Composite unmanned aerial vehicle wing leading edge assembly, thermoforming method and thermoforming tooling

Abstract

The application discloses a composite unmanned aerial vehicle wing leading edge assembly, a hot forming method and a hot forming tool. The wing leading edge assembly comprises a leading edge piece, a sandwich piece and a front beam piece. The leading edge piece comprises an equal-thickness piece and a sharp corner piece. The hot forming tool comprises a leading edge forming die, a front beam forming die and an assembly forming die. The leading edge forming die is provided with a first forming surface and is used for hot forming of the equal-thickness piece. The front beam forming die is provided with a second forming surface and is used for hot forming of the front beam piece. The assembly forming die is provided with a third forming surface and is used for hot forming of the leading edge piece and assembly of the wing leading edge assembly. The wing leading edge assembly is divided into different parts, and the wing leading edge assembly is formed by using a hot pressing process. Compared with a die pressing process, the surface quality of the wing leading edge piece formed by the application is the same, and the strength and rigidity are significantly improved.

Description

Technical Field

[0001] This invention relates to the field of composite material molding technology, and in particular to a composite material unmanned aerial vehicle wing leading edge component, thermoforming method, and thermoforming tooling. Background Technology

[0002] The wing is a crucial component of a drone, but it is also one of the most impact-sensitive parts. The wing not only needs to be lightweight but also possesses high strength to withstand the significant pressure and impact forces during rotation. The wing leading-edge assembly is a vital part of the wing structure; its surface quality and internal performance directly affect the drone's flight speed and stability. The structure of the wing leading-edge assembly is as follows: Figure 1 As shown, the length of the wing leading edge assembly is 0.8~1.4m, and the top of the wing leading edge assembly is a sharp corner with a height h of approximately 40~60mm, and the corner size is relatively small. Existing UAV wing leading edge assemblies are made of composite materials. Composite material structures are used more extensively due to their better specific strength and stiffness, reducing component weight, saving fuel, and lowering operating costs. The current molding method for wing leading edge assemblies is compression molding, which involves molding pre-prepared prepreg or hand lay-up blanks in a molding die, followed by heating and curing.

[0003] The existing methods for molding leading-edge components of wings have the following problems: ① The molding process lacks vacuum pressure during curing, resulting in non-destructive defects within the wing leading-edge components; ② Molded wing leading-edge components have low fiber content, and their mechanical properties only meet standard requirements, making it difficult to improve mechanical strength by increasing fiber content. Therefore, to improve the quality of manufactured wing leading-edge components, new molding methods need to be developed. While composite material parts can be prepared using autoclave processes in the prior art, the wing leading-edge components of this invention have complex structures, small sharp corner dimensions, and difficulty in laying prepreg. Furthermore, there are currently no suitable molding dies for autoclave processes of wing leading-edge components. Summary of the Invention

[0004] The purpose of this invention is to overcome the problem of the lack of molding dies suitable for thermoforming wing leading edge components in the existing autoclave process, and to provide a thermoforming tooling and thermoforming method for composite material UAV wing leading edge components.

[0005] To achieve the above-mentioned objectives, the present invention provides the following technical solution:

[0006] A thermoforming fixture for a composite material unmanned aerial vehicle (UAV) wing leading edge assembly includes a leading edge component, a sandwich component, and a front beam component. The sandwich component is located between the leading edge component and the front beam component. The leading edge component is divided into a component of equal thickness and a component with sharp corners. The leading edge component and the front beam component are made of composite materials, and the sandwich component is made of foam material. The thermoforming fixture includes a leading edge forming mold, a front beam forming mold, and an assembly forming mold.

[0007] The leading edge forming mold is provided with a first forming surface, which matches the outer surface of the equal thickness part. The leading edge forming mold is used for the thermoforming of the equal thickness part blank.

[0008] The front beam forming mold is provided with a second forming surface, which matches the outer surface of the front beam part. The front beam forming mold is used for the thermoforming of the front beam part.

[0009] The component forming mold is provided with a third forming surface; the third forming surface matches the outer surface of the wing leading edge component, and the component forming mold is used for the thermoforming of the leading edge part and the assembly of the wing leading edge component.

[0010] In the above technical solution, the wing leading edge assembly is divided into different parts: the leading edge component, the sandwich component, and the front sparsity component. The leading edge component and the front sparsity component are made of composite materials, while the sandwich component is made of foam material. The thermoforming tooling is also divided into different molds: the leading edge forming mold is used for thermoforming the uniform thickness blank, the front sparsity forming mold is used for thermoforming the front sparsity component, and the component forming mold is used for thermoforming the leading edge component and assembling the wing leading edge assembly. Different components are thermoformed using different molds, and then assembled using the component forming mold. This invention designs a method of dividing the wing leading edge assembly into different parts and using different molds for thermoforming different parts, solving the problem of the complex structure of the wing leading edge assembly. Furthermore, dividing the leading edge component into uniform thickness parts and sharp-corner parts for step-by-step forming solves the problems of small sharp-corner dimensions and difficulty in laying prepreg, thus providing a forming tooling suitable for the thermoforming of wing leading edge components.

