A design method for aircraft composite structures considering low-speed impact performance

Abstract

The present invention belongs to the technical field of composite materials, and discloses a method for designing aircraft composite material structures taking low-speed impact performance into consideration. A finite element model is established according to actual working conditions, material properties are assigned to the finite element model, and the finite element model is meshed. The boundary conditions of the finite element model are set according to actual working conditions, and an in-plane impact damage model and an out-of-plane impact damage model are constructed. Based on the above damage model, a damage mechanism is obtained, and a new impact-resistant structure is proposed with reference to biological impact-resistant damage structures. Data is then obtained through the above damage model, and an enhancement mechanism analysis is performed to finally obtain an optimal design. The beneficial effects of the present invention are: it can simultaneously take into account the in-plane impact resistance and the out-of-plane impact resistance, accurately, effectively, and economically comprehensively study the damage of composite laminate structures under low-speed impact loads, and thus perform structural design.

Description

Technical Field

[0001] The invention relates to the technical field of composite materials, and in particular to a method for designing aircraft composite material structures taking low-speed impact performance into consideration. Background Art

[0002] Compared with metal materials, composite materials have the characteristics of light weight, high specific strength and high specific stiffness, and are currently widely used in many fields. Unfortunately, composite laminates are extremely vulnerable to damage from accidental low-speed impact events, such as hail, debris, and dropped tools. These events are often mixed with in-plane impact and out-of-plane edge impact. Low-speed impact damage usually does not leave obvious marks on the impact surface, but it will cause serious internal damage, which greatly reduces the mechanical properties of the material and causes immeasurable damage. The impact sensitivity of composite laminates limits their application. Therefore, it is very necessary to study the failure mechanism of composite laminates under low-speed impact.

[0003] Typical damage of composite laminates under in-plane impact can be divided into fiber fracture, matrix cracking and delamination. Although edge impact damage is similar to in-plane impact damage, both of which are manifested as fiber fracture, matrix cracking and delamination, edge impact damage has obvious characteristics such as out-of-plane semi-elliptical damage area, wedge effect and multiple matrix cracks along the length direction. In addition, the mechanical response of out-of-plane edge impact test is more similar to progressive crushing test than in-plane impact. Generally, peak load is usually observed at the beginning of edge impact, followed by a progressive stage characterized by nearly constant force. It is worth noting that a larger damage area is detected for out-of-plane edge impact than in-plane impact and near-edge impact at the same impact energy. This is mainly due to the fact that the load of edge-impacted laminates is highly localized to the area below the impactor, which induces lower energy required for edge impact to achieve almost invisible impact damage. A large number of studies usually focus on normal in-plane center impact far away from the edge, both experimentally and numerically, and the failure mechanism is well established, while edge impact cases have not received enough attention, and methods for comprehensive analysis and enhanced design are even more lacking.

[0004] Therefore, it is necessary to provide a design method for aircraft composite structures that takes low-speed impact performance into consideration. By accurately simulating the in-plane impact and edge impact of the laminate structure, the design of the laminate structure's resistance to low-speed impact can be effectively carried out in all directions and the cost can be reduced. The damage of the composite laminate structure under low-speed impact loads can be accurately, effectively and economically studied in all directions, thereby carrying out structural design. Summary of the invention

[0005] The present invention discloses a method for designing aircraft composite material structures taking low-speed impact performance into consideration, and a method for designing composite material laminate structures to resist low-speed impact during service and maintenance. Composite material laminate structures are extremely susceptible to low-speed impacts, such as hail, falling hammers, etc. during service and maintenance. These low-speed impact events can be divided into in-plane low-speed impacts and edge low-speed impacts. By accurately simulating the in-plane impacts and edge impacts of the laminate structures, the design of the low-speed impact resistance of the laminate structures in all directions can be effectively carried out and the cost can be reduced, which can effectively solve the technical problems involved in the background technology.

