A molding method of a composite unmanned aerial vehicle body

Abstract

The application discloses a kind of composite unmanned aerial vehicle fuselage moulding forming methods, it includes the following steps: S1: flat plate and mould cleaning;S2: the laying of part;After laying with core mould / male die, it is assembled in the form of pre-pressing: the way of turning up after lower mould laying is changed to laying on core mould, according to drawing requirement, laying and the laying of reinforcing sheet are carried out, when laying reinforcing sheet, first half of core mould is laid, then core mould is turned over, after turning over, laying is completed, and it is packed and pre-vacuumized;S3: mould closing pre-pressing;S4: solidification;S5: demoulding;S6: edge cutting drilling;S7: inspection.The application adopts the form of pre-pressing after laying with core mould (male die), and the way of turning up after lower mould laying is changed to laying on core mould, which solves the problems of difficult turning up operation and pre-extraction during laying, reduces the difficulty of laying, and improves the laying efficiency.

Description

Technical Field

[0001] This invention relates to the field of composite material drone fuselage molding technology, specifically a molding method for composite material drone fuselage. Background Technology

[0002] Compression molding is a widely used process in the manufacturing of composite material drone fuselages. Traditionally, prepreg sheets are directly laid inside the lower mold cavity. The fuselage structure is built by laying the sheets as a whole and reinforcing them locally, with flanged sheets reserved in the lower mold. After all the sheets are laid, the mandrel assembly is assembled, the flanged sheets are wrapped around the outside of the mandrel, end blocks are installed, and the mold is closed with the upper mold before finally being pressed and cured. However, this process has revealed several insurmountable technical defects in practical applications. When the lower mold cavity has locating pins and sharp edges, the prepreg sheets are easily cut by the hard protrusions and sharp edges of the mold during the vacuum pre-compaction stage, causing internal fractures of the carbon fiber and directly weakening the mechanical properties of the part. Simultaneously, the multiple layers of sheets stacked and laid in the lower mold cavity adhere to each other, significantly increasing the resistance and difficulty of subsequent flanged operations. Furthermore, the mandrel assembly is only assembled after all layers are laid, making the positioning and alignment process cumbersome, time-consuming, and labor-intensive, severely restricting production efficiency. In the mold-closing stage, the end blocks of the machine body have inherent design flaws. The protruding bosses at the circular holes prevent the blocks from fully fitting and pressing against the mold end face. This results in uneven force during mold closing and locking, leading to insufficient pressure at both ends of the machine body and causing quality problems such as under-pressure at the ends, poor compaction, and missing material on the surface. The demolding stage also faces severe challenges. The segmented core mold is not equipped with a dedicated demolding ejection device. The core mold is tightly fitted to the composite material part, making disassembly and separation difficult and easily scratching or bumping the surface of the finished machine body during demolding. In addition, although the carbon fiber prepreg used in the existing technology undergoes plasma surface treatment to improve the interfacial bonding effect, the active groups on the carbon fiber surface gradually decay over time. The micro-rough structure is easily worn away and smoothed under cyclic alternating loads, resulting in insufficient interlaminar shear strength and difficulty in maintaining high mechanical properties stably over a long period of time. Summary of the Invention

[0003] The purpose of this invention is to provide a molding method for composite material drone fuselages, which solves the problems of easy cutting during prepreg laying, uneven mold force leading to end defects, and difficulty in demolding, and significantly improves production efficiency and product mechanical properties.

[0004] This invention provides a method for molding a composite material unmanned aerial vehicle (UAV) fuselage, comprising the following steps: S1: Cleaning of flat plates and molds Clean the tiling work surface, and use acetone to clean each section of the upper mold, lower mold, core mold, and end blocks. First, apply release agent to each individual core mold section, upper mold, and lower mold cavity, and then assemble the core mold assembly to ensure that the mold is clean and facilitates subsequent demolding operations. S2: Parts Placement Using the assembled core mold as a reference, the prepreg is laid first. Then, the core mold is flipped over to complete the overall outer perimeter laying. After laying, the core mold is vacuum-packed and pre-pressed. Reinforcing sheets are laid according to the outline of the core mold. In the area where the thickness of the machine body gradually changes, the prepreg layers are laid linearly. S3: Mold closing pre-compression The core mold assembly covered with prepreg is closed with the upper and lower molds. A fluoroplastic release film is inserted on the outside of the prepreg. The pressure block and end stop are assembled. The locking bolts, which are coated with grease that can withstand the curing temperature, are tightened. The whole assembly is pre-pressed for 15 to 20 minutes. The end stop adopts a flat and fitted structure with the round hole boss removed to ensure uniform force when the mold is closed and locked. S4: Curing and molding Remove the release liner to allow direct pressing and bonding between the mold cavity and the prepreg. Lock the pressure block and stop block again to complete the secondary mold closing. Place the entire mold into the press for molding and curing. Heat to 65–75°C at a rate of ≤2°C / min, hold for 30–40 minutes, and then apply pressure to 3.0–3.5 MPa. Check the mold closing gap and control it to <0.1 mm. If the gap is >0.1 mm, apply pressure gradually at 0.5 MPa until the gap meets the standard. Once the gap meets the requirements, continue heating to 120–130°C at ≤2°C / min, hold for 120–180 minutes, and then cool to below 60°C at a uniform rate of ≤2°C / min. Remove the mold from the press. S5: Demolding operation Disassemble the stop block, pressure block, upper mold, multiple connecting plates and core mold sections in sequence, and use the core mold section matching demolding structure to successfully demold the product. S6: Edge drilling Use a pneumatic cutting machine to roughly cut the edge along the theoretical outline of the part, leaving a machining allowance of 0.5 to 1 mm. Grind to remove burrs and excess material to the cutting line position. S7: Finished Product Inspection The external dimensions and surface appearance quality of the molded parts are inspected, and the inspection results are fully recorded.

[0005] Furthermore, the preparation method of the prepreg in S2 is as follows: By weight, 100–150 parts of epoxy resin are heated to 120–130°C, and 3–5 parts of polyethersulfone are added and thoroughly mixed. After cooling to 60–70°C, 7–15 parts of carboxyl-terminated liquid nitrile rubber and 1–3 parts of silane coupling agent are added and stirred until homogeneous. Then, 3–10 parts of nanofiller are added, followed by 0.1–2 parts of toughening agent, and stirring continues. Finally, 15–30 parts of curing agent are added to obtain a mixed resin matrix. The mixed resin matrix is ​​degassed under vacuum conditions of -0.09 to -0.095 MPa for 20–30 minutes. After degassed, the resin matrix is ​​impregnated with composite carbon fiber cloth using a hot-melt prepreg process, controlling the final resin content to 35–45%, to obtain a prepreg. The reinforcing sheet is prepared by cutting the above prepreg to the required dimensions.

[0006] Further, the epoxy resin is selected from at least one of bisphenol A type epoxy resin and bisphenol F type epoxy resin; the curing agent is selected from at least one of m-phenylenediamine, phthalic anhydride, and methyltetrahydrophthalic anhydride; the toughening agent is compounded from hydroxyl-terminated polybutadiene and dibutyl phthalate in a mass ratio of (1-3):1, and the total amount of the compounded toughening agent added to the resin matrix is ​​0.1-2 parts by weight; the nanofiller is selected from at least one of nano-silica, carbon nanotubes, and nano-graphene, and the average particle size of the nanofiller is 10-50 nm; the silane coupling agent is selected from at least one of vinyltris(β-methoxyethoxy)silane, N-(β-aminoethyl)-γ-aminopropylmethyldimethoxysilane, and γ-aminopropylmethyldiethoxysilane.

[0007] Furthermore, the preparation steps of the composite carbon fiber cloth are as follows: 1) Mix concentrated hydrochloric acid with deionized water and stir, add tetrabutyl titanate dropwise and mix well, then add anhydrous ethanol and stir thoroughly to prepare a precursor solution. 2) The modified carbon fiber cloth was immersed in the precursor solution, stirred and soaked, and then transferred to the reactor and kept at a temperature for 10-15 hours. After the reaction was completed, it was naturally cooled to room temperature, and then centrifuged, washed and dried to obtain the composite carbon fiber cloth.

[0008] Further, in step 1), the volume ratio of concentrated hydrochloric acid, deionized water, tetrabutyl titanate, and anhydrous ethanol is (50-100):(40-80):(1.5-2.5):(10-20); the concentration of concentrated hydrochloric acid is 12-15 mol / L.