[0011] In a preferred embodiment of the present invention, the leading edge forming mold is a male mold, and the leading edge forming mold includes a first base and a first forming part, the first forming part being disposed on the first base, and the first forming part having a first forming surface. Furthermore, the first forming surface matches the bottom surface of the outer surface of the equal-thickness part.

[0012] In a preferred embodiment of the present invention, the front beam forming mold is a male mold, and the front beam forming mold includes a second base and a second forming part, the second forming part being disposed on the second base, and the second forming part having a second forming surface. Furthermore, the second forming surface matches the bottom surface of the outer shape of the front beam component.

[0013] As a preferred embodiment of the present invention, the component molding die includes an upper die and a lower die, the upper die is provided with a fourth molding surface, the lower die is provided with a fifth molding surface, and the fourth molding surface and the fifth molding surface constitute the third molding surface.

[0014] As a preferred embodiment of the present invention, the upper mold is provided with a first mating surface, and the lower mold is provided with a second mating surface. The first mating surface and the second mating surface match each other for mold closing of the upper mold and the lower mold.

[0015] As a preferred embodiment of the present invention, the leading edge and the front beam are made of carbon fiber or glass fiber composite material, and the foam material of the sandwich component is any one of PMI, PU, ​​PVC, PET, or aramid paper honeycomb, which is formed by foaming process and processed into the required shape and size of the sandwich component.

[0016] As a more preferred embodiment of the present invention, the upper mold and the lower mold are provided with matching positioning holes, and at least two positioning holes are provided for positioning the first mating surface and the second mating surface.

[0017] As a preferred embodiment of the present invention, the upper mold is provided with a first threaded hole, and the lower mold is provided with a second threaded hole. The first threaded hole and the second threaded hole are connected by bolts to realize the connection between the upper mold and the lower mold.

[0018] The present invention discloses a thermoforming method for a composite material unmanned aerial vehicle (UAV) wing leading edge assembly, which is prepared using the aforementioned thermoforming mold. The thermoforming method includes the following steps:

[0019] Step S1, preparing the leading edge component: Composite prepreg is sequentially laid on the first forming surface of the leading edge forming mold according to the shape and layup sequence of the equal-thickness component. After laying, a sealed bag is assembled, and the component undergoes a first curing process in a heating and pressurizing device to obtain an equal-thickness component blank. Composite prepreg is sequentially laid on the third forming surface of the component forming mold according to the shape and layup sequence of the pointed corner component. The equal-thickness component blank is then laid inside the component forming mold. After laying, a sealed bag is assembled, and the component undergoes a second curing process in a heating and pressurizing device to obtain the leading edge component.

[0020] Preparation of the front beam component: Composite prepreg is laid sequentially on the second forming surface of the front beam forming mold according to the shape and layup sequence of the front beam component. After laying, the prepreg is assembled and sealed, and then cured in a heating and pressurizing device to obtain the front beam component.

[0021] Step S2: The leading edge component and the sandwich component are sequentially laid in the component molding mold, an adhesive film is laid on the contact surface of the leading edge component and the sandwich component, and then pre-pressing is performed;

[0022] Step S3: Continue to lay the front beam component in the component molding mold. Lay an adhesive film on the contact surface of the front beam component with the leading edge component and the sandwich component, and fill the gap after the front edge component, the sandwich component and the front beam component are combined with the adhesive film. After laying, assemble and seal the bag, and cure it in a heating and pressurizing device to obtain the wing leading edge component.

[0023] In the above technical solution, different molds are used for the thermoforming of different components. The leading edge component is divided into equal-thickness parts and sharp-corner parts for step-by-step forming. First, the equal-thickness part blank is formed using a leading edge forming mold. Then, the prepreg for the sharp-corner part and the leading edge component are laid out and formed using a component forming mold. The front beam component is formed on the front beam forming mold. Finally, the leading edge component, the core component, and the front beam component are assembled using a component forming mold. This invention divides the wing leading edge component into different parts. The core component uses foam material, which reduces the overall weight of the wing leading edge component. It achieves the formation of the wing leading edge component using a thermoforming process. Compared with the molding process, the wing leading edge component formed by this invention has the same surface quality, while its strength and rigidity are significantly improved.