[0006] To achieve the above object, the technical solution of the present invention is:

[0007] A method for designing a composite structure of an aircraft considering low-speed impact performance comprises the following steps:

[0008] Step 1: Establish a finite element model based on actual working conditions;

[0009] Step 2: Assign material properties to the finite element model;

[0010] Step 3: Mesh the finite element model;

[0011] Step 4: Set the boundary conditions of the finite element model according to the actual working conditions;

[0012] Step 5: Construct in-plane impact damage model and out-of-plane impact damage model;

[0013] Step 6: Get and integrate simulation results;

[0014] Step 7: Compare and analyze all impact responses and failure models;

[0015] Step 8: Refer to the impact-resistant structure, comprehensively consider the in-plane low-speed impact resistance performance and the out-of-plane low-speed impact resistance performance, and carry out structural reinforcement design;

[0016] Step 9: Carry out steps 5 to 7 again for the designed new structure;

[0017] Step 10: Analyze the strengthening mechanism of the structure with strengthening effect;

[0018] Step 11: Based on the reinforcement mechanism analysis, comprehensively consider the in-plane low-speed impact resistance performance and the out-of-plane low-speed impact resistance performance, optimize the structural parameters, repeat steps 8 to 10 for optimization iteration, and finally obtain the optimal design.

[0019] A design method for low-speed impact resistance of composite laminate structures during service and maintenance. Composite laminate structures are extremely susceptible to low-speed impacts during service and maintenance, such as hail and falling hammers. These low-speed impact events can be divided into in-plane low-speed impacts and edge low-speed impacts. By accurately simulating the in-plane impact and edge impact of the laminate structure, the design of the low-speed impact resistance of the laminate structure in all directions can be effectively carried out and the cost can be reduced. The rate range of low-speed impact is generally 1~10m / s, and the specific value depends on the stiffness, component properties, punch mass, stiffness, etc. of the impacted material. The composite laminate structure is a laminate structure formed by stacking and bonding multiple single-layer plates. Based on the design method of the present invention, the damage of the composite laminate structure under low-speed impact loads can be accurately, effectively and economically studied in all directions, so as to carry out structural design.

[0020] As a preferred improvement of the present invention: in the step eight, the impact-resistant structure includes a biological impact-resistant structure.

[0021] As a preferred improvement of the present invention: the in-plane impact damage model is as follows, the failure modes of the fiber and the matrix are mainly divided into five failure modes, namely, fiber tensile fracture, fiber compression buckling and kinking, matrix cracking under transverse tension, matrix crushing under transverse compression, and in-plane shear failure, and the corresponding damage criteria are as follows:

[0022] Fiber tensile fracture : (1);

[0023] Fiber compression buckling and kinking : (2);

[0024] Matrix tensile cracking : (3);

[0025] Matrix compression and crushing : (4);

[0026] In-plane shear failure : (5);

[0027] In the formula, , and is the effective stress tensor, , , , and They are fiber tensile strength, fiber compressive strength, matrix tensile strength, matrix compressive strength, and shear strength in 12 directions; F ftis the longitudinal tensile failure factor, F fc is the longitudinal compression failure factor, F mt is the transverse tensile failure factor, F mc is the lateral compression failure factor, F s is the in-plane shear failure factor.

[0028] As a preferred improvement of the present invention: the out-of-plane impact damage model is as follows. When the composite material is subjected to a low-speed impact, the specific failure criterion is as follows:

[0029] Fiber tensile fracture ( ): (26);

[0030] Fiber compression failure ( ): (27);

[0031] Matrix cracking ( ): (28);

[0032] Matrix extrusion ( ):

[0033] (29);

[0034] In the formula (i,j=1,2,3) are stress components, 1 is the fiber direction, 2 is the matrix direction, 3 is the out-of-plane normal direction, s is the 12 shear direction, X 13 , X 23 They are the shear strength in the 13 and 23 directions respectively. As long as the stress state of the unit satisfies the equation, the corresponding failure mode will appear, and then the material will begin to be damaged.

[0035] As a preferred improvement of the present invention: the composite material is a laminate structure formed by stacking and bonding multiple single-layer boards, and adjacent single-layer boards are bonded via a bonding interface, and the bonding interface is made by the following steps:

[0036] S1. Laying the fiber cloth in layers according to a preset laying angle, applying a layer of a mixture of resin and curing agent after each layer is laid;

[0037] S2. Apply release agent to the upper and lower mold plates, and place the laid fiber cloth between the molds after the release agent is dry;

[0038] S3, hot pressing and curing the fiber cloth in the mold;

[0039] S4. Open the mold.

