Composite material layup structure, method of manufacture and use

Abstract

The present application relates to a kind of composite material layer structure and preparation method and application, belong to composite material structure design and manufacturing technical field.Composite material layer structure includes multiple sublayers symmetrically distributed along symmetry plane, each sublayer includes multiple composite material unidirectional layers parallel to symmetry plane;Each sublayer has the laying angle α of the sublayer, each sublayer is sequentially arranged according to the direction towards symmetry plane, and the laying direction of composite material unidirectional layer is + α,- α,- α,+ α,- α,+ α,+ α and- α respectively;The laying angle of multiple sublayers arranged along the direction towards symmetry plane gradually increases, and each laying angle is greater than 0 °, and less than 90 °.Each sublayer is independent mechanical decoupling structure unit, and by specific angle sequence, the in-plane tensile-shear coupling stiffness term of the sublayer itself, bending-torsional coupling stiffness term is completely zero, while avoiding hidden remote secondary delamination, ensure the compression residual strength and structural integrity of main load-bearing component after low-speed impact.

Description

Technical Field

[0001] This invention belongs to the field of composite material structure design and manufacturing technology, specifically relating to a composite material layup structure, its preparation method, and its application. Background Technology

[0002] The information disclosed in this background section is intended only to enhance understanding of the overall background of the invention and is not necessarily to be construed as an admission or in any way implying that such information constitutes prior art known to those skilled in the art.

[0003] In fields such as aerospace and rail transportation, structural weight is extremely sensitive. Using fiber-reinforced composite materials to fabricate load-bearing components is an effective way to achieve extreme lightweighting, such as in the fabrication of aircraft wing leading edges or high-speed train body sidewalls. However, in service environments susceptible to low-speed impacts from discrete sources (such as runway debris, hail, and falling objects), such impacts can easily induce internal damage in fiber-reinforced composite components, including matrix cracking, delamination, and even fiber breakage, seriously threatening the remaining strength and service safety of the structure.

[0004] Bouligand structures offer valuable biomimetic insights into overcoming the performance bottlenecks of impact damage tolerance and fracture toughness in composite laminates. Bouligand structures require unidirectional fiber layers to be stacked and rotated layer by layer at specific angle differences, forming an asymmetric helical configuration. In this structure, the continuous change in interlayer fiber orientation induces repeated deflection and bifurcation of matrix cracks induced by impact loads, forcing the cracks to propagate along complex three-dimensional tortuous paths. This significantly increases the fracture surface area through multiple microcrack initiation and propagation mechanisms, dissipating impact kinetic energy and thus achieving superior penetration resistance and energy absorption performance compared to traditional ply structures. However, its non-orthogonal and asymmetric characteristics disrupt the symmetry of the structure's macroscopic stiffness, introducing significant tension-shear coupling and bending-torsion coupling effects into the material constitutive model. This coupling effect causes asymmetric distortion of the local stress field. While enhancing the penetration resistance of areas directly subjected to impact, it generates high levels of interlaminar shear and tensile stresses far from the impact zone, inducing large-scale delamination damage at non-impact points. This hidden secondary damage significantly weakens the overall residual compressive strength of fiber-reinforced composite materials, severely restricting their structural integrity after damage. Summary of the Invention

[0005] To address the shortcomings of existing technologies, the present invention aims to provide a composite material layup structure, its preparation method, and its application. Starting from the constitutive relationship, a layup system that combines mechanical decoupling and helical configuration is constructed. While fully preserving the biomimetic toughening mechanism, the harmful stiffness coupling is systematically eliminated, thereby obtaining a highly reliable composite layer structure with high impact tolerance, no distal secondary damage, and suitable for main load-bearing structures.

[0006] To achieve the above objectives, the technical solution of the present invention is as follows: In a first aspect, a composite material layup structure includes multiple sublayers symmetrically distributed along a plane of symmetry, each sublayer comprising multiple unidirectional composite material layers parallel to the plane of symmetry. Each sublayer has a lay-up angle α, and within each sublayer, in order of orientation toward the plane of symmetry, there are sequentially laid composite unidirectional layers with lay-up directions of +α, -α, -α, +α, -α, +α, +α and -α respectively; The laying angles of the multiple sub-layers arranged along the direction toward the plane of symmetry gradually increase, with each laying angle being greater than 0° and less than 90°.