[0024] As a preferred embodiment of the present invention, the component molding mold includes an upper mold and a lower mold. The upper mold has a fourth molding surface, and the lower mold has a fifth molding surface. In step S1, when laying the sharp corner piece, the composite material prepreg is first laid on the fourth molding surface of the upper mold, and then the equal-thickness blank is laid at the corresponding position in the component molding mold. Then, the upper mold and the lower mold are closed. If the upper mold and the lower mold cannot be closed, the carbon fiber composite material prepreg at the sharp corner piece is trimmed until the upper mold and the lower mold are closed. Furthermore, the closing condition of the upper mold and the lower mold is determined by the fit between the first mating surface of the upper mold and the second mating surface of the lower mold. When the distance between the first mating surface and the second mating surface after mold closing is less than or equal to 0.05 mm, it is determined that the upper mold and the lower mold can be closed.

[0025] As a preferred embodiment of the present invention, the process of the leading edge component includes two curing processes. The conditions for the first curing are: vacuuming to -0.9 bar or higher, applying a positive pressure of 6-7 bar, heating to 70-90°C at a rate of 0.5-1.5°C / min and maintaining for 20-60 min. The positive pressure is achieved by filling with N2. The conditions for the second curing are: vacuuming to -0.9 bar or higher, applying a positive pressure of 6-7 bar, heating to 180-220°C and maintaining for 60-150 min. The curing conditions during the preparation of the front beam component are: vacuuming to -0.9 bar or higher, applying a positive pressure of 6-7 bar, heating to 210°C and maintaining for 90 min. Furthermore, during the second curing of the leading edge component and the curing of the front beam component, a stepped heating process can be adopted. For example, the temperature can be increased from room temperature to the first curing temperature at a first heating rate, and maintained at the first curing temperature for 20-40 minutes. The first heating rate is 0.5-1.5℃ / min, and the first curing temperature is 90-110℃. Then, the temperature can be increased from the first curing temperature to the second curing temperature at a second heating rate, and maintained at the second curing temperature for 40-150 minutes. The second heating rate is 0.5-1.5℃ / min, and the second curing temperature is 180-220℃.

[0026] As a preferred embodiment of the present invention, the pre-pressing in step S2 is to perform preliminary hot pressing on the film under vacuum conditions. The pre-pressing conditions are: vacuuming at -0.9 bar or above, applying positive pressure of 6-7 bar, heating to 70~100℃ and maintaining it for 10-30 minutes.

[0027] As a preferred embodiment of the present invention, the curing conditions in step S3 are: vacuuming to -0.9 bar or above, applying positive pressure of 6-7 bar, heating to 90~120°C and maintaining for 30-90 min.

[0028] The present invention also provides a composite material unmanned aerial vehicle (UAV) wing leading edge assembly, which is prepared by the above-described thermoforming method.

[0029] Compared with the prior art, the beneficial effects of the present invention are as follows:

[0030] 1. This invention provides a forming fixture suitable for the thermoforming of wing leading-edge components. The wing leading-edge component is divided into different parts: a leading-edge part, a sandwich part, and a front sparsity part. The leading-edge part and the front sparsity part are made of composite materials, while the sandwich part is made of foam material. The thermoforming fixture is also divided into different molds: a leading-edge forming mold for thermoforming the uniform-thickness blank, a front sparsity forming mold for thermoforming the front sparsity part, and a component forming mold for thermoforming the leading-edge part and assembling the wing leading-edge component. Different parts are thermoformed using different molds, and then assembled using the component forming mold. This invention designs a method of dividing the wing leading-edge component into different parts and using different molds for thermoforming different parts, solving the problem of the complex structure of the wing leading-edge component. Furthermore, dividing the leading-edge part into uniform-thickness parts and sharp-corner parts for step-by-step forming solves the problems of small sharp-corner dimensions and difficulty in laying prepreg.

[0031] 2. The thermoforming method provided by this invention uses different molds for thermoforming different components. The leading edge part is divided into equal-thickness parts and sharp-corner parts for step-by-step forming. First, the equal-thickness part blank is formed using a leading edge forming mold. Then, the prepreg for the sharp-corner part and the leading edge part are laid out and formed using a component forming mold. The front beam part is formed on the front beam forming mold. Finally, the leading edge part, the core part, and the front beam part are assembled using a component forming mold. This invention divides the wing leading edge assembly into different parts. The core part uses foam material, which reduces the overall weight of the wing leading edge assembly. It achieves the formation of the wing leading edge assembly using a hot-pressing process. Compared with the molding process, the wing leading edge part formed by this invention has the same surface quality, while its strength and rigidity are significantly improved. Attached Figure Description

[0032] Figure 1 This is a three-dimensional structural diagram of the leading edge component of the UAV wing of the present invention;

[0033] Figure 2 This is a schematic diagram showing the segmentation of the wing leading edge assembly;

[0034] Figure 3 This is a three-dimensional structural diagram of the leading edge component of the present invention;

[0035] Figure 4 This is a three-dimensional structural diagram of the sandwich component of the present invention;

[0036] Figure 5 This is a three-dimensional structural diagram of the front beam component of the present invention;