[0009] Further, in step 2), the solid-liquid ratio of the modified carbon fiber cloth to the precursor liquid is 1g:(30-50)mL; the temperature of the reactor is 110-130℃, and the heating rate is controlled at 1-2℃ / min.

[0010] Furthermore, the preparation steps of the modified carbon fiber cloth are as follows: 1) Carbon fibers are placed in an acetone solution and heated under reflux at 70-80°C for 20-30 hours. After treatment, the carbon fibers are repeatedly washed with deionized water to remove residual acetone and dried to obtain desized carbon fiber cloth. Electrochemical oxidation etching is performed using dilute sulfuric acid as electrolyte, the desized carbon fiber cloth as anode, and graphite as cathode. The electrode spacing is 2-4 cm, the anode current density is 0.3-0.6 mA / cm², and the treatment time is 10-20 s. After removal, the carbon fibers are washed with deionized water until the filtrate is neutral and dried to obtain pretreated carbon fiber cloth. 2) Dissolve bismuth nitrate in deionized water and stir to prepare solution A, and dissolve sodium tungstate in deionized water and stir to prepare solution B; pour solution B into solution A according to the volume ratio of solution A: solution B 1: (1.5~1.8), and stir and mix at a constant temperature of 40~43℃ to obtain the reaction solution; 3) After the pretreated carbon fiber cloth is fully immersed in the reaction solution, it is transferred to the reaction vessel and hydrothermally reacted at 160-170℃ for 15-20 hours. After the reaction is completed, it is naturally cooled to room temperature, and after centrifugation, washing and drying, the modified carbon fiber cloth is obtained.

[0011] Further, in step 1), the carbon fiber cloth is at least one of T300, T700, T800 carbon fiber plain weave fabric and twill weave fabric, with an areal density of 160-220 g / m²; the acetone heating reflux temperature is 70-80°C; and the mass fraction of dilute sulfuric acid is 0.3-1.0 wt%.

[0012] Further, in step 2), the ratio of bismuth nitrate to deionized water in solution A is (1.0-1.5) g: (30-50) mL; and the ratio of sodium tungstate to deionized water in solution B is (0.4-1.0) g: (45-100) mL.

[0013] Further, in step 3), the solid-liquid ratio of the pretreated carbon fiber cloth to the reaction solution is 1g:(50-80)mL; the hydrothermal reaction temperature is 160-170℃.

[0014] The present invention provides a compression molding method for composite material drone fuselage, which has the following advantages: The compression molding method of the present invention avoids the prepreg being cut in the lower mold cavity by using the assembly core mold as a reference for the laying operation, and uses the end block flat fitting structure to ensure uniform mold force. At the same time, it achieves smooth demolding by using the core mold segment matching demolding structure. It has the advantages of solving the problems of prepreg cutting, uneven mold force, and demolding difficulty, and significantly improving production efficiency and product mechanical properties. Detailed Implementation

[0015] The technical solutions in the embodiments of the present invention will be clearly and completely described below.

[0016] Traditional composite material compression molding processes for drone fuselages are prone to sheet tearing and fiber damage during prepreg application to the lower mold, affecting the mechanical properties of the finished part. Multiple sheets adhering together make flanging difficult, core mold assembly and alignment cumbersome, and production efficiency low. Existing block structure design flaws often result in insufficient pressure at both ends of the fuselage, leading to underpressure and poor compaction. Furthermore, the lack of specialized equipment during demolding makes the finished product susceptible to scratches.

[0017] In response, this invention proposes a molding method for composite material drone fuselages, including steps such as cleaning of flat plates and molds, laying of parts, pre-pressing of molds, curing and molding, demolding, edge trimming and drilling, and finished product inspection, aiming to solve the above-mentioned technical problems.

[0018] To facilitate a clear understanding of the technical solution, key terms in this embodiment are defined as follows: Composite material UAV fuselage refers to a lightweight, high-strength component made of multiple materials that bears the load of the main UAV structure; Compression molding method refers to a composite material processing technology in which prepreg is placed in a mold cavity and cured by heating and pressurizing; Prepreg refers to sheet-like composite material raw material that has been pre-impregnated with reinforcing fibers and pre-cured; Upper mold, lower mold, and core mold are the core mold components for compression molding. The upper and lower molds form the external cavity of the part, and the core mold forms the internal cavity structure of the part; Stop blocks, pressure blocks, and locking bolts are auxiliary components for mold positioning and locking, used to ensure mold closing accuracy and pressure uniformity; Fluoroplastic release film is a special release medium for compression molding, possessing high temperature resistance and anti-adhesion properties; Press is a special curing equipment that provides molding temperature and pressure; Pneumatic cutting machine is a special processing equipment for precise trimming of composite material parts.

[0019] The specific implementation method of the composite material unmanned aerial vehicle fuselage of the present invention is as follows: Cleaning the mold and mold is a crucial pre-molding process. Its core purpose is to remove oil, dust, and impurities from the mold surface, ensuring the surface quality of the parts and laying the foundation for smooth demolding. The process involves first cleaning the work surface, then using acetone to thoroughly clean each section of the upper mold, lower mold, core mold, and end stops, completely removing any residual contaminants. After cleaning, a release agent is evenly applied to each core mold section and the inner cavity of the upper and lower molds. Finally, the entire core mold assembly is assembled. This step-by-step process ensures uniform cleanliness across all areas of the mold, preventing issues such as the inability to apply release agent to the internal cavities after assembly, and eliminating defects like part contamination and demolding difficulties from the outset.

[0020] The component layup process directly determines the structural integrity and mechanical properties of the part, and is one of the core improvements of this invention. Traditional processes involve directly laying prepreg in the lower mold cavity. This cavity contains hard protrusions such as locating pins and sharp-angled steps. During vacuum pre-compaction, the prepreg is pulled tightly against these hard structures by negative pressure, easily tearing the sheets and damaging the carbon fibers. Furthermore, multiple layers of sheets adhere to each other after being stacked in the closed cavity, requiring subsequent flanging to cover the mandrel. This flanging operation is subject to significant resistance due to the limited space in the mold, making it difficult and cumbersome to assemble and align the mandrel, resulting in low production efficiency. This invention abandons the traditional method of laying prepreg inside the lower mold cavity, instead using the smooth outer surface of the assembled mandrel as the sole laying reference for external prepreg laying. The outer contour of the core mold is a smooth and continuous surface that fits the inner cavity of the machine body. The core mold body does not have hard protrusions such as positioning pins or sharp corners. During the prepreg laying and vacuum pre-extraction compaction process, it only contacts the smooth outer peripheral surface of the core mold, thus isolating the hard and sharp components of the upper and lower molds in space. Therefore, it can completely avoid damage to the material sheet caused by the hard structure of the mold.

[0021] In practice, the prepreg is first laid on one half of the mandrel, then flipped over to complete the full coverage of the outer perimeter. This laying process is a continuous, integral wrapping process, with the prepreg sheets laid before and after flipping the mandrel overlapping each other to form a seamless, complete wrapping layer, ensuring the overall continuity and mechanical integrity of the fuselage layer. After laying, the material is vacuum-packed and pre-compacted to completely remove air between the prepreg layers and improve the layer density. Based on the stress requirements of the fuselage structure, local reinforcing sheets are precisely laid along the outline of the mandrel. For areas with gradual thickness changes in the fuselage, the number of prepreg layers is adjusted linearly to match the mechanical load-bearing requirements of the gradual structure of the fuselage. This invention directly completes the overall wrapping on the outer perimeter of the mandrel, completely eliminating the traditional process of "stacked laying in the lower mold cavity + subsequent edge wrapping," thus eliminating the problems of multi-layer sheet stacking and adhesion and difficult edge-flipping operations from the source of the process. It eliminates the need for subsequent edge-flipping operations, significantly simplifying the assembly process and improving laying accuracy and production efficiency.

[0022] There are different ways to flip the core mold. For example, the operator can hold the end of the core mold directly or use the pre-set gripping structure on the core mold to flip the core mold 180° along its axis. Another example is that the operator can push the core mold to rotate 180° around its axis. For example, for heavy core molds, the operator can use hoisting equipment to lift the core mold and flip it 180° in the air.

