Insect exoskeleton morph-layered composite panel, sandwich structure and preparation method

Abstract

The application discloses an insect exoskeleton morphological layup composite panel, a sandwich structure and a preparation method. The sandwich structure comprises two pieces of bionic layup composite panels and a piece of aramid honeycomb core layer. The bionic layup composite panel draws lessons from the spiral stacking microstructure of the insect exoskeleton, the layup design considers the symmetry and balance requirements of the composite material, a fixed rotation angle is introduced between adjacent fiber layers, and the rotation angle is taken as a period of 180 degrees, and after a plurality of periods are laid and heat-pressed and cured, the bionic layup composite panel is obtained. The bionic layup design effectively reduces the interlaminar shear stress of the composite material, weakens the inherent anisotropy of the composite material, and improves the transverse bearing capacity and impact resistance. The bionic layup composite panel and the aramid honeycomb core layer have the advantages of high specific strength, high specific stiffness, impact energy absorption, vibration and noise reduction, electromagnetic shielding and the like, and a structure and function integrated material with more excellent comprehensive performance is obtained, and the structure and function integrated material is expected to be widely used in main load-bearing structures of various spacecrafts.

Description

Technical Field

[0001] This application relates to the technical field of composite material structures in the aerospace industry, and in particular to an insect exoskeleton-shaped layup composite material panel, sandwich structure and preparation method. Background Technology

[0002] As my country accelerates its exploration in fields such as satellites, manned spaceflight, and deep space exploration, the requirements for lightweight, high load-bearing capacity, and impact resistance in spacecraft such as satellites and probes are becoming increasingly stringent. Carbon-skinned aluminum honeycomb sandwich panels, currently the most commonly used structural panels in spacecraft, can no longer meet the demands of future lightweight design. Traditional sandwich panels consist of carbon fiber composite panels and aluminum honeycomb core layers, with the composite panel layup primarily using orthogonal stacking. When the sandwich panel is subjected to external mechanical loads, the composite panel is prone to matrix cracking and delamination, while the aluminum honeycomb core layer exhibits defects such as cell distortion and uneven cell walls, leading to a degradation of the overall structural mechanical properties, reduced strength and stiffness, and weakened impact energy absorption. To meet the evolving requirements of spacecraft for their load-bearing structures, there is an urgent need to develop lightweight, high-strength, multifunctional structural panels that also possess impact energy absorption, vibration reduction, noise reduction, and electromagnetic shielding properties. Summary of the Invention

[0003] This application provides an insect exoskeleton-shaped layup composite material panel, a sandwich structure, and a preparation method thereof. The biomimetic layup composite material panel and the aramid honeycomb core are connected by a two-component epoxy resin adhesive to obtain a high-strength and tough all-composite sandwich structure.

[0004] In a first aspect, an insect exoskeleton-shaped plywood composite panel is provided, the panel being formed by hot pressing multiple layers of unidirectional fiber prepreg, wherein if the panel is an even-numbered layer composite material, the fiber directions of each layer on the panel are sequentially arranged as follows: n ×180° / n ×180°- i / n ×180°-2 i / … / 2 i / i / 0°] s Arrangement, if the panel is an odd-layer composite material, the fiber directions of each layer on the panel are in the following order [( n ×180° / n ×180°- i / n ×180°-2 i / … / 2 i / i / 0°) s Arranged as / 0°].

[0005] In conjunction with the first aspect, in some implementations of the first aspect, 5° ≤ i ≤ 30°, and 180° is i Integer multiples of.

[0006] In conjunction with the first aspect, in some implementations of the first aspect, the fiber is at least one of the following: carbon fiber, glass fiber, basalt fiber, or natural fiber.