[0037] Figure 6 This is a three-dimensional structural diagram of the leading edge forming mold of the present invention;

[0038] Figure 7 This is a three-dimensional structural diagram of the front beam forming mold of the present invention;

[0039] Figure 8This is a three-dimensional structural diagram of the component molding die of the present invention;

[0040] Figure 9 This is a schematic diagram illustrating the assembly process using component molding dies according to the present invention;

[0041] The markings in the diagram are: 1-wing leading edge assembly, 11-leading edge component, 12-sandwich component, 13-front beam component, 131-front beam component outer surface A, 132-front beam component outer surface B, 14-equal thickness component, 141-equal thickness component outer surface A, 15-sharp corner component, 151-sharp corner component outer surface A, 2-leading edge forming mold, 21-first base, 22-first forming part, 23-first forming surface, 3-front beam forming mold, 31-second base, 32-second forming part, 33-second forming surface, 4-component forming mold, 41-upper mold, 42-lower mold, 43-fourth forming surface, 44-fifth forming surface, 45-first mating surface, 46-second mating surface. Detailed Implementation

[0042] To more clearly describe the inventive objectives, technical solutions, and advantages of the specific embodiments of this invention, the solutions in the specific embodiments will be described in detail below with reference to the accompanying drawings. The specific technical solutions involved in the following embodiments are merely for the purpose of clearly and completely describing the innovative technical solutions of this invention. They are only a part of the specific implementation methods that this invention can adopt, not all embodiments, and should not be construed as limiting the innovative solutions of this invention. Any solution that adopts the same inventive concept as this invention should be included within the protection scope of this invention.

[0043] Secondly, the descriptions in the accompanying drawings of the specific embodiments of this invention are merely for the convenience of those skilled in the art to understand the invention. The details shown in the drawings are for the purpose of clearly presenting the technical solution, and should not be construed as including all technical features in the drawings in the specific implementation examples, nor should the details in the drawings be considered as additional limitations on the innovative technical solution of this invention. The components in the various embodiments described and shown in the drawings can be combined and arranged in different configurations. These variations in combination and arrangement should be considered as part of all embodiments of the innovative solution of this invention and included within the scope of protection of this invention.

[0044] In summary, the solutions or descriptions presented in the specific embodiments and accompanying drawings of this invention are not intended to limit the scope of protection claimed, but merely to illustrate selected embodiments / examples to help those skilled in the art understand the relevant innovative solutions. All other equivalent or parallel embodiments obtained by those skilled in the art based on these embodiments without inventive effort are within the scope of protection claimed by this invention.

[0045] It should be noted that, unless otherwise specified, the use of terms such as "upper," "lower," "left," "right," "center," "inner," and "outer" to indicate orientation or positional relationships in the description of specific embodiments of the present invention is based on the orientation or positional relationships shown in the accompanying drawings, or the orientation or positional relationship in which the product / equipment / device is typically placed during use. These terms are merely for the purpose of facilitating the description of the present invention or simplifying the description in specific embodiments, enabling those skilled in the art to quickly understand the solution, and do not indicate or imply that a particular device / component / element must have a specific orientation, or be constructed and operated in a specific positional relationship. Therefore, they should not be construed as limitations on the present invention.

[0046] Furthermore, the use of terms such as "first," "second," "third," etc. in terminology is merely for distinguishing identical or similar components and should not be interpreted as emphasizing or implying the relative importance of a particular component.

[0047] Furthermore, in the description of the technical solution of this invention, unless otherwise explicitly specified / limited / restricted, the terms "set," "install," "connect," and "link" should be interpreted broadly. For example, they can refer to fixed connections, detachable connections, or integral connections; they can refer to common connection methods in the art, such as welding, riveting, bolting, and threaded connections. Such connections can be mechanical or electrical; they can be direct connections or indirect connections through an intermediate medium; and they can refer to the internal communication between two components.

[0048] Example 1

[0049] The structure of the wing leading edge assembly is as follows Figure 1 As shown, the length of the wing leading edge assembly 1 is 0.8~1.4m. The top of the wing leading edge assembly 1 is a sharp corner with a height h of about 40~60mm, a bottom width of 45~50mm, and a top width of 15~20mm. Therefore, the size of the sharp corner is relatively small. Existing wing leading edge assembly molding methods have the problems of low fiber content, no vacuum pressure in the preparation process, and the mechanical properties of the prepared wing leading edge assembly can only meet the standard requirements, making it difficult to improve the mechanical strength. In order to improve the quality of the wing leading edge assembly, this invention designs to use a hot pressing molding process for preparation. However, the current wing leading edge assembly has the problems of complex structure, small sharp corner size, and difficulty in laying prepreg.