[0023] The pre-pressing process is used to initially compact and shape the prepreg, eliminating defects such as loose layers and air bubbles, and providing a foundation for secondary curing. The core mold assembly covering the prepreg is precisely aligned with the upper and lower molds to complete the initial mold closing. A fluoroplastic release film is placed on the outside of the prepreg to effectively prevent the prepreg resin from sticking to the mold. Subsequently, the pressure block and end stops are assembled. Lubricant sufficient to withstand the curing temperature is applied to the surface of the locking bolts before tightening to ensure mold locking stability. Overall, a constant temperature pre-pressing process is performed for 15–20 minutes to achieve uniform compaction of the prepreg. Addressing the defects of poor fit and insufficient end pressure caused by the round hole boss in traditional stop blocks, this invention adopts a flat, fitted stop block structure without the round hole boss, ensuring complete fit between the stop block and the mold end face. After mold closing and locking, the force is uniform, completely solving the molding defects of insufficient pressure and poor density at both ends of the machine body.

[0024] Curing and molding is the core process for cross-linking and curing the resin matrix and shaping the parts, directly determining the final mechanical properties and structural stability of the machine body. After pre-pressing, the mold locking structure is loosened, the internal fluoroplastic release film is removed, and the mold cavity is directly and tightly fitted to the surface of the prepreg. The pressure block and end stop are then locked again to complete the second precise mold closing. The entire mold is then smoothly fed into the press for segmented temperature and pressure controlled curing and molding. During the curing process, the heating rate is strictly controlled to ≤2℃ / min. First, the temperature is uniformly raised to 65~75℃ and held for 30~40min to ensure uniform resin melting and flow and full impregnation of the fibers. Then, the pressure is slowly applied to 3.0~3.5MPa. The mold closing gap is monitored in real time and strictly controlled to <0.1mm. If the gap exceeds the standard, pressure is gradually increased in increments of 0.5MPa until the gap meets the standard, preventing problems such as loose parts and material shortages caused by excessive mold closing gap. After the gap meets the standard, continue to heat to 120-130℃ at a rate of ≤2℃ / min, and maintain constant temperature and pressure for curing for 120-180 minutes to ensure complete cross-linking and curing of the resin. After curing, cool down to below 60℃ at a uniform rate of ≤2℃ / min, release the pressure of the press, and remove the entire mold. Slow cooling can effectively release the internal stress of the parts and avoid deformation and cracking of the machine body.

[0025] Demolding is used to smoothly separate the cured part from the mold and core mold. Traditional segmented core molds are tightly fitted to composite material parts, lacking dedicated demolding aids, making disassembly difficult and prone to scratches and damage to the finished product. This invention employs a step-by-step, orderly demolding process, sequentially disassembling the end blocks, pressure blocks, upper mold, and mold connecting plate, and finally disassembling each segment of the core mold one by one. By ejecting each core mold segment individually, the internal support constraints of the part are gradually released, smoothly completing the demolding of the finished product. This step-by-step demolding method distributes force evenly, maximizing the protection of the finished product's structural integrity and avoiding demolding damage defects.

[0026] Edge trimming and drilling are finishing processes in the post-processing of parts, used to correct forming allowances and ensure dimensional accuracy. A pneumatic cutting machine is used for rough cutting along the theoretical contour line of the machine body, leaving a uniform machining allowance of 0.5–1 mm to avoid damaging the effective structure of the part. After rough cutting, fine grinding removes burrs and excess material from the edges of the part, precisely trimming it to the theoretical cutting line position to ensure a regular machine body contour and accurate dimensions.

[0027] Finished product inspection is the final step in ensuring product quality. A comprehensive inspection is conducted on the formed drone fuselage, focusing on verifying the accuracy of the external dimensions and the quality of the surface appearance. Molding defects such as bubbles, cracks, scratches, and missing materials are checked and all inspection data are fully recorded to achieve product quality traceability and ensure that the finished product meets the usage standards.

[0028] In summary, this invention completely solves the problems of fiber damage, layup adhesion, and inefficient assembly caused by traditional lower mold laying through the core mold reference laying process; eliminates defects such as insufficient pressure and density at the fuselage end through the flat, boss-free block structure; improves the overall density and structural uniformity of the part through segmented curing and precise control of gap filling pressure; and avoids scratches and bumps on the finished product through the step-by-step ejection demolding process, significantly improving the molding quality, dimensional accuracy, and production efficiency of composite material UAV fuselage.

[0029] Based on the aforementioned fuselage compression molding process, conventional carbon fiber prepreg resin matrices exhibit significant performance limitations. Ordinary resin systems struggle to balance post-curing mechanical strength, toughness, and fiber-interfacial bonding stability, leading to interlaminar shear failure and fatigue damage in UAV fuselages under complex alternating loads. Furthermore, the resin matrix suffers from poor rheological properties and degassing, resulting in internal porosity defects that negatively impact the overall reliability of the fuselage. To address this, this invention further optimizes the prepreg preparation process, constructing a multi-component synergistic toughening and strengthening high-performance resin matrix system. The specific preparation method is as follows: According to the weight ratio, 100-150 parts of epoxy resin are heated to 120-130℃ and stirred thoroughly. Then, 3-5 parts of polyethersulfone are added and mixed evenly. After cooling to 60-70℃, 7-15 parts of carboxyl-terminated liquid nitrile rubber and 1-3 parts of silane coupling agent are added sequentially and stirred until homogeneous. Then, 3-10 parts of nanofiller and 0.1-2 parts of compound toughening agent are added and stirred continuously. Finally, 15-30 parts of curing agent are added to obtain a uniform and stable mixed resin matrix. The prepared resin matrix is ​​placed in a vacuum environment and degassed at a vacuum degree of -0.09 to -0.095 MPa for 20-30 minutes to completely remove microbubbles in the system. After degassed, the resin matrix is ​​uniformly impregnated with composite carbon fiber cloth using a hot melt prepreg process. The final resin content is precisely controlled within the range of 35-45% to obtain a high-performance modified prepreg. The reinforcing sheets required for the machine body are all prepared by cutting this prepreg as needed.

[0030] The synergistic effect of the resin matrix components comprehensively enhances the overall performance of the prepreg: Epoxy resin, as the main matrix, is selected from bisphenol A or bisphenol F types, possessing excellent adhesion, chemical stability, and molding processability, providing the composite material with basic structural strength and rigidity. Polyethersulfone, as a thermoplastic modifier, effectively improves the curing brittleness of epoxy resin, enhances matrix toughness and processing fluidity, and alleviates internal stress during resin curing. Carboxyl-terminated liquid nitrile rubber can chemically bond with epoxy resin, forming a micro-phase separation structure during curing, effectively absorbing impact energy, inhibiting crack propagation, and significantly improving the material's impact resistance and fatigue resistance. Silane coupling agent, acting as an interface modification bridge, binds to the carbon fiber surface at one end and grafts the resin matrix at the other, significantly improving the interfacial bonding strength between the fiber and the matrix and preventing interfacial debonding failure.

[0031] The nanofiller is selected from at least one of nano-silica, carbon nanotubes, and nano-graphene, with a particle size controlled between 10 and 50 nm. It fills the micropores of the resin through the nano-size effect, constructing a nano-reinforcing network and improving the matrix modulus, hardness, and fatigue resistance. The toughening agent is a compound of hydroxyl-terminated polybutadiene and dibutyl phthalate in a mass ratio of (1–3):1, with a total addition of 0.1–2 parts by weight. The hydroxyl-terminated polybutadiene provides a flexible rubber phase to achieve the core toughening effect, while dibutyl phthalate reduces resin viscosity and improves the dispersibility of the nanofiller, synergistically enhancing the system's toughness and processing performance. The curing agent is selected from at least one of m-phenylenediamine, phthalic anhydride, and methyltetrahydrophthalic anhydride. It precisely controls the crosslinking reaction rate and crosslinking density of the epoxy resin to ensure full curing and the formation of a stable and dense three-dimensional crosslinked network structure.

[0032] High-vacuum degassing completely eliminates dissolved gases and microbubbles within the resin system, preventing porosity defects in composite materials at the source and improving the density and interlayer properties of the manufactured parts. Hot-melt prepreg technology eliminates solvent involvement, avoiding solvent evaporation and residue defects. Precisely controllable resin content balances prepreg flowability and fiber volume content, ensuring sufficient resin impregnation of the fibers during compression molding. This ultimately produces high-strength, high-toughness, and high-density composite materials, effectively solving the technical challenges of poor interfacial stability, susceptibility to fatigue damage, and insufficient molding density inherent in conventional prepregs.