[0007] Secondly, a method for preparing an insect exoskeleton morphology layup composite material panel as described in any of the implementations of the first aspect above is provided, comprising:

[0008] For unidirectional fiber prepregs, with the fiber direction as 0°, the deflection is linearly increased sequentially according to the layup order. i Cutting is performed; if the panel is an even-numbered layer composite material, then 2 layers are cut. n+2 Prepreg layer ， If the panel is an odd-numbered layer composite material, then it is cut to obtain 2 n+3 One prepreg layer;

[0009] The cut prepregs are stacked layer by layer according to the layup sequence. The stacked prepregs are then hot-pressed and cooled. For even-numbered layer composites, the layup sequence is […]. n ×180° / n ×180°- i / n ×180°-2 i / … / 2 i / i / 0°] s For odd-layer composite materials, the layup sequence is [( n ×180° / n ×180°- i / n ×180°-2 i / … / 2 i / i / 0°) s / 0°).

[0010] In conjunction with the second aspect, in some implementations of the second aspect, the unidirectional fiber prepreg has a unit area mass of 50 ~ 250 g / m². 2 The thickness is 0.05 ~ 0.25 mm.

[0011] In conjunction with the second aspect, in some implementations of the second aspect, the hot pressing treatment is an autoclave molding process, with a molding pressure of 0.1 ~ 0.8 MPa, a heating rate of 1 ~ 5 ℃ / min, a molding temperature of 125 ~ 150 ℃, and a holding time of 0.5 ~ 2 h.

[0012] In conjunction with the second aspect, in some implementations of the second aspect, the cooling process involves reducing the temperature to 80°C at a rate of 2 to 4°C / min, and then allowing it to cool naturally to room temperature.

[0013] Thirdly, a high-strength and tough composite material sandwich structure with insect exoskeleton morphology is provided. The sandwich structure includes upper and lower layers of insect exoskeleton morphology plywood composite material panels as described in any of the implementations of the first to second aspects above, and a honeycomb core, wherein the honeycomb core is sandwiched between the two layers of insect exoskeleton morphology plywood composite material panels.

[0014] In conjunction with the third aspect, in some implementations of the third aspect, the honeycomb core is an aramid honeycomb, and the core height... T The nominal pore size is 5 ~ 25 mm, and the side length of the pore is... a The core thickness is 3~10 mm. N Its thickness is 0.1 ~ 0.5 mm, and its density is 30 ~ 160 kg / m³. 3 .

[0015] Fourthly, a method for preparing a high-strength and tough composite material sandwich structure with an insect exoskeleton morphology as described in any of the implementations of the third aspect above is provided, comprising:

[0016] Apply a uniform two-component epoxy resin adhesive to one side of the upper and lower layers of insect exoskeleton morphology composite material panel, and then attach it to the upper and lower sides of the honeycomb core of equal area.

[0017] The well-fitted insect exoskeleton-shaped high-strength and tough composite material sandwich structure is clamped with a clamp and cured at room temperature for 8-10 hours. After removing the clamp, an epoxy resin-bonded insect exoskeleton-shaped high-strength and tough composite material sandwich structure is obtained.

[0018] Compared with the prior art, the solution provided in this application has at least the following beneficial technical effects:

[0019] 1) Lightweight: Using composite sandwich structures as the main load-bearing structure can reduce the overall weight of the spacecraft and improve its fuel efficiency and payload capacity; 2) High stiffness and strength: Composite sandwich structures have excellent stiffness and strength, which can withstand the mechanical loads faced by the spacecraft during operation; 3) Fatigue resistance: Composite sandwich structures have excellent fatigue resistance, which can extend the service life of the spacecraft; 4) Impact resistance: Composite sandwich structures can absorb and disperse the impact energy of space debris, improving the protective performance of the spacecraft.

[0020] High-strength and high-toughness composite sandwich structures with insect exoskeleton morphology have significant application prospects in the aerospace field, improving the safety, reliability, and performance of spacecraft. 1) The orientation variation of composite fiber optimizes the overall strength and stiffness of the material; 2) Under the same load conditions, biomimetic fiber-reinforced composites exhibit lower interlaminar stress and are less prone to delamination failure, thus improving their interfacial properties; 3) The biomimetic helical structure allows for helical crack propagation. The change in the local fracture mode at the crack tip leads to a decrease in the local strain energy release rate, thereby increasing the external energy release rate required for crack propagation and enhancing its toughness.