[0040] As a preferred improvement of the present invention: the velocity range of the low-speed impact is generally 1-10 m / s, and the specific value depends on the stiffness and component properties of the impacted material, as well as the mass and stiffness of the punch.

[0041] The beneficial effects of the present invention are as follows:

[0042] 1. The present invention provides a set of failure criteria and corresponding damage evolution criteria, including an in-plane impact damage model and an out-of-plane impact damage model. The established finite element model can accurately capture the failure mode and load response of the laminate structure under various impact loads, providing important data for analyzing and understanding the damage mechanism;

[0043] 2. The present invention provides a complete low-speed impact resistance design method for composite laminates during service maintenance, etc. Based on the above-mentioned damage model, the damage mechanism is obtained, and a new impact-resistant structure is proposed with reference to the biological impact-resistant damage structure. Then, data is obtained through the above-mentioned damage model, and the enhancement mechanism analysis is carried out. Finally, the optimal design is obtained, which can take into account both the in-plane impact resistance and the out-of-plane impact resistance, and accurately, effectively and economically study the damage of the composite laminate structure under low-speed impact load in an all-round way, so as to carry out structural design. BRIEF DESCRIPTION OF THE DRAWINGS

[0044] In order to more clearly illustrate the technical solutions in the embodiments of the present invention, the following briefly introduces the drawings required for describing the embodiments. Obviously, the drawings described below are only some embodiments of the present invention. For those skilled in the art, other drawings can be obtained based on these drawings without creative work, among which:

[0045] Figure 1 A schematic diagram of a method for designing a composite structure of an aircraft taking low-speed impact performance into consideration according to the present invention;

[0046] Figure 2 A flowchart of a method for designing a composite structure of an aircraft taking low-speed impact performance into consideration according to the present invention;

[0047] Figure 3 Schematic diagram of the in-plane impact finite element model and the out-of-plane impact finite element model;

[0048] Figure 4 This is the validation diagram of the in-plane impact finite element model;

[0049] Figure 5 This is the verification diagram of the out-of-plane impact finite element model;

[0050] Figure 6 Schematic diagrams of three helical configurations;

[0051] Figure 7 The comparison diagram of in-plane impact response between bionic spiral laminate and traditional laminate;

[0052] Figure 8 Comparison of out-of-plane impact responses of bionic spiral laminates and traditional laminates.

[0053] In the figure: 1 - in-plane impact finite element model, 2 - out-of-plane impact finite element model. DETAILED DESCRIPTION

[0054] The technical solutions in the embodiments of the present invention will be described clearly and completely below in combination with the embodiments of the present invention. Obviously, the described embodiments are only part of the embodiments of the present invention, not all of the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by ordinary technicians in this field without creative work are within the scope of protection of the present invention.

[0055] It should be noted that all directional indications (such as up, down, left, right, front, back, etc.) in the embodiments of the present invention are only used to explain the relative position relationship, movement status, etc. between the components under a certain specific posture (as shown in the accompanying drawings). If the specific posture changes, the directional indication will also change accordingly.

[0056] In addition, in the present invention, descriptions such as "first", "second", etc. are only used for descriptive purposes and cannot be understood as indicating or implying their relative importance or implicitly indicating the number of the indicated technical features. Therefore, the features defined as "first" or "second" may explicitly or implicitly include at least one of the features. In the description of the present invention, the meaning of "plurality" is at least two, such as two, three, etc., unless otherwise clearly and specifically defined.

[0057] In the present invention, unless otherwise clearly specified and limited, the terms "connection", "fixation", etc. should be understood in a broad sense. For example, "fixation" can be a fixed connection, a detachable connection, or an integral connection; it can be a mechanical connection or an electrical connection; it can be a direct connection or an indirect connection through an intermediate medium, it can be the internal connection of two elements or the interaction relationship between two elements, unless otherwise clearly defined. For ordinary technicians in this field, the specific meanings of the above terms in the present invention can be understood according to specific circumstances.