[0033] Based on the optimization of the resin matrix formulation, conventional carbon fiber surface modification techniques have inherent defects. Traditional plasma modification can only introduce temporary active functional groups, which gradually decay with prolonged storage and use. Furthermore, the micro-rough structure exhibits weak anti-slip ability, low interlaminar shear strength in the composite material, and the anchoring structure is prone to wear and failure under long-term alternating loads, failing to maintain the high mechanical properties of the fabric over the long term. To address this, this invention constructs a multi-step composite modification process to prepare composite carbon fiber fabric with high interfacial stability. The specific steps are as follows: The first step is to prepare the tetrabutyl titanate hydrolysis precursor solution: Concentrated hydrochloric acid (12–15 mol / L), deionized water, tetrabutyl titanate, and anhydrous ethanol are mixed and stirred in a volume ratio of (50–100):(40–80):(1.5–2.5):(10–20). First, the concentrated hydrochloric acid and deionized water are mixed, then tetrabutyl titanate is slowly added dropwise and mixed thoroughly. Finally, anhydrous ethanol is added and stirred continuously to prepare a uniform and stable precursor solution. Concentrated hydrochloric acid acts as an acidic catalyst, precisely controlling the hydrolysis and polycondensation reaction rate of tetrabutyl titanate and preventing rapid aggregation of reactants. Anhydrous ethanol acts as a dispersion and retardation medium, stabilizing the precursor solution system and ensuring uniform deposition of the subsequent coating.

[0034] The second step is inorganic coating composite modification: the pretreated modified carbon fiber cloth is immersed in the precursor solution at a solid-liquid ratio of 1g:(30-50)mL, and thoroughly stirred to ensure complete penetration of the precursor solution into the fiber bundle. The impregnated carbon fiber cloth is then transferred to a reactor, and the temperature is increased to 110-130℃ at a rate of 1-2℃ / min, and the reaction is maintained at this temperature for 10-15 hours. After the reaction, the cloth is naturally cooled to room temperature, centrifuged, washed alternately with deionized water and ethanol, and dried to finally obtain a composite carbon fiber cloth with a stable inorganic titanium oxide coating on its surface. This inorganic coating can construct a durable and stable micro-rough structure and active sites on the carbon fiber surface, which will not decay or wear down over time, completely solving the performance shortcomings of traditional plasma modification.

[0035] This invention further optimizes the carbon fiber pretreatment process, providing a high-quality substrate for inorganic coating growth. The specific pretreatment steps are as follows: T300, T700, and T800 carbon fiber plain or twill fabrics with an areal density of 160–220 g / m² are selected. The carbon fibers are placed in an acetone solution at 70–80°C and heated under reflux for 20–30 hours to thoroughly remove impurities such as sizing agents and oil stains from the fiber surface. After repeated washing with deionized water, the fabric is dried to obtain a clean, desized carbon fiber cloth. Using 0.3–1.0 wt% dilute sulfuric acid as the electrolyte, the desized carbon fiber cloth is used as the anode, and graphite as the cathode. The electrode spacing is set to 2–4 cm, and the anode current density to 0.3–0.6 mA / cm². Electrochemical oxidation etching is performed for 10–20 seconds to construct a uniform micro-rough morphology and oxygen-containing active functional groups on the fiber surface. After washing with deionized water until neutral, the fabric is dried to obtain the pretreated carbon fiber cloth.

[0036] After pretreatment, a bismuth tungstate functional layer is grown in situ via hydrothermal reaction to further enhance the fiber interface properties: A precursor solution is prepared by dissolving 1.0–1.5 g of bismuth nitrate in 30–50 mL of deionized water to obtain solution A, and 0.4–1.0 g of sodium tungstate in 45–100 mL of deionized water to obtain solution B. Solution B is slowly poured into solution A at a volume ratio of 1:(1.5–1.8), and the mixture is stirred at a constant temperature of 40–43°C to obtain a homogeneous reaction solution. The pretreated carbon fiber cloth is immersed in the reaction solution at a solid-liquid ratio of 1 g:(50–80) mL to fully impregnate it, then transferred to a reactor and subjected to hydrothermal reaction at 160–170°C for 15–20 h. After natural cooling, the mixture is centrifuged, washed, and dried to obtain high-performance modified carbon fiber cloth.

[0037] This multi-step composite modification process has a clear mechanism and stable effects: acetone reflux treatment thoroughly removes impurities from the fiber surface, ensuring a clean and uniform modified substrate; electrochemical etching precisely controls the microstructure of the fiber surface, introducing a large number of active anchoring points, laying the foundation for the adhesion of inorganic particles; hydrothermally grown bismuth tungstate crystals and subsequent titanium oxide coating form a composite anchoring structure, combining chemical bonding and physical interlocking, significantly improving the interfacial bonding strength between carbon fiber and the resin matrix. This composite interface structure exhibits extremely strong chemical stability, without the problems of active group decay or structural wear failure. It can significantly improve the interlaminar anti-slip and shear resistance of the composite material, and can still maintain excellent mechanical properties under long-term alternating loads, solving the core problems of easy fatigue and interlaminar failure of UAV fuselages from the raw material level.

[0038] This invention systematically solves various defects in traditional UAV fuselage compression molding processes through a comprehensive technical solution involving process optimization, mold structure improvement, resin formulation modification, and carbon fiber interface reinforcement. The molding process avoids fiber damage, improves production efficiency, enhances end-forming quality, and protects the integrity of the finished product. The modified resin matrix combines high strength, high toughness, and high density. The composite modified carbon fiber fabric constructs a durable and stable fiber-resin interface, comprehensively improving the mechanical properties, structural stability, and service life of the composite material UAV fuselage, thus meeting the long-term usage requirements of UAVs under complex operating conditions.

[0039] Example 1

[0040] This embodiment provides a method for molding a composite material drone fuselage, specifically including the following steps: S1: Flat Plate and Mold Cleaning: Clean the operating flat plate that is compatible with the mold size. All laying and placement processes are completed on the flat plate. Use acetone to thoroughly clean the upper mold, lower mold, each core mold segment, and end blocks to remove oil, dust, and residual impurities. After cleaning, assemble all core mold segments into a complete core mold assembly. Then, evenly apply release agent to all cavity surfaces and segment mating surfaces of the mold to ensure smooth demolding.

[0041] S2: Parts Laying: A molding method using a core mold and male mold for laying, followed by mold closing and pre-pressing, is adopted, abandoning the traditional lower mold laying and flanging process. Using the assembled core mold as a reference, the main prepreg sheets and local reinforcing sheets are laid sequentially according to the design drawings. First, one side of the core mold is laid; after flipping the core mold, the entire outer perimeter is fully covered. After laying, the parts are vacuum-packed and pre-compressed to remove interlayer air. The reinforcing sheets are precisely laid to conform to the core mold's outline. In areas with gradual thickness changes in the machine body, the number of layers is adjusted using a linear increasing and decreasing method. During the laying process, backing paper is pre-reserved in the flanging step, and each layer is inspected after flanging to prevent backing paper inclusion defects.

[0042] S3: Mold Closure and Pre-compression: After all the prepreg sheets are laid, precisely align the prepreg-covered core mold assembly with the upper and lower molds and close the mold. Place a fluoroplastic release liner at the contact point between the prepreg and the mold. Assemble the pressure block and the flat, boss-free end blocks. Apply grease that can withstand the curing temperature evenly to the surface of the locking bolts and tighten them. Hold the pressure for 15 minutes. After pre-compression, observe the adhesion of the release liner and the compaction of the part. Add material sheets to any areas with gaps or insufficient compaction. Thoroughly clean any excess material sheets from the mold closing gap to eliminate overflow and clamping defects.

[0043] S4: Curing and Molding: After the pre-pressing process, loosen the mold locking bolts, open the mold and remove the internal release film. Then, precisely close the mold again, lock the pressure block and stop block, and send the entire mold into the press for curing. Curing process parameters: Heat to 65℃ at a uniform rate of 1℃ / min, hold for 30 minutes to allow the resin to fully melt and flow; then slowly increase the pressure to 3.0MPa, monitoring the mold closing gap in real time to ensure it is <0.1mm; if the gap exceeds the standard, gradually increase the pressure in increments of 0.5MPa until the gap meets the standard. After the gap is acceptable, continue heating to 120℃ at a rate of 1℃ / min, and maintain constant temperature and pressure for 120 minutes. After curing, cool to below 60℃ at a uniform rate of 1℃ / min, release the pressure, open the mold, and remove the entire mold set.