[0021] The innovations of this application include: 1) The composite sandwich panel is designed with a microstructure of layered fiber layers stacked in a spiral pattern resembling an insect exoskeleton, which reduces the interlaminar shear stress of the composite material, weakens the inherent anisotropy of the composite material, and improves the mechanical properties of the sandwich structure against out-of-plane loads; 2) The design of the spiral stacked microstructure of the insect exoskeleton is combined with symmetry and balance design, which improves the molding quality of the composite material. The rotary cutting of the prepreg avoids edge warping and secondary processing of the molded panel, and simplifies the manufacturing process; 3) The high specific strength, high specific stiffness, impact energy absorption, vibration reduction and noise reduction, and electromagnetic shielding of the biomimetic layup composite panel and the aramid honeycomb core layer are fully utilized to achieve multifunctional coupling and obtain a structural-functional integrated material with better overall performance.

[0022] In the aerospace field, weight reduction and structural integrity are paramount. The practical engineering value of this application lies in: 1) High specific strength allows for the construction of lightweight yet robust components, reducing the overall weight of spacecraft and thus improving fuel efficiency and increasing payload capacity; 2) Excellent toughness and damage resistance provide better impact and fatigue resistance, ensuring the structural integrity and safety of spacecraft; 3) The sandwich design effectively absorbs energy from high-speed collisions, such as those encountered during space debris collisions or atmospheric reentry, improving the survivability of spacecraft under extreme conditions; 4) The aramid honeycomb core layer provides thermal insulation, vibration damping, noise reduction, and electromagnetic shielding, further enhancing the performance and comfort of spacecraft. In summary, the high-strength, high-toughness biomimetic composite sandwich structure with insect exoskeleton morphology has significant application value in the aerospace field. These structures have potential in weight reduction, maintaining structural integrity, impact resistance, and energy absorption, and are expected to bring about changes in spacecraft design and performance, enabling more efficient, safe, and durable spacecraft.

[0023] In summary, the high-strength and tough composite sandwich structure of insect exoskeleton morphology and its preparation method of this application have excellent mechanical properties and capabilities, which can meet the needs of the engineering field for high-performance materials and have broad application prospects. Attached Figure Description

[0024] Figure 1 This is a schematic diagram of the high-strength and tough composite material sandwich structure of the insect exoskeleton involved in this application.

[0025] The markings in the diagram are as follows: 1- Bionic layup composite panel, 2- Aramid honeycomb core.

[0026] Figure 2 This is a schematic diagram of the composite material layup for the insect exoskeleton morphology involved in this application.

[0027] Figure 3 This is a schematic diagram of the prepreg cutting involved in this application.

[0028] Figure 4 This is a schematic diagram of the aramid honeycomb core involved in this application.

[0029] Explanation of markings in the diagram: T -Core height, a -Nominal kunz side length, N - Core thickness.

[0030] Figure 5 This is a schematic diagram of the quasi-static indentation test involved in this application.

[0031] Figure 6 The force-displacement curve results of the quasi-static indentation test of Embodiment 1 involved in this application.

[0032] Figure 7 The results show the peak load and energy absorption of Embodiment 1 as described in this application.

[0033] Figure 8 The force-displacement curve results of the quasi-static indentation test of Embodiment 2 involved in this application.

[0034] Figure 9 The results show the peak load and energy absorption of Embodiment 2 as described in this application. Detailed Implementation

[0035] The present application will now be described in further detail with reference to the accompanying drawings and specific embodiments.

[0036] This application provides a laminated composite panel with an insect exoskeleton morphology. The panel is formed by hot pressing multiple layers of unidirectional fiber prepreg. If the panel is an even-numbered composite material, the fiber directions of each layer on the panel are sequentially arranged as follows: n ×180° / n ×180°- i / n ×180°-2 i / … / 2 i / i / 0°] s Arrangement, if the panel is an odd-numbered layer composite material, the fiber directions of each layer on the panel are in the following order [( n ×180° / n ×180°- i / n ×180°-2 i / … / 2 i / i / 0°) s Arranged as / 0°].

[0037] This application also provides a method for preparing an insect exoskeleton morphology layup composite material panel, including the following steps.