[0058] In addition, the technical solutions between the various embodiments of the present invention can be combined with each other, but it must be based on the fact that ordinary technicians in the field can implement it. When the combination of technical solutions is contradictory or cannot be implemented, it should be deemed that such combination of technical solutions does not exist and is not within the scope of protection required by the present invention.

[0059] See also Figure 1As shown, the present invention provides a method for designing a composite structure of an aircraft considering low-speed impact performance, comprising the following steps:

[0060] Step 1: Establish a finite element model based on actual working conditions;

[0061] Step 2: Assign material properties to the finite element model;

[0062] Step 3: Mesh the finite element model;

[0063] Step 4: Set the boundary conditions of the finite element model according to the actual working conditions (mainly the displacement and force of the boundary);

[0064] Step 5: Construct in-plane impact damage model and out-of-plane impact damage model;

[0065] As an implementation mode, the in-plane impact damage model is as follows. The failure modes of the fiber and the matrix are mainly divided into five failure modes, namely, fiber tensile fracture, fiber compression buckling and kinking, matrix cracking under transverse tension, matrix crushing under transverse compression, and in-plane shear failure. The corresponding damage criteria are as follows:

[0066] Fiber tensile fracture : (1);

[0067] Fiber compression buckling and kinking : (2);

[0068] Matrix tensile cracking : (3);

[0069] Matrix compression and crushing : (4);

[0070] In-plane shear failure : (5);

[0071] In the formula, , and is the effective stress tensor, , , , and They are fiber tensile strength, fiber compressive strength, matrix tensile strength, matrix compressive strength, and shear strength in 12 directions; F ft is the longitudinal tensile failure factor, F fc is the longitudinal compression failure factor, F mtis the transverse tensile failure factor, F mc is the lateral compression failure factor, F s is the in-plane shear failure factor. Once one of the above five damage criteria is met, the material will gradually degrade according to the nonlinear damage evolution rule, as shown in Formula 7.

[0072] (6)

[0073] Where: dij ( i=ft,fc,mt,mc )and d s The damage variable is used to indicate the degree of material damage; is the elastic energy of the corresponding material; is the corresponding material fracture energy; is the unit characteristic length; is the shear response parameter used for calibration; is the maximum shear damage variable; is the time damage threshold variable, is the damage threshold, and The calculation formula is as follows:

[0074] (7)

[0075] (8)

[0076] in t is the current time, is the time variable; and are the modulus of the laminate in the fiber direction and the modulus in the matrix direction, About time The damage threshold function.

[0077] (9)

[0078] in, = ; ; For in time The corresponding effective stress tensor is For the time The corresponding intensity.

[0079] As another preferred option:

[0080] Fiber stretch damage: (10);

[0081] Fiber compression damage: (11);

[0082] Matrix tensile damage: (12);

[0083] Matrix compression damage: (13);

[0084] In the formula , , , They are fiber tensile damage, fiber compression damage, matrix tensile damage, and matrix compression damage; is the coefficient of the contribution of shear stress to the fiber tension onset criterion, and the response of the damaged material is calculated as follows:

[0085] (14);

[0086] (15)

[0087] In the formula is the stress vector, is the strain vector, is the damage elastic matrix, , , They are shear modulus, Poisson's ratio in 12 direction, and Poisson's ratio in 21 direction; is the intra-layer damage variable; , , , , They are fiber tensile damage variable, fiber compression damage variable, matrix tensile damage variable, matrix compression damage variable, and shear damage variable, and the calculation formula is as follows:

[0088] (16)

[0089] where d f d m are fiber damage variables and matrix damage variables, respectively. The evolution of intralayer damage is based on the crack band model. Assuming linear degradation, the damage variables are calculated as follows:

[0090] (17)

[0091] Where: In destruction mode The equivalent displacement of and The material reaches the failure mode The initial damage equivalent displacement and final damage equivalent displacement of are expressed as follows:

[0092] (18)

[0093] (19)

[0094] Where: is the initial damage criterion value of the component material, is the equivalent stress of the initial damage of the component material, is the fracture energy density of the component material.

