] Table 1 Performance parameters of C100R3724S unidirectional carbon fiber prepreg Tab l e2 C100R3724S un i direc ti onal carbon fi b er prepreg

[0025] performance parameters

[0026]

[0027] The effects of using the above-mentioned unidirectional carbon fiber prepreg can meet the design requirements.

[0028] Certainly, the above description is not a limitation of the present invention, and the present invention is not limited to the above examples. Changes, modifications, additions or substitutions made by those skilled in the art within the scope of the essence of the present invention should also fall within the protection scope of the present invention.

Claims

1. A method for forming an ultralight composite wing structure, characterized in that: The ultralight composite wing structure comprises a central wing box, wing bodies on the left and right sides of the central wing box, and wing tips on the left and right sides of the wing body, wherein an I-beam is provided in the wing body and the I-beam extends and is embedded in the wing tip, and the molding method is as follows: firstly, the central wing box, the wing tip and the I-beam are solidified and formed respectively, and then the skin is laid and secondary solidified to form the ultralight composite wing structure; wherein, the I-beam is formed by laying carbon fiber unidirectional prepreg, and the vertical ribs of the I-beam are laid with positive and negative C-type prepregs, and then the upper and lower surfaces are laid respectively, and finally, the composite I-beam structure is solidified under the temperature conditions of 90-135°C and the pressure conditions of 0.3-0.6MPa.

2. The method for forming an ultralight composite wing structure according to claim 1, characterized in that: The thickness of the I-beam is 1.2-2.4 mm, the thickness of the carbon fiber / epoxy single-layer prepreg is 0.1-0.2 mm, and 6 layers are laid on the upper and lower surfaces and the vertical ribs of the I-beam.

3. The method for forming an ultralight composite wing structure according to claim 1 or 2, characterized in that: The laying angles of the reinforcements on both sides of the I-beam are symmetrically distributed.

4. The method for forming an ultralight composite wing structure according to claim 3, characterized in that: The proportion of 0° plies in the length direction of the I-beam is not less than 50%.

5. The method for forming an ultralight composite wing structure according to claim 1, 2 or 4, characterized in that: The same ply in the I-beam is arranged at differential angles and evenly dispersed in the entire ply.

6. The method for forming an ultralight composite wing structure according to claim 5, characterized in that: The angle difference between adjacent plies in the I-beam should be less than 60°.

7. The method for forming an ultralight composite wing structure according to claim 1, 2, 4 or 6, characterized in that: The I-beam includes 0° plies and ±45° plies.

8. The method for forming an ultralight composite wing structure according to claim 7, characterized in that: The skin is made of carbon fiber / epoxy resin prepreg, and is laid out in 3 layers. The ply angles of each layer are designed to be 0° and 90°, and the designed ply sequence is: [90 / 0 / 90].

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