[0007] Secondly, the preparation method of the above-mentioned composite material layup structure includes the following steps: Set the laying angle α, and lay the composite material unidirectional layer according to the laying directions of +α, -α, -α, +α, -α, +α, and -α to obtain the first sublayer; After increasing the laying angle, continue laying the composite material unidirectional layer according to the increased laying angle to obtain a sublayer with an increased laying angle; Until the calculated sublayer layup angle is greater than 90°, the surface of the last laid sublayer is used as the plane of symmetry. The composite unidirectional layer on the other side of the plane of symmetry is then laid, so that the composite material layup structure is symmetrically distributed along the plane of symmetry.

[0008] Thirdly, a composite material component includes the aforementioned composite material layup structure distributed in a matrix.

[0009] Fourthly, the applications of the aforementioned composite material components include: the manufacture of load-bearing components for use in service environments susceptible to low-velocity impacts from discrete sources.

[0010] The beneficial effects of this invention are as follows: This invention provides a composite material layup structure, in which each sublayer is an independent mechanically decoupled structural unit. The fiber orientation of the unidirectional composite material layer within each sublayer is configured to be geometrically symmetrical about the sublayer's midplane. A specific angle sequence is used to completely zero the in-plane tensile-shear coupling stiffness, bending-torsional coupling stiffness, and in-plane-out-plane coupling stiffness matrix of each sublayer. All sublayers are successively rotated and stacked along the thickness direction of the layup structure at a set angular increment Δθ, preserving the toughening mechanism of three-dimensional deflection and torsional propagation of cracks induced by the biomimetic spiral structure. This design successfully blocks the energy transfer path to the far end caused by the coupling effect, significantly improving the structure's impact damage tolerance, avoiding hidden secondary delamination at the far end, while ensuring the compressive residual strength of the main load-bearing component after low-speed impact. It achieves decoupling of tension / compression and bending moment, and under localized lateral impact, effectively blocks the coupling and transformation of bending deformation into in-plane tensile / compressive strain, thereby avoiding the problem of large-scale secondary damage induced in locations far from the impact zone in the composite material layup structure. Attached Figure Description

[0011] The accompanying drawings, which form part of this invention, are used to provide a further understanding of the invention. The illustrative embodiments of the invention and their descriptions are used to explain the invention and do not constitute an improper limitation of the invention.

[0012] Figure 1 This is a schematic diagram of the structure in Example 1.

[0013] Figure 2 The graph shows the relationship between axial strain and out-of-plane displacement in Test Example 2.

[0014] Figure 3 This is the internal energy-time relationship curve for test example 3.

[0015] Among them, 1. First sublayer; 2. Second sublayer; 3. Third sublayer; 4. Fourth sublayer; 5. Fifth sublayer; 6. Symmetry plane; 11. First unidirectional layer; 12. Second unidirectional layer; 13. Third unidirectional layer; 14. Fourth unidirectional layer; 15. Fifth unidirectional layer; 16. Sixth unidirectional layer; 17. Seventh unidirectional layer; 18. Eighth unidirectional layer. Detailed Implementation

[0016] It should be noted that the following detailed descriptions are exemplary and intended to provide further illustration of the invention. Unless otherwise specified, all technical and scientific terms used herein have the same meaning as commonly understood by one of ordinary skill in the art to which this invention pertains.

[0017] It should be noted that the terminology used herein is for the purpose of describing particular embodiments only and is not intended to limit the scope of exemplary embodiments according to the invention. As used herein, the singular form is intended to include the plural form as well, unless the context clearly indicates otherwise. Furthermore, it should be understood that when the terms "comprising" and / or "including" are used in this specification, they indicate the presence of features, steps, operations, devices, components, and / or combinations thereof.