[0044] S5: Demolding Operation: A step-by-step, orderly demolding method is adopted. First, the end blocks and upper pressure blocks of the mold are removed, and the upper mold is taken off. Then, the four connecting plates and auxiliary blocks of the mold are removed in sequence. Finally, each block of the core mold is disassembled one by one to release the internal support constraints of the part and complete the demolding of the machine body parts. During the demolding process, only copper spatulas or composite material wedges are used to gently pry the mold in the remaining area of ​​the part, strictly avoiding damage to the part surface and the mold cavity.

[0045] S6: Edge Cutting and Drilling: S6-1 Contour Cutting: Use a pneumatic cutting machine to roughly cut along the theoretical contour line of the part, leaving a uniform machining allowance of 0.5mm; after rough cutting, use a grinder to remove edge burrs, trim the allowance to the theoretical cutting line, and after initial inspection, use wet sandpaper to polish and eliminate the cutting line protrusion to ensure a smooth contour. S6-2 Drilling: Select a drill bit of appropriate specification to complete the hole machining of the part, and finally use a reamer to ream and trim the hole, remove burrs from the hole opening, and ensure the accuracy of the hole position.

[0046] S7: Finished Product Inspection: Conduct comprehensive inspection of the external dimensions, contour accuracy, and surface quality of the formed drone fuselage, check for defects such as bubbles, cracks, scratches, and missing materials, record all inspection data completely, archive them into the inspection procedures, and achieve quality traceability.

[0047] The prepreg preparation method used in this embodiment is as follows: 100 parts by weight of bisphenol A epoxy resin are heated to 120°C, 3 parts of polyethersulfone are added, and the mixture is stirred at 600 r / min for 60 min until the system is uniform and transparent; the temperature is naturally cooled to 60°C, and 7 parts of carboxyl-terminated liquid nitrile rubber and 1 part of vinyltris(β-methoxyethoxy)silane are added sequentially and stirred for 10 min to mix homogeneously; 3 parts of nano-silica filler with a particle size of 10-50 nm are added and stirred at 400 r / min for 30 min, and then 0.1 parts of compound toughening agent (the mass ratio of hydroxyl-terminated polybutadiene to dibutyl phthalate is (1-3):1, preferably 2:1 in this embodiment) are added and stirred for 5 min; finally, 15 parts of m-phenylenediamine curing agent are added and stirred continuously for 15 min to obtain a uniformly mixed resin matrix. The resin matrix was degassed in a vacuum environment of -0.09MPa for 20 minutes, and the composite carbon fiber cloth was impregnated using a hot melt prepreg process. The resin content was precisely controlled at 35% to obtain the finished prepreg.

[0048] The method for preparing the composite carbon fiber cloth used in this embodiment is as follows: 1) Pretreatment and Desizing Etching: T700 plain weave carbon fiber cloth (area density 180 g / m²) was placed in acetone solution and heated under reflux at 70°C for 20 h to thoroughly remove the sizing agent and impurities on the surface. After repeated washing with deionized water, it was vacuum dried at 60°C to constant weight to obtain desizing carbon fiber cloth. Using 0.3 wt% dilute sulfuric acid as electrolyte, the desizing carbon fiber cloth was used as the anode and graphite as the cathode, with an electrode spacing of 2 cm. An anodic current density of 0.3 mA / cm² was applied for electrochemical oxidation etching for 10 s. After removal, it was washed with deionized water until the filtrate was neutral and vacuum dried at 60°C to constant weight to obtain pretreated carbon fiber cloth.

[0049] 2) Preparation of bismuth tungstate reaction solution: Dissolve 1.1g of bismuth nitrate in 30mL of deionized water and stir to dissolve to obtain solution A; dissolve 0.4g of sodium tungstate in 45mL of deionized water and stir to dissolve to obtain solution B; slowly pour solution B into solution A at a volume ratio of 1:1.5, and stir at 40℃ and 500r / min for 50min to obtain a homogeneous reaction solution.

[0050] 3) Hydrothermal growth modified layer: The pretreated carbon fiber cloth was immersed in the reaction solution at a solid-liquid ratio of 1g:50mL and ultrasonically dispersed at 200W for 10min to fully wet it; it was transferred to the reaction vessel and hydrothermally reacted at 160℃ for 15h; it was naturally cooled to room temperature, centrifuged, washed alternately with deionized water and anhydrous ethanol, and dried to obtain bismuth tungstate modified carbon fiber cloth.

[0051] 4) Preparation of titanium oxide precursor solution: Mix 50 mL of 12 mol / L concentrated hydrochloric acid with 40 mL of deionized water and stir. Slowly add 1.5 mL of tetrabutyl titanate, mix well, and then add 10 mL of anhydrous ethanol. Stir thoroughly to obtain a stable precursor solution.

[0052] 5) Secondary inorganic coating modification: The modified carbon fiber cloth was immersed in the precursor solution at a solid-liquid ratio of 1g:30mL and stirred at 200r / min for 10min to fully wet it; it was then transferred to a reaction vessel and heated to 110℃ at 1℃ / min and kept at that temperature for 10h; after natural cooling, it was centrifuged and separated, and then washed and dried alternately to finally obtain a composite carbon fiber cloth with high interfacial stability.

[0053] Example 2

[0054] This embodiment provides a method for molding a composite material drone fuselage, specifically including the following steps: S1: Flat Plate and Mold Cleaning: Clean the operating flat plate that is compatible with the mold size. All laying and placement processes are completed on the flat plate. Use acetone to thoroughly clean the upper mold, lower mold, each core mold segment, and end blocks to remove oil, dust, and residual impurities. After cleaning, assemble all core mold segments into a complete core mold assembly. Then, evenly apply release agent to all cavity surfaces and segment mating surfaces of the mold to ensure smooth demolding.

[0055] S2: Parts Laying: A molding method using a core mold and male mold for laying, followed by mold closing and pre-pressing, is adopted, abandoning the traditional lower mold laying and flanging process. Using the assembled core mold as a reference, the main prepreg sheets and local reinforcing sheets are laid sequentially according to the design drawings. First, one side of the core mold is laid; after flipping the core mold, the entire outer perimeter is fully covered. After laying, the parts are vacuum-packed and pre-compressed to remove interlayer air. The reinforcing sheets are precisely laid to conform to the core mold's outline. In areas with gradual thickness changes in the machine body, the number of layers is adjusted using a linear increasing and decreasing method. During the laying process, backing paper is pre-reserved in the flanging step, and each layer is inspected after flanging to prevent backing paper inclusion defects.

[0056] S3: Mold Closure and Pre-compression: After all the prepreg sheets are laid, precisely align the prepreg-covered core mold assembly with the upper and lower molds and close the mold. Place a fluoroplastic release film at the contact point between the prepreg and the mold. Assemble the pressure block and the flat, boss-free end blocks. Apply grease that can withstand the curing temperature evenly to the surface of the locking bolts and tighten them. Hold the pressure for 18 minutes. After pre-compression, observe the adhesion of the release film and the compaction of the part. Add material sheets to any areas with gaps or insufficient compaction. Thoroughly clean any excess material sheets from the mold closing gap to eliminate overflow and clamping defects.

[0057] S4: Curing and Molding: After the pre-pressing process, loosen the mold locking bolts, open the mold and remove the internal release film. Then, precisely close the mold again, lock the pressure block and stop block, and send the entire mold into the press for curing. Curing process parameters: Heat to 70℃ at a uniform rate of 1℃ / min, hold for 35 minutes to allow the resin to fully melt and flow; then slowly increase the pressure to 3.2MPa, monitoring the mold closing gap in real time to ensure it is <0.1mm; if the gap exceeds the standard, gradually increase the pressure in increments of 0.5MPa until the gap meets the standard. After the gap is acceptable, continue heating to 125℃ at a rate of 1℃ / min, and maintain constant temperature and pressure for curing for 150 minutes. After curing, cool to below 60℃ at a uniform rate of 1℃ / min, release the pressure, open the mold, and remove the entire mold set.

[0058] S5: Demolding Operation: A step-by-step, orderly demolding method is adopted. First, the end blocks and upper pressure blocks of the mold are removed, and the upper mold is taken off. Then, the four connecting plates and auxiliary blocks of the mold are removed in sequence. Finally, each block of the core mold is disassembled one by one to release the internal support constraints of the part and complete the demolding of the machine body parts. During the demolding process, only copper spatulas or composite material wedges are used to gently pry the mold in the remaining area of ​​the part, strictly avoiding damage to the part surface and the mold cavity.