[0050] like Figure 2 As shown, the present invention divides the wing leading edge assembly 1 into a leading edge component 11, a sandwich component 12 and a front beam component 13. The sandwich component 12 is located between the leading edge component 11 and the front beam component 13. The leading edge component 11 includes a component of equal thickness 14 and a pointed corner component 15. The leading edge component 11 and the front beam component 13 are made of composite materials, and the sandwich component 12 is made of foam material.

[0051] like Figure 3 As shown, the leading edge 11 is divided into a uniform thickness part 14 and a pointed corner part 15 located at the upper end of the uniform thickness part 14. The uniform thickness part 14 has the same thickness in different parts. This arrangement facilitates the laying of the composite material prepreg in the uniform thickness part 14 during the thermoforming process. The outer surface of the uniform thickness part includes uniform thickness outer surface A and uniform thickness outer surface B. The uniform thickness outer surface A is the outer surface located at the bottom of the uniform thickness part 14. The bottom surface of the pointed corner part 15 is in contact with the top surface of the uniform thickness part 14. The pointed corner part 15 is small in size, and the prepreg is not easy to lay during the thermoforming process. The outer surface of the pointed corner part includes pointed corner part outer surface A and pointed corner part outer surface B. The pointed corner part outer surface B is the outer surface located at the bottom of the pointed corner part 15 and is in contact with part of the outer surface of the uniform thickness part outer surface B. The pointed corner part outer surface A is the outer surface of the side and top of the pointed corner part 15.

[0052] like Figure 4 As shown, the sandwich member 12 is located between the leading edge member 11 and the front spar member 13. By making the sandwich member 12 out of foam material, the overall weight of the wing leading edge assembly 1 can be reduced. Figure 5 As shown, the front beam 13 is located at the bottom of the wing leading edge assembly 1. The outer surface of the front beam includes front beam outer surface A and front beam outer surface B. Front beam outer surface A is the outer surface located at the bottom of the front beam 13. In some embodiments, the corners of the top surface of the front beam 13 are chamfered. The corners of the top surface of the front beam 13 are the corners of the top edge. The chamfering prevents stress concentration during the processing of the front beam 13. On the other hand, it leaves a gap for placing the adhesive film when assembling the leading edge part 11, the sandwich part 12 and the front beam 13, which facilitates the thermoforming process, increases the connectivity between the individually molded parts leading edge part 11, sandwich part 12 and front beam 13, enhances the overall integrity of the finally molded wing leading edge assembly 1, and thus improves the mechanical properties of the parts.

[0053] This embodiment provides a thermoforming fixture for a composite material UAV wing leading edge component 1. The thermoforming fixture includes a leading edge forming mold 2, a front beam forming mold 3, and a component forming mold 4. The leading edge forming mold 2 is provided with a first forming surface 23, which matches the outer surface of a uniform thickness part. The leading edge forming mold 2 is used for thermoforming the uniform thickness part blank. The front beam forming mold 3 is provided with a second forming surface 33, which matches the outer surface of the front beam part. The front beam forming mold 3 is used for thermoforming the front beam part 13. The component forming mold 4 is provided with a third forming surface, which matches the outer surface of the wing leading edge component 1. The component forming mold 4 is used for thermoforming the leading edge part 11 and assembling the wing leading edge component 1.

[0054] like Figure 6As shown, the leading edge forming mold 2 is a male mold. The leading edge forming mold 2 includes a first base 21 and a first forming part 22. The first forming part 22 is disposed on the first base 21. The first forming part 22 is provided with a first forming surface 23. The first forming surface 23 matches the bottom surface of the outer surface of the equal thickness part, that is, the first forming surface 23 matches the outer surface A of the equal thickness part. During the forming process, the composite material prepreg can be directly laid on the first forming surface 23.

[0055] like Figure 7 As shown, the front beam forming mold 3 is a male mold. The front beam forming mold 3 includes a second base 31 and a second forming part 32. The second forming part 32 is disposed on the second base 31. The second forming part 32 is provided with a second forming surface 33. The second forming surface 33 matches the bottom surface of the outer surface of the front beam part, that is, the second forming surface 33 matches the outer surface A of the front beam part. During the forming process, the composite material prepreg can be directly laid on the second forming surface 33.

[0056] like Figure 8 As shown, the component molding mold 4 includes an upper mold 41 and a lower mold 42. The upper mold 41 is provided with a fourth molding surface 43, and the lower mold 42 is provided with a fifth molding surface 44. The fourth molding surface 43 and the fifth molding surface 44 form a third molding surface. The third molding surface matches the outer surface of the side and top of the wing leading edge component 1. When the wing leading edge component 1 is divided, the outer surface of the side and top of the leading edge part 11 is the same as the outer surface of the same position of the wing leading edge component 1. Therefore, the component molding mold 4 can be used to perform hot pressing molding of the leading edge part 11.