[0018] One or more embodiments of the present invention provide a composite material layup structure, comprising a plurality of sublayers symmetrically distributed along a plane of symmetry, each sublayer comprising a plurality of unidirectional composite material layers parallel to the plane of symmetry; Each sublayer has a lay-up angle α, and within each sublayer, in order of orientation toward the plane of symmetry, there are sequentially laid composite unidirectional layers with lay-up directions of +α, -α, -α, +α, -α, +α, +α and -α respectively; The laying angles of the multiple sub-layers arranged along the direction toward the plane of symmetry gradually increase, with each laying angle being greater than 0° and less than 90°.

[0019] The above structure integrates a biomimetic spiral structure and a decoupled layup structure: in the thickness direction, the 0° layup direction of each sublayer remains the same while the layup angle increases sequentially, achieving a macroscopic biomimetic spiral configuration and leveraging its advantages in penetration resistance and energy absorption performance; each sublayer is also an independent mechanical decoupling unit, comprising multiple symmetrically or antisymmetrically arranged composite unidirectional layers in the thickness direction; the symmetrical distribution of multiple sublayers along symmetrical planes further optimizes the macroscopic symmetry of the laminate structure. This multi-level symmetrical structure can solve the complex surface caused by asymmetrical layup in existing biomimetic spiral laminates. Deformation problems such as in-plane-out-plane coupling (e.g., tension-bending coupling, compression-bending coupling), tension-shear coupling, compression-shear coupling, and bending-torsion coupling are specifically manifested as follows: In the characterization of classical laminated plate theory, the coupling stiffness matrix is ​​0, the bending-torsion coupling term in the bending stiffness matrix is ​​0, and the in-plane coupling term in the in-plane stiffness matrix is ​​0; thus, a customized laminated structure is constructed that has both biomimetic crack deflection high toughness in macroscopic mechanical performance and the ability to eliminate tension-shear coupling, in-plane-out-plane coupling, and bending-torsion coupling effects, which can suppress the generation of far-end secondary delamination under local impact and improve the residual strength of the structure after impact.

[0020] Optionally, the composite material layup structure has a first sub-layer, a second sub-layer, a third sub-layer, ..., an nth sub-layer arranged sequentially on one side of the plane of symmetry, in the direction closest to the plane of symmetry. The layup angle α1 of the first sub-layer is θ, the layup angle α2 of the second sub-layer is θ+Δθ, the layup angle α3 of the third sub-layer is θ+2Δθ, ..., the layup angle α of the nth sub-layer is α nIt is θ + (n-1)Δθ; by successively accumulating Δθ, the biomimetic toughening mechanism can be completely preserved, while the harmful stiffness coupling effect is eliminated by the ply structure in the sub-plate and the macroscopic symmetry structure set along the symmetry plane.

[0021] Optionally, 0° < θ < 10°, 10° ≤ Δθ ≤ 20°, θ + (n-1)Δθ < 90°, and θ + nΔθ > 90°; the above angle range can achieve a balance between manufacturing cost and mechanical performance.

[0022] Optionally, multiple unidirectional composite material layers in each sublayer are cured into a single composite material matrix, and multiple sublayers in the composite material layup structure are cured into a single composite material matrix; consistent with common layup structures in the art.

[0023] Optionally, the direction of the maximum tensile modulus of the unidirectional composite material layer is consistent with the layup direction.

[0024] One or more embodiments of the present invention provide a method for preparing the above-mentioned composite material layup structure, comprising the steps of: Set the laying angle α, and lay the composite material unidirectional layer according to the laying directions of +α, -α, -α, +α, -α, +α, and -α to obtain the first sublayer; After increasing the laying angle, continue laying the composite material unidirectional layer according to the increased laying angle to obtain a sublayer with an increased laying angle; Until the calculated next sublayer layup angle is greater than 90°, the surface of the last laid sublayer is used as the plane of symmetry, and the composite unidirectional layer on the other side of the plane of symmetry is laid, so that the composite material layup structure is symmetrically distributed along the plane of symmetry.