[0038] 1) For unidirectional fiber prepregs, with the fiber direction as 0°, the deflection is linearly increased sequentially according to the layup order. i Cutting is performed; if the panel is an even-numbered layer composite material, then 2 layers are cut. n+2 Prepreg layer ， If the panel is an odd-numbered layer composite material, then it is cut to obtain 2 n+3 One prepreg layer;

[0039] 2) Stack the cut prepregs layer by layer according to the layup sequence, and perform hot pressing and cooling treatment on the stacked prepregs. For even-numbered layer composite materials, the layup sequence is [ n ×180° / n ×180°- i / n ×180°-2 i / … / 2 i / i / 0°] s For odd-layer composite materials, the layup sequence is [( n ×180° / n ×180°- i / n ×180°-2 i / … / 2 i / i / 0°) s / 0°).

[0040] Figure 1 This is a schematic diagram of a high-strength and tough composite material sandwich structure with an insect exoskeleton morphology, as provided in this application. The sandwich structure may include an aramid honeycomb core 2 and upper and lower biomimetic plywood composite panel 1, with the aramid honeycomb core 2 sandwiched between the upper and lower biomimetic plywood composite panel 1. The biomimetic plywood composite panel 1 is the insect exoskeleton morphology plywood composite material panel described above.

[0041] This application also provides a method for preparing a high-strength and tough composite sandwich structure with insect exoskeleton morphology, including the following steps.

[0042] 1) First, a biomimetic spiral design was used for the layup sequence of the composite panel. Considering the requirements of symmetry and balance, the spiral period was determined. n With helix angle i For even-layer composite materials, the layup sequence is [ n ×180° / n ×180°- i / n ×180°-2 i / … / 2 i / i / 0°] s ,like Figure 2 As shown; for odd-layer composite materials, the layup sequence is [( n ×180° / n ×180°- i / n ×180°-2 i / … / 2 i / i / 0°) s / 0°).

[0043] The above-mentioned biomimetic layup design method takes into account the symmetry requirement and eliminates the in-plane-out-plane and out-plane-in-plane coupling effects of composite laminates, thereby avoiding warping deformation of the resulting laminates.

[0044] The above-mentioned biomimetic layup design method takes into account the requirements of balance and eliminates the tension-shear coupling effect of composite laminates under uniaxial tension.

[0045] In some embodiments, the helical angle in the biomimetic layup design method should satisfy 5° ≤ i ≤ 30°, and 180° is i Integer multiples of.

[0046] 2) Secondly, for unidirectional fiber prepregs, with the fiber direction as 0°, the deflection is linearly increased sequentially according to the layup design. i Cut, such as Figure 3 As shown.

[0047] In some embodiments, the unidirectional fiber prepreg is an epoxy resin prepreg made of carbon fiber, glass fiber, basalt fiber, natural fiber, etc., with a unit area mass of 50~250 g / m². 2 The thickness is 0.05 ~ 0.25 mm.

[0048] 3) Next, the cut prepreg is stacked layer by layer according to the layup design, and the stacked prepreg is hot-pressed and cooled.

[0049] In some embodiments, the hot pressing treatment is an autoclave molding process, with a molding pressure of 0.1 ~ 0.8 MPa, a heating rate of 1 ~ 5 ℃ / min, a molding temperature of 125 ~ 150 ℃, and a holding time of 0.5 ~ 2 h.

[0050] In some embodiments, the cooling process involves reducing the temperature to 80°C at a rate of 2 to 4°C / min, and then allowing it to cool naturally to room temperature.

[0051] 4) Finally, cut two panels of equal size from the composite material. Apply a uniform layer of two-component epoxy resin adhesive to one side and attach them to the top and bottom sides of the core of equal area. The composite material core is as follows: Figure 4 As shown. The bonded composite honeycomb sandwich panels are clamped with a fixture and cured at room temperature for 8 to 10 hours. After removing the fixture, an epoxy resin bonded all-composite sandwich panel is obtained.

[0052] In some embodiments, the honeycomb core is an aramid honeycomb, and the core height is... T The nominal pore size is 5 ~ 25 mm, and the side length of the pore is... a The core thickness is 3~10 mm. NIts thickness is 0.1 ~ 0.5 mm, and its density is 30 ~ 160 kg / m³. 3 .