[0095] The damage is layered as follows:

[0096] According to the linear elastic constitutive relation, the linear elastic deformation stage is expressed as:

[0097] (20)

[0098] (twenty one)

[0099] in, is the normal traction stress, and are two tangential stresses; is the normal interface stiffness coefficient, and are the two tangential interface stiffness coefficients; is the normal displacement, and are two tangential displacements; is the interface damage variable (initial value is 0); , , are the three current traction stresses.

[0100] As the interface traction stress increases, initial damage will occur on the interface. The secondary nominal stress is used to describe the initial damage of the interface:

[0101] (twenty two)

[0102] in, is the normal interface strength, and are the two tangential interface strengths.

[0103] After the initial interface damage criterion is met, the bonding interface begins to debond or delaminate. The damage evolution model of the interface is represented by the linear damage evolution criterion or the nonlinear damage evolution criterion:

[0104] Linear damage evolution criterion:

[0105] (twenty three)

[0106] in, is the maximum damage displacement during the interface damage process, is the initial damage displacement of the interface, is the displacement of complete interface damage.

[0107] Nonlinear damage evolution criterion:

[0108] (twenty four)

[0109] in, is the effective traction stress, is the total fracture energy of the interface, is the interface elastic energy;

[0110] After reaching a certain energy, the interface begins to completely debond and delaminate:

[0111] (25)

[0112] In the formula, is the fracture energy, is the critical fracture energy, , ; is the viscosity parameter; , , is the work done by the stress in the normal direction and the two tangential directions under the corresponding displacement.

[0113] As for out-of-plane impact, the damage model is as follows:

[0114] When composite materials are subjected to low-speed impact, their damage forms are mainly fiber damage, matrix damage and delamination. The specific damage criteria are as follows:

[0115] Fiber tensile fracture ( ): (26);

[0116] Fiber compression failure ( ): (27);

[0117] Matrix cracking ( ): (28);

[0118] Matrix extrusion ( ):

[0119] (29);

[0120] In the formula (i,j=1,2,3) are stress components, 1 is the fiber direction, 2 is the matrix direction, 3 is the out-of-plane normal direction, s is the 12 shear direction, X 13 , X 23 They are the shear strength in the 13 and 23 directions respectively. As long as the stress state of the unit satisfies the equation, the corresponding failure mode will appear, and then the material will begin to be damaged.

[0121] Fiber tensile fracture; (30)

[0122] Fiber compression failure: (31)

[0123] Matrix tensile cracking: (32)

[0124] Matrix compression failure: (33)

[0125] In the formula, and are the reduced variables of the fiber direction modulus and the matrix direction modulus, respectively. , is the reduction variable of shear modulus in 12 and 23 directions; is the modulus in the 23 direction before reduction; , , , They are fiber tensile reduction variable, fiber compression reduction variable, matrix tensile reduction variable, matrix compression reduction variable, and their calculation formulas are as follows:

[0126] (34)

[0127] In the formula t T It is the total time from the start of impact to the time when the kinetic energy of the punch reaches zero.

[0128] As another preferred option:

[0129] The formation of microcracks is caused by lateral and shear loading of the matrix. These microcracks include splitting, shear banding and microcracks. W E The micro damage evolution equation is inverted and can be expressed as:

[0130] (35)

[0131] In the formula , , are the strains in the corresponding directions, As a constant, the reduced damage variable is introduced , S is the micro damage dissipation, which is considered to be a constant at the macro crack extension point, and the micro damage evolution equation is obtained:

[0132] (36)

[0133] In the formula , The original transverse modulus and shear modulus, , are the lateral and shear micro-damage functions respectively, and the failure criterion for the onset of failure is as follows:

[0134] (37)

[0135] In the formula , are the tensile and compressive failure strains in the fiber direction, respectively.

[0136] (38)

[0137] In the formula , are the transverse tensile and compressive failure strains, Z Longitudinal shear failure strain. Assuming that when either of the equations (37) or (38) is satisfied, i.e., the fiber has failed, the matrix-dominated failure mode initiates the post-peak response at its current traction force, where matrix failure does not initiate fiber failure. The moduli associated with each failure mode can now be degraded in a continuous and smooth manner by the traction separation law, and the material properties can be described by the stress-strain constitutive relation:

[0138] (39)

[0139] In the formula , , They are the failure parameters that control the degradation in the fiber direction, matrix direction, and shear direction, respectively. The parameter value range is from 0 (no failure) to 1 (complete failure). The out-of-plane impact delamination model is consistent with the in-plane impact delamination model.