[0025] In the above process, the laying direction of the unidirectional layer of composite material in the sublayer is determined according to the laying angle of the sublayer. When the base angle of the next sublayer plate group calculated by incremental increment exceeds 90 degrees, it is used as a limiting boundary condition to stop the spiral incremental superposition. Then, taking the top surface of the currently laid structure as the geometric symmetry plane, all the laid sublayer plate groups are globally positive symmetrically stacked to complete the construction of the entire composite plate.

[0026] Optionally, the first sub-layer is the 1st sub-layer with a laying angle α1 of θ; the second sub-layer is the 2nd sub-layer with a laying angle α2 of θ + Δθ; the third sub-layer has a laying angle α3 of θ + 2Δθ, and so on, with the nth sub-layer having a laying angle α... n Let θ be θ + (n-1)Δθ.

[0027] Optionally, after the layup is completed, all unidirectional layers of the composite material are cured together through the composite matrix; this is a conventional method in the field of composite materials.

[0028] One or more embodiments of the present invention provide a composite material component, comprising the aforementioned composite material layup structure distributed in a matrix; wherein the included composite material matrix comprises: a thermosetting resin or a thermoplastic resin, capable of forming a three-dimensional network structure after curing, fixing the position of the internal fibers; wherein the included composite material unidirectional layer is a unidirectional fabric of carbon fiber, basalt fiber, glass fiber or aramid fiber, the layup structure contains only a single type of unidirectional fabric, the single type of unidirectional fabric has consistent performance data, consistent with the assumption that the performance of each composite material unidirectional layer in the aforementioned composite material layup structure is the same, and a composite material layup structure product that conforms to the calculation results can be obtained.

[0029] One or more embodiments of the present invention provide applications of the above-mentioned composite material components, including: for preparing load-bearing components used in service environments susceptible to low-speed impacts from discrete sources, specifically including: the leading edge of an aircraft wing or the side wall of a high-speed train body.

[0030] The present invention will be further described below with reference to specific embodiments.

[0031] Example 1 An impact-resistant composite laminate integrating decoupled layup and biomimetic helical structure comprises multiple unidirectional composite material layers cured into a single composite matrix, such as... Figure 1 As shown, the center position in the thickness direction is the plane of symmetry 6. Multiple sub-layers are arranged from the surface of the impact-resistant composite laminate towards the plane of symmetry 6, including the first sub-layer 1, the second sub-layer 2, the third sub-layer 3, the fourth sub-layer 4, and the fifth sub-layer 5. Each sub-layer includes 8 composite material unidirectional layers parallel to the plane of symmetry 6. For example, in the first sub-layer, the layers arranged in order towards the plane of symmetry 6 are the first unidirectional layer 11, the second unidirectional layer 12, the third unidirectional layer 13, the fourth unidirectional layer 14, the fifth unidirectional layer 15, the sixth unidirectional layer 16, the seventh unidirectional layer 17, and the eighth unidirectional layer 18.

[0032] Each sublayer has a layup angle α, which gradually increases along the direction toward the plane of symmetry 6, with a difference of Δθ. In this embodiment, the layup angle of the first sublayer 1 is α1=θ=5°, and the difference in layup angle between each sublayer and the previous sublayer is Δθ=20°. Therefore, the layup angle of the second sublayer 2 is α2=θ+Δθ=25°, and the layup angle of the third sublayer 3 is α3=θ+2Δθ=45°. Correspondingly, the layup angle of the fourth sublayer 4 is 65°, and the layup angle of the fifth sublayer 5 is 85°. If a sixth sublayer is added, the layup angle of the sixth sublayer is 105°, which exceeds the set boundary condition. Therefore, the surface of the fifth sublayer 5 away from the first sublayer 1 becomes the plane of symmetry 6, and the other side of the plane of symmetry 6 is a completely symmetrical composite unidirectional layer.