[0059] S6: Edge Cutting and Drilling: S6-1 Contour Cutting: Use a pneumatic cutting machine to roughly cut along the theoretical contour line of the part, leaving a uniform machining allowance of 0.8mm; after rough cutting, use a grinder to remove edge burrs, trim the allowance to the theoretical cutting line, and after initial inspection, use wet sandpaper to polish and eliminate the cutting line protrusion to ensure a smooth contour. S6-2 Drilling: Select a drill bit of appropriate specification to complete the hole machining of the part, and finally use a reamer to ream and trim the hole, remove burrs from the hole opening, and ensure the accuracy of the hole position.

[0060] S7: Finished Product Inspection: Conduct comprehensive inspection of the external dimensions, contour accuracy, and surface quality of the formed drone fuselage, check for defects such as bubbles, cracks, scratches, and missing materials, record all inspection data completely, archive them into the inspection procedures, and achieve quality traceability.

[0061] The prepreg preparation method used in this embodiment is as follows: 120 parts by weight of bisphenol F epoxy resin are heated to 125°C, 4 parts of polyethersulfone are added, and the mixture is stirred at 800 r / min for 80 min until the system is uniform and transparent; the temperature is naturally cooled to 65°C, and 12 parts of carboxyl-terminated liquid nitrile rubber and 2 parts of N-(β-aminoethyl)-γ-aminopropylmethyldimethoxysilane are added sequentially and stirred for 15 min to mix homogeneously; 6 parts of carbon nanotube filler with a particle size of 10-50 nm are added and stirred at 450 r / min for 35 min, then 1 part of compound toughening agent (the mass ratio of hydroxyl-terminated polybutadiene to dibutyl phthalate is (1-3):1, preferably 2:1 in this embodiment) is added and stirred for 4 min; finally, 25 parts of methyltetrahydrophthalic anhydride curing agent are added and stirred continuously for 20 min to obtain a uniformly mixed resin matrix. The resin matrix was degassed in a vacuum environment of -0.092MPa for 25 minutes, and the composite carbon fiber cloth was impregnated using a hot melt prepreg process. The resin content was precisely controlled to 40% to obtain the finished prepreg.

[0062] The method for preparing the composite carbon fiber cloth used in this embodiment is as follows: 1) Pretreatment and Desizing Etching: T700 twill carbon fiber cloth (area density 200 g / m²) was placed in acetone solution and heated under reflux at 75°C for 25 h to thoroughly remove the sizing agent and impurities on the surface. After repeated washing with deionized water, it was vacuum dried at 65°C to constant weight to obtain desizing carbon fiber cloth. Using 0.8 wt% dilute sulfuric acid as electrolyte, the desizing carbon fiber cloth was used as the anode and graphite as the cathode, with an electrode spacing of 3 cm. An anodic current density of 0.5 mA / cm² was applied for electrochemical oxidation etching for 15 s. After removal, it was washed with deionized water until the filtrate was neutral and vacuum dried at 65°C to constant weight to obtain pretreated carbon fiber cloth.

[0063] 2) Preparation of bismuth tungstate reaction solution: Dissolve 1.3g of bismuth nitrate in 40mL of deionized water and stir to dissolve to obtain solution A; dissolve 0.7g of sodium tungstate in 75mL of deionized water and stir to dissolve to obtain solution B; slowly pour solution B into solution A at a volume ratio of 1:1.6, and stir at 42℃ and 600r / min for 60min to obtain a homogeneous reaction solution.

[0064] 3) Hydrothermal growth of modified layer: The pretreated carbon fiber cloth was immersed in the reaction solution at a solid-liquid ratio of 1g:70mL and ultrasonically dispersed at 250W for 20min to fully impregnate it; it was transferred to the reaction vessel and hydrothermally reacted at 165℃ for 18h; it was naturally cooled to room temperature, centrifuged, washed alternately with deionized water and anhydrous ethanol, and dried to obtain bismuth tungstate modified carbon fiber cloth.

[0065] 4) Preparation of titanium oxide precursor solution: Mix 80 mL of 13 mol / L concentrated hydrochloric acid with 60 mL of deionized water and stir. Slowly add 2 mL of tetrabutyl titanate, mix well, and then add 15 mL of anhydrous ethanol. Stir thoroughly to obtain a stable precursor solution.

[0066] 5) Secondary inorganic coating modification: The modified carbon fiber cloth was immersed in the precursor solution at a solid-liquid ratio of 1g:40mL and stirred at 250r / min for 20min to fully wet it; it was then transferred to a reaction vessel and heated to 120℃ at 2℃ / min and kept at that temperature for 12h; after natural cooling, it was centrifuged and separated, and then washed and dried alternately to finally obtain a composite carbon fiber cloth with high interfacial stability.

[0067] Example 3

[0068] This embodiment provides a method for molding a composite material drone fuselage, specifically including the following steps: S1: Flat Plate and Mold Cleaning: Clean the operating flat plate that is compatible with the mold size. All laying and placement processes are completed on the flat plate. Use acetone to thoroughly clean the upper mold, lower mold, each core mold segment, and end blocks to remove oil, dust, and residual impurities. After cleaning, assemble all core mold segments into a complete core mold assembly. Then, evenly apply release agent to all cavity surfaces and segment mating surfaces of the mold to ensure smooth demolding.

[0069] S2: Parts Laying: A molding method using a core mold and male mold for laying, followed by mold closing and pre-pressing, is adopted, abandoning the traditional lower mold laying and flanging process. Using the assembled core mold as a reference, the main prepreg sheets and local reinforcing sheets are laid sequentially according to the design drawings. First, one side of the core mold is laid; after flipping the core mold, the entire outer perimeter is fully covered. After laying, the parts are vacuum-packed and pre-compressed to remove interlayer air. The reinforcing sheets are precisely laid to conform to the core mold's outline. In areas with gradual thickness changes in the machine body, the number of layers is adjusted using a linear increasing and decreasing method. During the laying process, backing paper is pre-reserved in the flanging step, and each layer is inspected after flanging to prevent backing paper inclusion defects.

[0070] S3: Mold Closure and Pre-compression: After all the prepreg sheets are laid, precisely align the prepreg-covered core mold assembly with the upper and lower molds and close the mold. Place a fluoroplastic release liner at the contact point between the prepreg and the mold. Assemble the pressure block and the flat, boss-free end blocks. Apply grease that can withstand the curing temperature evenly to the surface of the locking bolts and tighten them. Hold the pressure for 20 minutes. After pre-compression, observe the adhesion of the release liner and the compaction of the part. Add material sheets to any areas with gaps or insufficient compaction. Thoroughly clean any excess material sheets from the mold closing gap to eliminate overflow and clamping defects.

[0071] S4: Curing and Molding: After the pre-pressing process, loosen the mold locking bolts, open the mold and remove the internal release film. Then, precisely close the mold again, lock the pressure block and stop block, and send the entire mold into the press for curing. Curing process parameters: Heat to 75℃ at a uniform rate of 1℃ / min, hold for 40 minutes to allow the resin to fully melt and flow; then slowly increase the pressure to 3.5MPa, monitoring the mold closing gap in real time to ensure it is <0.1mm; if the gap exceeds the standard, gradually increase the pressure in increments of 0.5MPa until the gap meets the standard. After the gap is acceptable, continue heating to 130℃ at a rate of 1℃ / min, and maintain constant temperature and pressure for curing for 180 minutes. After curing, cool to below 60℃ at a uniform rate of 1℃ / min, release the pressure, open the mold, and remove the entire mold set.

[0072] S5: Demolding Operation: A step-by-step, orderly demolding method is adopted. First, the end blocks and upper pressure blocks of the mold are removed, and the upper mold is taken off. Then, the four connecting plates and auxiliary blocks of the mold are removed in sequence. Finally, each block of the core mold is disassembled one by one to release the internal support constraints of the part and complete the demolding of the machine body parts. During the demolding process, only copper spatulas or composite material wedges are used to gently pry the mold in the remaining area of ​​the part, strictly avoiding damage to the part surface and the mold cavity.

[0073] S6: Edge Cutting and Drilling: S6-1 Contour Cutting: Use a pneumatic cutting machine to roughly cut along the theoretical contour line of the part, leaving a uniform machining allowance of 1mm; after rough cutting, use a grinder to remove edge burrs, trim the allowance to the theoretical cutting line, and after initial inspection, use wet sandpaper to polish and eliminate the cutting line protrusion to ensure a smooth contour. S6-2 Drilling: Select a drill bit of appropriate specification to complete the hole machining of the part, and finally use a reamer to ream and trim the hole, remove burrs from the hole opening, and ensure the accuracy of the hole position.