[0053] Example 1

[0054] 1) First, a biomimetic spiral design was used for the layup sequence of the composite material panel. Considering the requirements of symmetry and balance, the spiral period was determined to be 2, the spiral angle to be 10°, the number of layers to be odd, and the layup sequence to be [(360° / 350° / 340° / … / 20° / 10° / 0°)]. s / 0°], and then cut the unidirectional carbon fiber prepreg with the fiber direction as 0°, rotating it 10° each time according to the layup design;

[0055] 2) Next, the cut carbon fiber prepreg is stacked layer by layer according to the layup design. The stacked prepreg is hot-pressed and formed at a pressure of 0.1 MPa, a heating rate of 4 ℃ / min, a forming temperature of 130 ℃, and a holding time of 60 min. Then, the temperature is reduced to 90 ℃ at a rate of 4 ℃ / min and allowed to cool naturally to room temperature.

[0056] 3) Finally, the composite material is water-cut into two panels of the same size. A uniform two-component epoxy resin is applied to one side. An aramid honeycomb core is then used, with a nominal cell side length of 3.2 mm, a height of 10 mm, and a density of 48 kg / m³. 3 The panel is pasted onto the upper and lower sides of the composite core of equal area. The pasted composite honeycomb sandwich panel is clamped with a fixture and cured at room temperature for 9 hours. After removing the fixture, an epoxy resin bonded all-composite sandwich panel is obtained.

[0057] Indentation tests were performed on the prepared biomimetic plywood sandwich structure according to ASTM D6264 standard, such as... Figure 5 As shown, a hemispherical indenter with a diameter of 10 mm was used, and the pressing speed was 1 mm / min. The force-displacement curve results are as follows. Figure 6 As shown. The peak load of the traditional orthogonal layup composite sandwich structure under the same fabrication process is 1171 N, with an energy absorption of 1.054 J; while the peak load of the biomimetic layup composite sandwich structure is 1709 N, with an energy absorption of 1.965 J. Figure 7 As shown, peak load and energy absorption increased by 45.9% and 86.4% respectively, achieving a significant improvement in the overall structural strength and toughness.

[0058] Example 2

[0059] 1) First, a biomimetic spiral design was used for the layup sequence of the composite panel. Considering the requirements of symmetry and balance, the spiral period was determined to be 2, the spiral angle to be 10°, the number of layers to be even, and the layup sequence to be [360° / 350° / 340° / … / 20° / 10° / 0°]. s Then, the unidirectional carbon fiber prepreg is cut by rotating it 10° at a time according to the layup design, with the fiber direction as 0°.

[0060] 2) Next, the cut carbon fiber prepreg is stacked layer by layer according to the layup design. The stacked prepreg is hot-pressed and formed at a pressure of 0.2 MPa, a heating rate of 5 ℃ / min, a forming temperature of 140 ℃, and a holding time of 40 min. Then, the temperature is reduced to 80 ℃ at a rate of 3 ℃ / min and allowed to cool naturally to room temperature.

[0061] 3) Finally, cut two panels of the same size from the composite material. Apply a uniform two-component epoxy resin adhesive to one side. Take the aramid honeycomb core, whose nominal cell side length is 3.2 mm, height is 20 mm, and density is 72 kg / m³. 3 The panel is pasted onto the upper and lower sides of the composite core of equal area. The pasted composite honeycomb sandwich panel is clamped with a fixture and cured at room temperature for 8 hours. After removing the fixture, an epoxy resin bonded all-composite sandwich panel is obtained.

[0062] Indentation tests were conducted on the prepared biomimetic plywood sandwich structure according to ASTM D6264 standard. A hemispherical indenter with a diameter of 10 mm was used, and the indentation speed was 1 mm / min. The force-displacement curve results are as follows. Figure 8 As shown, the peak load of the traditional orthogonal layup composite sandwich structure under the same manufacturing process is 1567 N, with an energy absorption of 1.412 J; while the peak load of the biomimetic layup composite sandwich structure is 2136 N, with an energy absorption of 2.323 J. Figure 9 As shown, peak load and energy absorption increased by 36.3% and 64.6% respectively, achieving a significant improvement in the overall structural strength and toughness.