[0033] In each sublayer, the unidirectional composite material layers are laid in a direction according to the laying angle. Specifically, in the first sublayer 1, the laying direction of the first unidirectional layer 11 is +5°, the laying direction of the second unidirectional layer 12 is -5°, the laying direction of the third unidirectional layer 13 is -5°, the laying direction of the fourth unidirectional layer 14 is +5°, the laying direction of the fifth unidirectional layer 15 is -5°, the laying direction of the sixth unidirectional layer 16 is +5°, the laying direction of the seventh unidirectional layer 17 is +5°, and the laying direction of the eighth unidirectional layer 18 is -5°.

[0034] The impact-resistant composite laminate in this embodiment has 10 sub-layers and 80 layups.

[0035] Preparation methods include: Set the laying angle α1=θ=5°, and lay the composite material unidirectional layers sequentially in the laying directions of +5°, -5°, -5°, +5°, -5°, +5°, +5° and -5° to obtain the first sublayer 1; Set the laying angle α2=θ+Δθ=25°, and lay the composite material unidirectional layer sequentially in the laying directions of +25°, -25°, -25°, +25°, -25°, +25°, +25° and -25° to obtain the second sub-layer 2; The third sublayer 3, the fourth sublayer 4, and the fifth sublayer 5 are prepared by setting the laying angles to 45°, 65°, and 85° respectively. The surface of the fifth sublayer 5 away from the first sublayer 1 is taken as the symmetry plane 6, and a completely symmetrical composite unidirectional layer is set on the other side of the symmetry plane 6. The structure obtained by the layup is cured as a whole, so that all the unidirectional layers of composite materials are cured into one piece through the composite matrix, resulting in an impact-resistant composite plate with fused decoupled layup and biomimetic spiral structure.

[0036] Comparative Example 1 This comparative example provides an impact-resistant composite laminate with staggered ply, comprising multiple composite unidirectional layers cured into one piece by a composite matrix. The ply materials and composite matrix are the same as in Example 1. It consists of 80 ply layers, each with the same thickness as in Example 1. Each composite unidirectional layer is laid with its ply direction rotated by +9° compared to the previous composite unidirectional layer.

[0037] Comparative Example 2 This comparative example provides an impact-resistant composite laminate with staggered ply, comprising multiple composite unidirectional layers cured into one piece by a composite matrix. The ply materials and composite matrix are the same as in Example 1. It consists of 80 ply layers, each with the same thickness as in Example 1. Each composite unidirectional layer is laid with its ply direction rotated by +45° compared to the previous composite unidirectional layer.

[0038] Example 2 A composite material component includes the impact-resistant composite laminate of Example 1, wherein the matrix of the composite material is a thermosetting resin and the inner composite unidirectional layer is made of carbon fiber unidirectional fabric, which can be used in service environments where it is susceptible to low-speed impacts from discrete sources.

[0039] Test Example 1 Based on the stiffness matrix calculation method in the patent application number CN2026103258376, the performance of the impact-resistant composite laminates in Example 1, Comparative Example 1, and Comparative Example 2 is calculated, including the following steps: Step 1: Determine the engineering elastic constants of the single-layer plate; obtain the material properties of the reinforcing fiber and resin matrix, and determine the four basic engineering elastic constants in the principal directions of the composite unidirectional layer material, specifically including: : Elastic modulus in the fiber direction (longitudinal modulus); : Elastic modulus perpendicular to the fiber direction (transverse modulus); Poisson's ratio; In-plane shear modulus; Step 2: Using the engineering elastic constants obtained in Step 1, calculate the reduced stiffness matrix of the single-layer plate in the material principal direction coordinate system (axis 1-2). Each component Thus, the reduced stiffness matrix is ​​obtained. In this context, axis 1 is defined as the main fiber direction of the unidirectional layer (longitudinal direction, i.e., the direction of maximum tensile modulus), and axis 2 is defined as the direction perpendicular to axis 1 (transverse direction) within the plane of the unidirectional layer. Specifically, it includes: ; ; ; ; in, To satisfy the reciprocal relationship.