[0057] Furthermore, the upper mold 41 is provided with a first mating surface 45, and the lower mold 42 is provided with a second mating surface 46. The first mating surface 45 and the second mating surface 46 are matched with each other for mold closing of the upper mold 41 and the lower mold 42. The upper mold 41 and the lower mold 42 are provided with matching positioning holes, and at least two positioning holes are provided for positioning the first mating surface 45 and the second mating surface 46. The upper mold 41 is provided with a first threaded hole, and the lower mold 42 is provided with a second threaded hole. The first threaded hole and the second threaded hole are connected by bolts to realize the connection between the upper mold 41 and the lower mold 42. When the upper mold 41 and the lower mold 42 are closed, they are first positioned according to the position of the positioning holes so that the first mating surface 45 and the second mating surface 46 are in contact, and then the upper mold 41 and the lower mold 42 are connected by bolts.

[0058] The leading edge forming mold 2, the front beam forming mold 3, and the component forming mold 4 are made of Q235 metal, which has good high temperature resistance and stability.

[0059] Example 2

[0060] This embodiment provides a method for fabricating wing leading edge assembly 1 using the thermoforming tooling of Embodiment 1. The thermoforming method includes the following steps:

[0061] Step S1: Preparation of leading edge component 11: Carbon fiber composite prepreg is laid sequentially on the first forming surface 23 of the leading edge forming mold 2 according to the shape and layup sequence of the equal thickness component 14. After laying, the bag is assembled and sealed, and the first curing is carried out in a hot autoclave to obtain the equal thickness component 14 blank. Carbon fiber composite prepreg is laid sequentially on the third forming surface of the component forming mold 4 according to the shape and layup sequence of the sharp corner component 15. The equal thickness component 14 blank is laid in the component forming mold 4. After laying, the bag is assembled and sealed, and the second curing is carried out in a hot autoclave to obtain the leading edge component 11.

[0062] Preparation of front beam component 13: Carbon fiber composite prepreg is laid sequentially on the second forming surface 33 of the front beam forming mold 3 according to the shape and layup sequence of the front beam component 13. After laying, the prepreg is assembled and sealed, and then cured in a hot autoclave to obtain the front beam component 13.

[0063] Before manual layup in step S1, prepare auxiliary materials, carbon fiber composite prepreg, thermoforming tooling, etc. Auxiliary materials include peelable layer, non-porous release film, breathable felt, vacuum bag cover, sealing strip, etc. Then clean the leading edge forming mold 2, front beam forming mold 3 and component forming mold 4, and apply release agent to the forming surface of thermoforming tooling. Applying release agent makes it easy to demold after the parts are processed.

[0064] The carbon fiber composite prepreg used is T700-12K carbon fiber unidirectional tape and fabric, and the resin is high-temperature curing epoxy resin. The carbon fiber composite prepreg adopts a [0° / 90° / ±45°] cyclic zone design for the layup angle. The prepreg is laid manually. Specifically, during the layup process, after every 2-5 layers, the prepreg is assembled, sealed, and vacuum-pre-compacted. The vacuum pre-compacting conditions are: a vacuum degree of -0.9 bar or higher. Hot compaction can also be used to further remove air from the prepreg, ensuring the quality of the molded parts.

[0065] The component molding mold 4 used in this embodiment includes an upper mold 41 and a lower mold 42. The upper mold 41 is provided with a fourth molding surface 43, and the lower mold 42 is provided with a fifth molding surface 44. During the process of laying the sharp corner piece 15, carbon fiber composite prepreg is first laid on the fourth molding surface 43 of the upper mold 41, and then the blank of the equal thickness piece 14 is laid in the corresponding position in the component molding mold 4. Then the upper mold 41 and the lower mold 42 are closed. If the upper mold 41 and the lower mold 42 cannot be closed, the carbon fiber composite prepreg at the sharp corner piece 15 is trimmed until the upper mold 41 and the lower mold 42 are closed. The closing condition of the upper mold 41 and the lower mold 42 is determined by the fit between the first mating surface 45 of the upper mold 41 and the second mating surface 46 of the lower mold 42. When the distance between the first mating surface 45 and the second mating surface 46 after the mold is closed is less than or equal to 0.05mm, it is determined that the upper mold 41 and the lower mold 42 can be closed. In the actual manufacturing process, a 0.02mm feeler gauge is usually used to judge the mold closing condition of the upper mold 41 and the lower mold 42. When the 0.02mm feeler gauge can be placed between the first mating surface 45 and the second mating surface 46, the upper mold 41 and the lower mold 42 can be closed.