[0074] S7: Finished Product Inspection: Conduct comprehensive inspection of the external dimensions, contour accuracy, and surface quality of the formed drone fuselage, check for defects such as bubbles, cracks, scratches, and missing materials, record all inspection data completely, archive them into the inspection procedures, and achieve quality traceability.

[0075] The prepreg preparation method used in this embodiment is as follows: 150 parts by weight of bisphenol A epoxy resin are heated to 130°C, 5 parts of polyethersulfone are added, and the mixture is stirred at 1000 r / min for 100 min until the system is uniform and transparent; the temperature is naturally cooled to 70°C, and 15 parts of carboxyl-terminated liquid nitrile rubber and 3 parts of γ-aminopropylmethyldiethoxysilane are added sequentially and stirred for 20 min to mix homogeneously; 10 parts of nano-graphene filler with a particle size of 10-50 nm are added and stirred at 500 r / min for 40 min, then 2 parts of compound toughening agent (the mass ratio of hydroxyl-terminated polybutadiene to dibutyl phthalate is (1-3):1, preferably 2:1 in this embodiment) are added and stirred for 10 min; finally, 30 parts of phthalic anhydride curing agent are added and stirred continuously for 25 min to obtain a uniformly mixed resin matrix. The resin matrix was degassed in a vacuum environment of -0.095MPa for 30 minutes, and the composite carbon fiber cloth was impregnated using a hot melt prepreg process. The resin content was precisely controlled to 45% to obtain the finished prepreg.

[0076] The method for preparing the composite carbon fiber cloth used in this embodiment is as follows: 1) Pretreatment and Desizing Etching: T800 plain weave carbon fiber cloth (area density 220 g / m²) was placed in acetone solution and heated to reflux at 80°C for 30 h to thoroughly remove the sizing agent and impurities on the surface. After repeated washing with deionized water, it was vacuum dried at 70°C to constant weight to obtain desizing carbon fiber cloth. Using 1.0 wt% dilute sulfuric acid as electrolyte, the desizing carbon fiber cloth was used as the anode and graphite as the cathode, with an electrode spacing of 4 cm. An anodic current density of 0.6 mA / cm² was applied for electrochemical oxidation etching for 20 s. After removal, it was washed with deionized water until the filtrate was neutral and vacuum dried at 70°C to constant weight to obtain pretreated carbon fiber cloth.

[0077] 2) Preparation of bismuth tungstate reaction solution: Dissolve 1.5g of bismuth nitrate in 50mL of deionized water and stir to dissolve to obtain solution A; dissolve 1.0g of sodium tungstate in 100mL of deionized water and stir to dissolve to obtain solution B; slowly pour solution B into solution A at a volume ratio of 1:1.8, and stir at a constant temperature of 43℃ and 700r / min for 70min to obtain a homogeneous reaction solution.

[0078] 3) Hydrothermal growth modified layer: The pretreated carbon fiber cloth was immersed in the reaction solution at a solid-liquid ratio of 1g:80mL and ultrasonically dispersed at 300W for 30min to fully wet it; it was transferred to the reaction vessel and hydrothermally reacted at 170℃ for 20h; it was naturally cooled to room temperature, centrifuged, washed alternately with deionized water and anhydrous ethanol, and dried to obtain bismuth tungstate modified carbon fiber cloth.

[0079] 4) Preparation of titanium oxide precursor solution: Mix 100 mL of 15 mol / L concentrated hydrochloric acid with 80 mL of deionized water and stir. Slowly add 2.5 mL of tetrabutyl titanate, mix well, and then add 20 mL of anhydrous ethanol. Stir thoroughly to obtain a stable precursor solution.

[0080] 5) Secondary inorganic coating modification: The modified carbon fiber cloth was immersed in the precursor solution at a solid-liquid ratio of 1g:50mL and stirred at 300r / min for 30min to fully wet it; it was then transferred to a reaction vessel and heated to 130℃ at 2℃ / min and kept at that temperature for 15h; after natural cooling, it was centrifuged and separated, and then washed and dried alternately to finally obtain a composite carbon fiber cloth with high interfacial stability.

[0081] Comparative Example 1 The process steps, process parameters, resin matrix formulation, and molding conditions of this comparative example are completely consistent with those of Example 1. The only difference is that in the prepreg preparation process, unmodified ordinary T700 grade carbon fiber plain weave fabric (area density 160g / m²) is directly used to replace the high interface stability composite carbon fiber fabric in Example 1. The other raw material ratios and processing procedures remain unchanged.

[0082] Comparative Example 2 The process steps, process parameters, resin matrix formulation, and molding conditions of this comparative example are completely consistent with those of Example 1. The only difference is that in the preparation of the composite carbon fiber cloth, the carbon fiber pretreatment slurry removal and etching process in step 1) is omitted; at the same time, the pretreated carbon fiber cloth used in the hydrothermal reaction in step 3) is directly replaced with the original T700 carbon fiber cloth, while the subsequent bismuth tungstate growth and titanium oxide precursor liquid coating modification processes are retained, and the remaining process parameters remain unchanged.

[0083] Comparative Example 3 The process steps, process parameters, resin matrix formulation, and molding conditions of this comparative example are completely consistent with those of Example 1. The only difference is that in the preparation of the composite carbon fiber cloth, step 2) preparation of the bismuth tungstate reaction solution and step 3) hydrothermal growth of the modified layer are omitted; at the same time, the modified carbon fiber cloth used in step 5) secondary inorganic coating modification is directly replaced with the pretreated carbon fiber cloth prepared in step 1) of Example 1, and only the titanium oxide coating modification step is retained, while the other process parameters remain unchanged.

[0084] Comparative Example 4 The process steps, process parameters, resin matrix formulation, and molding conditions of this comparative example are completely consistent with those of Example 1. The only difference is that in the preparation of composite carbon fiber cloth, step 4) preparation of titanium oxide precursor liquid and step 5) secondary inorganic coating modification are omitted. Only the carbon fiber pretreatment and bismuth tungstate hydrothermal modification steps are retained, and the remaining process parameters remain unchanged.

[0085] Comparative Example 5 The process steps, process parameters, resin matrix formulation, and molding conditions of this comparative example are completely consistent with those of Example 1. The only difference is that in the preparation of the composite carbon fiber cloth, all modification processes from step 2) to step 5) are omitted. Only the original carbon fiber cloth that has been simply cleaned with acetone and dried is used as the reinforcing substrate. There is no electrochemical etching, inorganic crystal growth, or oxide coating modification treatment. The remaining process parameters remain unchanged.

[0086] Performance testing experiment To verify the effect of the modified carbon fiber composite system and molding process of this invention on improving the mechanical properties of UAV fuselage composite materials, composite UAV fuselage samples were prepared using the preparation methods of Examples 1-3 and Comparative Examples 1-5, respectively. Mechanical property comparison tests were conducted under unified environmental conditions and testing standards. The specific testing standards are as follows: 1. Tensile property test: According to ASTM D3039 / D3039M-17 "Standard Test Method for Tensile Properties of Polymer-Based Composites", the tensile strength and tensile modulus of each group of specimens are tested to characterize the axial load-bearing capacity and tensile deformation resistance of the composite material.

[0087] 2. Bending performance test: According to ASTM D7264 / D7264M-21 "Standard Test Method for Bending Performance of Polymer-based Composite Materials", the bending strength and bending modulus of each group of samples are tested by the three-point bending method to characterize the bending and flexural deformation resistance of the composite material and match the bending load-bearing conditions of the fuselage.

[0088] 3. Interlaminar shear performance test: According to ASTM D2344 / D2344M-22 "Standard Test Method for Strength of Polymer-Based Composite Short Beams", the interlaminar shear strength of each group of specimens was tested to characterize the interfacial bonding stability between the fiber and the resin matrix and the interlaminar peel resistance and shear failure resistance of the composite material.

[0089] All samples were tested using the same dimensions, grinding, and pretreatment methods. Five parallel tests were conducted on each sample group, and the average value was taken as the final test result to avoid test errors and ensure data accuracy and comparability.

[0090] Five parallel experiments were designed, and the average value of the experimental results was taken. The results are shown in Table 1.

[0091] Table 1

[0092] As shown in Table 1, the composite material drone fuselage obtained by the present invention can significantly improve the tensile strength, flexural strength, and interlaminar shear strength of the resin by introducing composite carbon fiber cloth into the resin, thereby enhancing the performance of the composite material and effectively improving the performance of the drone.