[0140] Step 6: Get and integrate simulation results;

[0141] Step 7: Compare and analyze all impact responses and failure models;

[0142] Step 8: Refer to the impact-resistant structure, comprehensively consider the in-plane low-speed impact resistance performance and the out-of-plane low-speed impact resistance performance, and carry out structural reinforcement design;

[0143] As a preferred solution: many biological structures with excellent impact resistance can be observed through an electron microscope to have a spiral arrangement of biological fibers. The fibers in the laminate are arranged in a spiral shape, and the spiral arrangement of fibers is designed to improve the laminate's resistance to frontal impact and side impact.

[0144] Step 9: Carry out steps 5 to 7 again for the designed new structure;

[0145] Step 10: Analyze the strengthening mechanism of the structure with strengthening effect as a preferred solution, and compare and analyze its structural parameters, such as the linear and nonlinear influence of the helical angle, the angle change rate, multi-helical distribution, the combination of helical distribution and non-helical distribution, etc., to obtain the parameter influence law;

[0146] Step 11: Based on the reinforcement mechanism analysis, comprehensively consider the in-plane low-speed impact resistance performance and the out-of-plane low-speed impact resistance performance, optimize the structural parameters, repeat steps 8 to 10 for optimization iteration, and finally obtain the optimal design. Example

[0147] See also Figure 1 and Figure 2 The present invention discloses a method for designing aircraft composite material structures taking low-speed impact performance into consideration. First, an in-plane impact finite element model and an out-of-plane impact finite element model are established according to actual working conditions. Figure 3 As shown;

[0148] Step 2: Assign material properties to the finite element model;

[0149] Step 3: Mesh the finite element model;

[0150] Step 4: Set the boundary conditions of the finite element model according to the actual working conditions;

[0151] Step 5: construct an in-plane impact damage model and an out-of-plane impact damage model; as a preferred improvement of the present invention, the in-plane impact damage model is as follows;

[0152] The failure modes of fibers and matrices are mainly divided into five failure modes, namely, fiber tensile fracture, fiber compression buckling and kinking, matrix cracking under transverse tension, matrix crushing under transverse compression, and in-plane shear failure. The corresponding damage criteria are as follows:

[0153] Fiber tensile fracture : (1);

[0154] Fiber compression buckling and kinking : (2);

[0155] Matrix tensile cracking : (3);

[0156] Matrix compression and crushing : (4);

[0157] In-plane shear failure : (5);

[0158] In the formula, , and is the effective stress tensor, , , , and They are fiber tensile strength, fiber compressive strength, matrix tensile strength, matrix compressive strength, and shear strength in 12 directions; F ft is the longitudinal tensile failure factor, F fc is the longitudinal compression failure factor, F mt is the transverse tensile failure factor, F mc is the lateral compression failure factor, F s is the in-plane shear failure factor. Once one of the above five damage criteria is met, the material will gradually degrade according to the nonlinear damage evolution rule, as shown in Formula 7.

[0159] (6)

[0160] Where: dij ( i=ft,fc,mt,mc )and d s The damage variable is used to indicate the degree of material damage; is the elastic energy of the corresponding material; is the corresponding material fracture energy; is the unit characteristic length; is the shear response parameter used for calibration; is the maximum shear damage variable; is the time damage threshold variable, is the damage threshold, and The calculation formula is as follows:

[0161] (7)

[0162] (8)

[0163] in t is the current time, is the time variable; and are the modulus of the laminate in the fiber direction and the modulus in the matrix direction, About time The damage threshold function.

[0164] According to the linear elastic constitutive relation, the linear elastic deformation stage is expressed as:

[0165] (20)

[0166] (twenty one)

[0167] in, is the normal traction stress, and are two tangential stresses; is the normal interface stiffness coefficient, and are the two tangential interface stiffness coefficients; is the normal displacement, and are two tangential displacements; is the interface damage variable (initial value is 0); , , are the three current traction stresses.