[0066] The fabrication of the leading edge component 11 involves two curing processes. The first curing conditions are: vacuum above -0.9 bar, applying a positive pressure of 6-7 bar, heating to 80°C at a rate of 0.5-1.5°C / min and maintaining the temperature for 30 min. The positive pressure is achieved by filling with N2. The second curing conditions are: vacuum above -0.9 bar, applying a positive pressure of 6-7 bar, heating to 180-220°C and maintaining the temperature for 60-150 min. The fabrication of the front beam component 13 involves: vacuum above -0.9 bar, applying a positive pressure of 6-7 bar, heating to 210°C and maintaining the temperature for 90 min. During the second curing of the leading edge component 11 and the curing of the front beam component 13, a stepped heating process can be adopted. For example, the temperature can be increased from room temperature to the first curing temperature at the first heating rate, and maintained at the first curing temperature for 20-40 minutes. The first heating rate is 0.5-1.5℃ / min, and the first curing temperature is 90-110℃. Alternatively, the temperature can be increased from the first curing temperature to the second curing temperature at the second heating rate, and maintained at the second curing temperature for 40-150 minutes. The second heating rate is 0.5-1.5℃ / min, and the second curing temperature is 180-220℃.

[0067] The assembly and sealing process includes the following steps: A peelable layer, a non-porous release film, and a breathable felt are sequentially placed on the surface of the laid-up carbon fiber composite prepreg. Then, a vacuum bag is used to cover the bag, and a high-temperature resistant sealing strip is placed at the seal to ensure a tight seal inside the vacuum bag. After assembly and sealing, a well-sealed vacuum system is formed.

[0068] After curing in the autoclave, the air is used for cooling at a rate of 1.5℃ / min. During the cooling process, the vacuum and pressure are maintained until the cooling is completed and then the pressure is released. This process eliminates thermal deformation and warping during heating, reduces internal stress, eliminates deformation caused by residual stress, and ensures the forming dimensions and quality of the wing leading edge assembly 1.

[0069] Step S2: Lay the leading edge 11 and the sandwich 12 sequentially in the component molding mold 4, lay an adhesive film on the contact surface of the leading edge 11 and the sandwich 12, and then perform pre-pressing.

[0070] During the installation process, an adhesive film is laid between the contact surfaces of the leading edge component 11 and the sandwich component 12. The adhesive film is made of epoxy resin, and its thickness and size are cut according to the combined dimensions of the leading edge component 11 and the sandwich component 12. The pre-pressing in step S2 involves preliminary hot pressing of the adhesive film under vacuum conditions. The pre-pressing conditions are: vacuuming to -0.9 bar or higher, applying a positive pressure of 6-7 bar, and heating to 80°C and maintaining it for 10-30 minutes.

[0071] Step S3: Continue laying the front beam component 13 inside the component molding mold 4. A film is laid on the contact surfaces of the front beam component 13 with the front edge component 11 and the sandwich component 12, such as... Figure 9 As shown, the gaps between the leading edge component 11, the sandwich component 12 and the front beam component 13 are filled with adhesive film. After the film is laid, the assembly is sealed and cured in an autoclave to obtain the wing leading edge component 1.

[0072] The front beam 13 is laid on the front edge 11 and the sandwich member 12. An adhesive film is applied to the contact surfaces between the front beam 13 and the front edge 11, and between the front beam 13 and the sandwich member 12. The thickness and size of the adhesive film are cut according to the combined dimensions of the front edge 11, sandwich member 12, and front beam 13. Since the top of the front beam 13 is chamfered, gaps will exist at the joints of the front edge 11, sandwich member 12, and front beam 13 after assembly. Figure 2 The gaps at the junctions of the front edge component 11, the sandwich component 12, and the front beam component 13 in the black shaded area are filled with cut adhesive film.

[0073] The curing conditions in step S3 are as follows: vacuuming to -0.9 bar or higher, applying a positive pressure of 6-7 bar, heating to 90~120℃ and maintaining it for 30-90 minutes. After curing, CNC machining is performed to the theoretical dimensions of the outer surface of the wing leading edge assembly 1, i.e., cutting the edge allowance to the part line to obtain the wing leading edge assembly 1.

[0074] For those skilled in the art, when understanding the solutions described in the specific embodiments of the present invention, conventional technical manuals in the field can be consulted. At the same time, appropriate understandings or adjustments can be made to the above-mentioned terms to deduce the same or similar technical solutions without creative effort.

[0075] The above embodiments describe only the basic principles, main features and / or 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 the description of the invention content in the specification are only the principles or specific cases of the present invention. Without departing from the essence of the innovative idea of ​​the present invention, there are various changes and improvements to the innovative solution of the present invention, and all such changes and improvements fall within the scope of protection claimed by the present invention.