[0040] Step 3: For ply angles of... For any unidirectional layer, its stiffness needs to be transformed from the material principal axis coordinate system to the laminate global coordinate system (xy axis) to obtain the off-axis stiffness matrix. The various components of . Thus, the off-axis stiffness matrix is ​​obtained. The specific formula for its component transformation is as follows: ; ; ; ; ; ; Step 4: Obtain the off-axis stiffness matrix of all unidirectional layers. Then, integration is performed along the thickness direction to solve for the overall stiffness matrix of the laminate, which is expressed as follows: , in, N For the combined force of internal forces, M For the resultant torque, A Here is the in-plane stiffness matrix. B The in-plane / out-of-plane coupling stiffness matrix is... D Here is the bending stiffness matrix. For mid-surface strain, The curvature of the mid-surface; Specifically: ; Wherein, the subscripts x and y represent the in-plane orthogonal principal directions in the global coordinate system of the laminate, and the subscript xy represents the shear or torsion direction in the xy plane of the global coordinate system; ; ; ; in, For ply numbering, From the geometric mid-surface of the laminate to the first The distance between the upper surfaces of the layer.

[0041] Step 5: Solve for Example 1, Comparative Example 1, and Comparative Example 2. Since the materials, thickness, and number of layers are the same in Example 1, Comparative Example 1, and Comparative Example 2, this calculation example is based on a typical carbon fiber / epoxy resin composite material system, specifically including: E1 = 131 GPa; E2 = 9 GPa; G 12 =5.4 GPa; v 12 =0.3; There are a total of 80 layers, therefore The range is 1 to 80; The single-layer ply thickness is 0.125mm, through Multiply by 0.125mm to obtain the corresponding values. .

[0042] Seeking After obtaining the matrix, the stress state or deformation of the entire laminate structure can be obtained through the overall stiffness matrix of the laminate.

[0043] The calculation results of Example 1 obtained As shown in Table 1, the calculation results of Comparative Example 1 As shown in Table 2, the calculation results of Comparative Example 2 As shown in Table 3.

[0044] Table 1. Overall stiffness matrix of laminated plates in Example 1

[0045] Table 2. Overall stiffness matrix of laminated plates in Comparative Example 1

[0046] Table 3. Overall stiffness matrix of laminated plates in Comparative Example 2

[0047] It can be seen that in Table 2 corresponding to Comparative Example 1 and Table 3 corresponding to Comparative Example 2, in-plane-out-of-plane coupling exists respectively. The matrix is ​​not zero, indicating significant bending-torsional coupling. (Not 0). Table 1 corresponding to Example 1 shows all coupling terms as 0, exhibiting pure shear characteristics. It achieves a complete zeroing of tension-shear, bending-torsion, and tension-bending coupling terms, exhibiting orthotropic force characteristics; and since the layups in Example 1, Comparative Example 1, and Comparative Example 2 are all uniform layups, the tension-shear coupling terms... All are 0.

[0048] On the other hand, during the actual preparation of composite products, the difference in the anisotropic thermal expansion coefficients between the layers during the high-temperature curing and cooling stage can induce significant thermal residual stress. When the coupling stiffness matrix [B] is not zero, the temperature stress generated by in-plane thermal deformation will cause the overall warping deformation of the laminate during the curing process, generating internal stress and affecting the molding accuracy; while the coupling stiffness matrix [B] of Example 1 is 0, which can effectively avoid curing warping deformation and further improve product quality.

[0049] Test Example 2 Axial strain-out-of-plane displacement relationship simulation was performed according to ASTM D3039 / D3039M, using a simulation time of 0.01 min. -1 The strain rate was increased until the strain reached 0.03, and the axial strain versus out-of-plane displacement curves of the laminates of Example 1, Comparative Example 1, and Comparative Example 2 under uniaxial tensile load were obtained, as shown in the figure. Figure 2As shown. Compared with Comparative Example 1 and Comparative Example 2, the layup design scheme of Example 1 successfully eliminated out-of-plane coupling deformation under axial tensile load by optimizing the stiffness matching inside the laminate, and the out-of-plane displacement remained at 0, exhibiting extremely excellent zero out-of-plane deformation characteristics.