[0168] As the interface traction stress increases, initial damage will occur on the interface. The secondary nominal stress is used to describe the initial damage of the interface:

[0169] (twenty two)

[0170] in, is the normal interface strength, and are the two tangential interface strengths.

[0171] After the initial interface damage criterion is met, the bonding interface begins to debond or delaminate. The damage evolution model of the interface is represented by the linear damage evolution criterion or the nonlinear damage evolution criterion:

[0172] Linear damage evolution criterion:

[0173] (twenty three)

[0174] in, is the maximum damage displacement during the interface damage process, is the initial damage displacement of the interface, is the displacement of complete interface damage.

[0175] Nonlinear damage evolution criterion:

[0176] (twenty four)

[0177] in, is the effective traction stress, is the total fracture energy of the interface, is the interface elastic energy;

[0178] After reaching a certain energy, the interface begins to completely debond and delaminate:

[0179] (25)

[0180] In the formula, is the fracture energy, is the critical fracture energy, , ; is the viscosity parameter; , , is the work done by the stress in the normal direction and the two tangential directions under the corresponding displacement.

[0181] As for out-of-plane impact, the damage model is as follows:

[0182] When composite materials are subjected to low-speed impact, their damage forms are mainly fiber damage, matrix damage and delamination. The specific damage criteria are as follows:

[0183] Fiber tensile fracture ( ): (26);

[0184] Fiber compression failure ( ): (27);

[0185] Matrix cracking ( ): (28);

[0186] Matrix extrusion ( ):

[0187] (29);

[0188] In the formula (i,j=1,2,3) are stress components, 1 is the fiber direction, 2 is the matrix direction, 3 is the out-of-plane normal direction, s is the 12 shear direction, X 13 , X 23 They are the shear strength in the 13 and 23 directions respectively. As long as the stress state of the unit satisfies the equation, the corresponding failure mode will appear, and then the material will begin to be damaged.

[0189] Fiber tensile fracture; (30)

[0190] Fiber compression failure: (31)

[0191] Matrix tensile cracking: (32)

[0192] Matrix compression failure: (33)

[0193] In the formula, and are the reduced variables of the fiber direction modulus and the matrix direction modulus, respectively. , is the reduction variable of shear modulus in 12 and 23 directions; is the modulus in the 23 direction before reduction; , , , They are fiber tensile reduction variable, fiber compression reduction variable, matrix tensile reduction variable, matrix compression reduction variable, and their calculation formulas are as follows:

[0194] (34)

[0195] In the formula t T It is the total time from the beginning of impact to the time when the punch kinetic energy reaches zero. The out-of-plane impact delamination model is consistent with the in-plane impact delamination model.

[0196] The results of the above model are compared with existing experiments, such as Figure 4 , 5 As shown. Figure 4As can be seen from the graph, the experimental results have an overall good correlation with the results of the present model. All curves go through two stages: the impact stage and the rebound stage. In the impact stage, both the experimental and simulated load-time responses show a rapid upward trend. Under the impact load, matrix cracks and fiber fractures appear at the impact point. At the same time, delamination occurs from the impact side to the back side. The impact energy is mainly dissipated by the above failure modes and stored in the elastic deformation of the laminate. As the punch penetrates deeper into the composite laminate, the kinetic energy of the punch gradually dissipates to 0, which can be seen at the impact time of about 5.1 ms in both curves. At the same time, the experimental data and simulation results reach almost the same maximum impact load of about 5.0 kN, which shows the accuracy of the finite element. In the rebound stage, the energy stored in the laminate is initially gradually transferred to the impactor. The response in the impact stage is almost symmetrical with that in the rebound stage, with the axis of symmetry at t=5.1 ms. The only difference in the stiffness of the laminate was found in the two stages of the experiment and simulation, which is mainly due to the simulated premature delamination damage. Comparison of the load-displacement curves of the out-of-plane impact experiment and simulation Figure 5 As shown. It can be divided into three typical stages, including (I) the initial impact stage, (II) the progressive failure stage and (III) the return stage, which can be observed in both the experimental curve and the simulation curve. In stage (I), before the initial damage, the load grows rapidly, which is considered to be elastic deformation. As the impactor deepens, the matrix and fiber begin to undergo compression damage. At the same time, the contact area between the impactor and the edge of the laminate also increases, resulting in a tortuous rise in the load. After that, delamination occurs near the impact area. The load continues to rise and reaches a maximum value until the impactor contacts the edge of the laminate layer. In stage (II), the delamination area increases with the increase in penetration depth. However, the number of layers in contact with the impactor has not increased, which also means that the laminate has reached the maximum load-bearing capacity of the impactor. Therefore, the load shows a downward trend with fluctuations. Of course, if the impact process continues, the damage within and between the laminar flows will maintain a certain balance with the penetration depth of the impactor, and the load will continue to be about 6000N. In stage (III), the kinetic energy of the impactor is completely converted into elastic potential energy and energy dissipated by various damage. The elastic potential energy stored in the laminate is initially transferred to the impactor. Within a small displacement range, the load quickly drops to 0, and the simulation results are in good agreement with the experimental results. The accuracy and effectiveness of the above model have been verified and can provide a reference for design.