[0093] The preferred embodiments of the present invention disclosed above are merely illustrative of the invention. These preferred embodiments do not exhaustively describe all details, nor do they limit the invention to the specific implementations described. Clearly, many modifications and variations can be made based on the content of this specification. This specification selects and specifically describes these embodiments to better explain the principles and practical applications of the invention, thereby enabling those skilled in the art to better understand and utilize the invention. The invention is limited only by the claims and their full scope and equivalents.

Claims

1. A method for molding a composite material unmanned aerial vehicle (UAV) fuselage, characterized in that, Includes the following steps: S1: Cleaning of flat plates and molds Clean the tiling work surface, and use acetone to clean each section of the upper mold, lower mold, core mold, and end blocks. First, apply release agent to each individual core mold section, upper mold, and lower mold cavity, and then assemble the core mold assembly to ensure that the mold is clean and facilitates subsequent demolding operations. S2: Parts Placement Using the assembled core mold as a reference, the prepreg is laid first. Then, the core mold is flipped over to complete the overall outer perimeter laying. After laying, the core mold is vacuum-packed and pre-pressed. Reinforcing sheets are laid according to the outline of the core mold. In the area where the thickness of the machine body gradually changes, the prepreg layers are laid linearly. S3: Mold closing pre-compression The core mold assembly covered with prepreg is closed with the upper and lower molds. A fluoroplastic release film is inserted on the outside of the prepreg. The pressure block and end stop are assembled. The locking bolts, which are coated with grease that can withstand the curing temperature, are tightened. The whole assembly is pre-pressed for 15 to 20 minutes. The end stop adopts a flat and fitted structure with the round hole boss removed to ensure uniform force when the mold is closed and locked. S4: Curing and molding Remove the release liner to allow direct pressing and bonding between the mold cavity and the prepreg. Lock the pressure block and stop block again to complete the secondary mold closing. Place the entire mold into the press for molding and curing. Heat to 65–75°C at a rate of ≤2°C / min, hold for 30–40 minutes, and then apply pressure to 3.0–3.5 MPa. Check the mold closing gap and control it to <0.1 mm. If the gap is >0.1 mm, apply pressure gradually at 0.5 MPa until the gap meets the standard. Once the gap meets the requirements, continue heating to 120–130°C at ≤2°C / min, hold for 120–180 minutes, and then cool to below 60°C at a uniform rate of ≤2°C / min. Remove the mold from the press. S5: Demolding operation Disassemble the stop block, pressure block, upper mold, multiple connecting plates and core mold sections in sequence, and use the core mold section matching demolding structure to successfully demold the product. S6: Edge drilling Use a pneumatic cutting machine to roughly cut the edge along the theoretical outline of the part, leaving a machining allowance of 0.5 to 1 mm. Grind to remove burrs and excess material to the cutting line position. S7: Finished Product Inspection The external dimensions and surface appearance quality of the molded parts are inspected, and the inspection results are fully recorded.

2. The molding method for a composite material unmanned aerial vehicle fuselage according to claim 1, characterized in that, The preparation method of the prepreg described in S2 is as follows: By weight, 100-150 parts of epoxy resin are heated to 120-130°C, and 3-5 parts of polyethersulfone are added and stirred thoroughly. After cooling to 60-70°C, 7-15 parts of carboxyl-terminated liquid nitrile rubber and 1-3 parts of silane coupling agent are added and stirred evenly. Then, 3-10 parts of nanofiller are added, followed by 0.1-2 parts of toughening agent and stirring. Finally, 15-30 parts of curing agent are added to obtain a mixed resin matrix. The mixed resin matrix is ​​degassed under vacuum conditions of -0.09 to -0.095 MPa for 20-30 minutes. After degassed, the resin matrix is ​​impregnated with composite carbon fiber cloth through a hot-melt prepreg process, controlling the final resin content to be 35-45%, to obtain the prepreg. The reinforcing sheet is prepared by cutting the above prepreg to the required size.

3. The molding method for a composite material unmanned aerial vehicle fuselage according to claim 2, characterized in that, The epoxy resin is selected from at least one of bisphenol A type epoxy resin and bisphenol F type epoxy resin; The curing agent is selected from at least one of m-phenylenediamine, phthalic anhydride, and methyltetrahydrophthalic anhydride; The toughening agent is a compound of hydroxyl-terminated polybutadiene and dibutyl phthalate in a mass ratio of (1-3):1, and the total amount of the compounded toughening agent added to the resin matrix is ​​0.1-2 parts by weight. The nanofiller is selected from at least one of nano-silica, carbon nanotubes, and nano-graphene, and the average particle size of the nanofiller is 10-50 nm. The silane coupling agent is selected from at least one of vinyltris(β-methoxyethoxy)silane, N-(β-aminoethyl)-γ-aminopropylmethyldimethoxysilane, and γ-aminopropylmethyldiethoxysilane.

4. The molding method for a composite material unmanned aerial vehicle fuselage according to claim 2, characterized in that, The preparation steps of the composite carbon fiber cloth are as follows: 1) Mix concentrated hydrochloric acid with deionized water and stir, add tetrabutyl titanate dropwise and mix well, then add anhydrous ethanol and stir thoroughly to prepare a precursor solution. 2) The modified carbon fiber cloth was immersed in the precursor solution, stirred and soaked, and then transferred to the reactor and kept at a temperature for 10-15 hours. After the reaction was completed, it was naturally cooled to room temperature, and then centrifuged, washed and dried to obtain the composite carbon fiber cloth.

5. The molding method for a composite material unmanned aerial vehicle fuselage according to claim 4, characterized in that, In step 1), the volume ratio of concentrated hydrochloric acid, deionized water, tetrabutyl titanate, and anhydrous ethanol is (50-100):(40-80):(1.5-2.5):(10-20); the concentration of the concentrated hydrochloric acid is 12-15 mol / L.

6. The molding method for a composite material unmanned aerial vehicle fuselage according to claim 4, characterized in that, In step 2), the solid-liquid ratio of the modified carbon fiber cloth to the precursor liquid is 1g:(30~50)mL; the temperature of the reactor is 110~130℃, and the heating rate is controlled at 1~2℃ / min.

7. The molding method for a composite material unmanned aerial vehicle fuselage according to claim 4, characterized in that, The preparation steps of the modified carbon fiber cloth are as follows: 1) Carbon fibers are placed in an acetone solution and heated under reflux at 70-80°C for 20-30 hours. After treatment, the carbon fibers are repeatedly washed with deionized water to remove residual acetone and dried to obtain desized carbon fiber cloth. Electrochemical oxidation etching is performed using dilute sulfuric acid as electrolyte, the desized carbon fiber cloth as anode, and graphite as cathode. The electrode spacing is 2-4 cm, the anode current density is 0.3-0.6 mA / cm², and the treatment time is 10-20 s. After removal, the carbon fibers are washed with deionized water until the filtrate is neutral and dried to obtain pretreated carbon fiber cloth. 2) Dissolve bismuth nitrate in deionized water and stir to prepare solution A, and dissolve sodium tungstate in deionized water and stir to prepare solution B; pour solution B into solution A according to the volume ratio of solution A: solution B 1: (1.5~1.8), and stir and mix at a constant temperature of 40~43℃ to obtain the reaction solution; 3) After the pretreated carbon fiber cloth is fully immersed in the reaction solution, it is transferred to the reaction vessel and hydrothermally reacted at 160-170℃ for 15-20 hours. After the reaction is completed, it is naturally cooled to room temperature, and after centrifugation, washing and drying, the modified carbon fiber cloth is obtained.

8. The molding method for a composite material unmanned aerial vehicle fuselage according to claim 7, characterized in that, The carbon fiber cloth mentioned in step 1) is at least one of T300, T700, T800 carbon fiber plain weave fabric and twill weave fabric, with an areal density of 160-220 g / m²; the acetone heating reflux temperature is 70-80°C; and the mass fraction of dilute sulfuric acid is 0.3-1.0 wt%.

9. A method for molding a composite material unmanned aerial vehicle fuselage according to claim 7, characterized in that: Step 2) The ratio of bismuth nitrate to deionized water in solution A is (1.0-1.5) g: (30-50) mL; the ratio of sodium tungstate to deionized water in solution B is (0.4-1.0) g: (45-100) mL.

10. A method for molding a composite material unmanned aerial vehicle fuselage according to claim 7, characterized in that, In step 3), the solid-liquid ratio of the pretreated carbon fiber cloth to the reaction solution is 1g:(50~80)mL; the hydrothermal reaction temperature is 160~170℃.

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