[0050] Test Example 3 Low-velocity impact simulations were performed according to ASTM D7136 standard to obtain the internal energy-time relationship curves of the laminates in Example 1, Comparative Example 1, and Comparative Example 2, as follows: Figure 3 As shown, Example 1 exhibits a peak internal energy of 54.5 J at the impact limit position, demonstrating excellent dynamic energy absorption capability. Furthermore, during the subsequent unloading and rebound phase, the layup releases a rebound energy of up to 41 J, with a structural recovery rate significantly superior to Comparative Example 1 and Comparative Example 2. This fully demonstrates the superior structural safety protection effect of the layup structure in the face of large-energy impacts.

[0051] The above description is merely a preferred embodiment of the present invention and is not intended to limit the invention. Various modifications and variations can be made to the present invention by those skilled in the art. Any modifications, equivalent substitutions, improvements, etc., made within the spirit and principles of the present invention should be included within the scope of protection of the present invention.

Claims

1. A composite material layup structure, characterized in that, It includes multiple sub-layers symmetrically distributed along the plane of symmetry, and each sub-layer includes multiple unidirectional composite material layers parallel to the plane of symmetry; Each sublayer has a lay-up angle α, and within each sublayer, in order of orientation toward the plane of symmetry, there are sequentially laid composite unidirectional layers with lay-up directions of +α, -α, -α, +α, -α, +α, +α and -α respectively; The laying angles of the multiple sub-layers arranged along the direction toward the plane of symmetry gradually increase, with each laying angle being greater than 0° and less than 90°.

2. The composite material layup structure as described in claim 1, characterized in that, The composite material layup structure has sublayers arranged sequentially on one side of the plane of symmetry, with sublayers numbered 1, 2, 3, ... n, in the direction closest to the plane of symmetry. The layup angle α1 of the 1st sublayer is θ, the layup angle α2 of the 2nd sublayer is θ + Δθ, the layup angle α3 of the 3rd sublayer is θ + 2Δθ, ... the layup angle α of the nth sublayer is α n Let θ be θ + (n-1)Δθ.

3. The composite material layup structure as described in claim 1, characterized in that, 0°<θ<10°, 10°≤Δθ≤20°, θ+(n-1)Δθ<90°, and θ+nΔθ>90°.

4. The composite material layup structure as described in claim 1, characterized in that, Multiple unidirectional composite material layers in each sublayer are cured into a single composite material matrix, and multiple sublayers in the composite material layup structure are cured into a single composite material matrix.

5. A composite material layup structure as described in claim 1, characterized in that, The direction of the maximum tensile modulus of the unidirectional layer of the composite material is consistent with the laying direction.

6. A method for preparing a composite material layup structure as described in any one of claims 1 to 5, characterized in that, Including the following steps: Set the laying angle α, and lay the composite material unidirectional layer according to the laying directions of +α, -α, -α, +α, -α, +α, and -α to obtain the first sub-layer; After increasing the laying angle, continue laying the composite material unidirectional layer according to the increased laying angle to obtain a sublayer with an increased laying angle; Until the calculated next sublayer layup angle is greater than 90°, the surface of the last laid sublayer is used as the plane of symmetry, and the composite unidirectional layer on the other side of the plane of symmetry is laid, so that the composite material layup structure is symmetrically distributed along the plane of symmetry.

7. The method for preparing the composite material layup structure as described in claim 6, characterized in that, After the layup is completed, all unidirectional layers of composite material are cured together through the composite matrix.

8. The method for preparing the composite material layup structure as described in claim 6, characterized in that, The first sublayer is the 1st sublayer with a laying angle α1 of θ; the second sublayer is the 2nd sublayer with a laying angle α2 of θ + Δθ; the third sublayer has a laying angle α2 of θ + 2Δθ, ... the nth sublayer has a laying angle αn of θ + (n-1)Δθ.

9. A composite material component, characterized in that, Includes composite material layup structures as described in any one of claims 1-5 distributed in the matrix.

10. An application of a composite material component as described in claim 9.

Images