Claims

1. A thermoforming tooling for the leading edge component of a composite material unmanned aerial vehicle wing, characterized in that, The wing leading edge assembly includes a leading edge component, a sandwich component, and a front spar component. The sandwich component is located between the leading edge component and the front spar component. The leading edge component includes a component of equal thickness and a component with sharp corners. The leading edge component and the front spar component are made of composite materials. The sandwich component is made of foam material. The thermoforming tooling includes a leading edge forming mold, a front spar forming mold, and a component forming mold. The leading edge forming mold is provided with a first forming surface, which matches the outer surface of the equal thickness part. The leading edge forming mold is used for the thermoforming of the equal thickness part blank. The front beam forming mold is provided with a second forming surface, which matches the outer surface of the front beam part. The front beam forming mold is used for the thermoforming of the front beam part. The component forming mold is provided with a third forming surface; the third forming surface matches the outer surface of the wing leading edge component; the component forming mold is used for the thermoforming of the leading edge part and the assembly of the wing leading edge component. The component molding die includes an upper die and a lower die. The upper die is provided with a fourth molding surface, and the lower die is provided with a fifth molding surface. The fourth molding surface and the fifth molding surface together form the third molding surface. When laying the sharp corner piece, first lay the composite prepreg on the fourth forming surface of the upper mold, then lay the equal thickness blank in the corresponding position in the component forming mold, and then close the upper mold and the lower mold. If the upper mold and the lower mold cannot be closed, the carbon fiber composite prepreg at the sharp corner piece is trimmed until the upper mold and the lower mold are closed.

2. The thermoforming tooling for the leading edge assembly of a composite material UAV wing according to claim 1, characterized in that, The leading edge forming mold includes a first base and a first forming part, the first forming part is disposed on the first base, and the first forming part is provided with the first forming surface.

3. The thermoforming tooling for the leading edge assembly of a composite material UAV wing according to claim 1, characterized in that, The front beam forming mold includes a second base and a second forming part, the second forming part is disposed on the second base, and the second forming part is provided with a second forming surface.

4. The thermoforming tooling for the leading edge assembly of a composite material UAV wing according to claim 1, characterized in that, The upper mold is provided with a first mating surface, and the lower mold is provided with a second mating surface. The first mating surface and the second mating surface match each other for mold closing of the upper mold and the lower mold.

5. The thermoforming tooling for the leading edge assembly of a composite material UAV wing according to claim 4, characterized in that, The upper mold and the lower mold are provided with matching positioning holes, and at least two positioning holes are provided for positioning the first mating surface and the second mating surface.

6. A method for thermoforming a composite material unmanned aerial vehicle (UAV) wing leading edge assembly, comprising the following steps: (Thermoforming tooling for the composite material UAV wing leading edge assembly as described in any one of claims 1-5) Step S1, preparing the leading edge component: Composite prepreg is sequentially laid on the first forming surface of the leading edge forming mold according to the shape and layup sequence of the equal-thickness component. After laying, a sealed bag is assembled, and the component undergoes a first curing process in a heating and pressurizing device to obtain an equal-thickness component blank. Composite prepreg is sequentially laid on the third forming surface of the component forming mold according to the shape and layup sequence of the pointed corner component. The equal-thickness component blank is then laid inside the component forming mold. After laying, a sealed bag is assembled, and the component undergoes a second curing process in a heating and pressurizing device to obtain the leading edge component. Preparation of the front beam component: Composite prepreg is laid sequentially on the second forming surface of the front beam forming mold according to the shape and layup sequence of the front beam component. After laying, the prepreg is assembled and sealed, and then cured in a heating and pressurizing device to obtain the front beam component. Step S2: The leading edge component and the sandwich component are sequentially laid in the component molding mold, an adhesive film is laid on the contact surface of the leading edge component and the sandwich component, and then pre-pressing is performed; Step S3: Continue to lay the front beam component in the component molding mold. Lay an adhesive film on the contact surface of the front beam component with the leading edge component and the sandwich component, and fill the gap after the front edge component, the sandwich component and the front beam component are combined with the adhesive film. After laying, assemble and seal the bag, and cure it in a heating and pressurizing device to obtain the wing leading edge component.

7. The thermoforming method for the leading edge component of a composite material UAV wing according to claim 6, characterized in that, The process of preparing the leading edge component includes two curing processes. The conditions for the first curing are: vacuuming to -0.9 bar or higher, applying a positive pressure of 6 to 7 bar, heating to 70 to 90 °C at a rate of 0.5 to 1.5 °C / min and maintaining the temperature for 20 to 60 min. The positive pressure is achieved by filling with N2. The conditions for the second curing are: vacuuming to -0.9 bar or higher, applying a positive pressure of 6 to 7 bar, heating to 180 to 220 °C and maintaining the temperature for 60 to 150 min. The curing conditions during the preparation of the front beam component are: vacuuming to -0.9 bar or higher, applying a positive pressure of 6 to 7 bar, heating to 210 °C and maintaining the temperature for 90 min.

8. A composite material unmanned aerial vehicle (UAV) wing leading edge assembly, characterized in that, The wing leading edge assembly is prepared by the thermoforming method for composite material UAV wing leading edge assembly as described in any one of claims 6-7.

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