[0063] Although this application discloses preferred embodiments as described above, it is not intended to limit this application. Any person skilled in the art can make possible changes and modifications without departing from the spirit and scope of this application. Therefore, the scope of protection of this application should be determined by the scope defined in the claims of this application.

Claims

1. A high-strength and tough composite material sandwich structure with insect exoskeleton morphology, characterized in that, The sandwich structure includes upper and lower layers of insect exoskeleton-shaped plywood composite material panels and a honeycomb core, with the honeycomb core sandwiched between the two layers of insect exoskeleton-shaped plywood composite material panels. The honeycomb core is made of aramid fiber, and the core height is... T The nominal pore size is 5 ~ 25 mm, and the side length of the pore is... a The core thickness is 3~10 mm. N Its thickness is 0.1 ~ 0.5 mm, and its density is 30 ~ 160 kg / m³. 3 ; The insect exoskeleton-shaped plywood composite panel is formed by hot pressing multiple layers of unidirectional fiber prepreg. The plywood design simultaneously satisfies the requirements of symmetry and balance. If the panel is an even-numbered layer composite material, the fiber directions of each layer on the panel are sequentially arranged as follows: n ×180° / n ×180°- θ / n ×180°-2 θ / … / 2 θ / θ / 0°] s Arrangement, if the panel is an odd-numbered layer composite material, the fiber directions of each layer on the panel are in the following order [( n ×180° / n ×180°- θ / n ×180°-2 θ / … / 2 θ / θ / 0°) s Arranged as / 0°; 5° ≤ θ ≤ 30°, and 180° is θ Integer multiples of.

2. The high-strength and tough composite material sandwich structure for insect exoskeleton morphology according to claim 1, characterized in that, The fiber is at least one of the following: carbon fiber, glass fiber, basalt fiber, or natural fiber.

3. A method for preparing a high-strength and tough composite material sandwich structure with insect exoskeleton morphology as described in claim 1, characterized in that, include: Apply a uniform two-component epoxy resin adhesive to one side of the upper and lower layers of insect exoskeleton morphology composite material panel, and then attach it to the upper and lower sides of the honeycomb core of equal area. The well-fitted insect exoskeleton-shaped high-strength and tough composite material sandwich structure is clamped with a clamp and cured at room temperature for 8 to 10 hours. After removing the clamp, an epoxy resin-bonded insect exoskeleton-shaped high-strength and tough composite material sandwich structure is obtained.

4. The preparation method according to claim 3, characterized in that, The panel preparation method includes: For unidirectional fiber prepregs, with the fiber direction as 0°, the deflection is linearly increased sequentially according to the layup order. θ Cutting is performed; if the panel is an even-numbered layer composite material, then 2 layers are cut. n+2 Prepreg layer ， If the panel is an odd-numbered layer composite material, then it is cut to obtain 2 n +3 prepreg layers; The cut prepregs are stacked layer by layer according to the layup sequence. The stacked prepregs are then hot-pressed and cooled. For even-numbered layer composites, the layup sequence is […]. n ×180° / n ×180°- θ / n ×180°-2 θ / … / 2 θ / θ / 0°] s For odd-layer composite materials, the layup sequence is [( n ×180° / n ×180°- θ / n ×180°-2 θ / … / 2 θ / θ / 0°) s / 0°]; The hot pressing treatment is an autoclave molding process, with a molding pressure of 0.1 ~ 0.8 MPa, a heating rate of 1 ~ 5 ℃ / min, a molding temperature of 125 ~ 150 ℃, and a holding time of 0.5 ~ 2 h.

5. The preparation method according to claim 4, characterized in that, The unidirectional fiber prepreg has a unit area mass of 50 ~ 250 g / m². 2 The thickness is 0.05 ~ 0.25 mm.

6. The preparation method according to claim 4, characterized in that, The cooling process involves reducing the temperature to 80°C at a rate of 2 to 4°C / min, and then allowing it to cool naturally to room temperature.

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