[0197] Three spiral arrangement methods are proposed for the bionic spiral laminate reinforcement design: Figure 6 As shown. The helix angles are: Linear helical configuration (LRA): [7.5 / 15 / 22.5 / 30 / 37.5 / 45 / 52.5 / 60 / 67.5 / 75 / 82.5 / 90] s, Sine helical configuration (SRA): [0 / 12.81 / 25.36 / 37.39 / 48.66 / 58.94 / 68.02 / 75.71 / 81.87 / 86.35 / 89.08 / 90] s , Exponential chord helical configuration (ERA): [0.05 / 0.42 / 1.41 / 3.33 / 6.51 / 11.25 / 17.87 / 26.67 / 37.97 / 52.08 / 69.32 / 90] s The ply angle configurations used in the experiment are all reference configurations (Re): [-45 / 0 / 0 / 45 / 0 / -45 / 0 / 45 / 0 / 0] s ; [902 / -452 / 04 / 452 / 04 / 452 / 04 / -452 / 902] s , and then the designed structure is subjected to finite element calculation, and the results are as follows Figure 7-8 shown.

[0198] from Figure 7 It can be seen that in the initial stage of the impact, due to the lack of damage, the slopes of all the curves are similar. In the range of 0.8~1.9ms, the slope or stiffness of the bionic spiral laminate is similar to that of the reference laminate, and is steeper. With further loading of the impactor, subtle differences begin to appear in the three types of spiral composite laminates. The slope or stiffness of the bionic spiral laminate is significantly greater than that of Re, indicating that the spiral composite laminate has good resistance to low-speed impact. At 3.2ms, due to the initial matrix damage and delamination, the slopes of all curves began to decrease until they reached zero (reaching the peak load), and the peak time of the impact load (peak time) of Re was later than that of the bionic spiral laminate. In the initial stage of the impact, due to the slightly different helix angles of the layers initially involved in the impact, the impact responses of the bionic spiral laminates are very similar. From Figure 8 It can be seen that the bionic spiral laminate has greater stiffness than the reference laminate, and the loads of SRA and ERA decrease rapidly from 0.25 ms to around 0.6 ms, which may be due to the large initial stiffness. In summary, the bionic spiral laminate has better in-plane impact resistance and out-of-plane impact resistance than the conventional laminate.

[0199] Although the embodiments of the present invention have been disclosed as above, they are not limited to the applications listed in the specification and the embodiments. They can be fully applied to various fields suitable for the present invention. For those familiar with the art, additional modifications can be easily implemented. Therefore, without departing from the general concept defined by the claims and the scope of equivalents, the present invention is not limited to the specific details and the illustrations shown and described herein.

Claims

1. A method for designing aircraft composite material structures taking low-speed impact performance into consideration, characterized in that: The following steps are involved: Step 1: Establish a finite element model based on actual working conditions; Step 2: Assign material properties to the finite element model; Step 3: Mesh the finite element model; Step 4: Set the boundary conditions of the finite element model according to the actual working conditions; Step 5: Construct in-plane impact damage model and out-of-plane impact damage model; The out-of-plane impact damage model is as follows. When the composite material is subjected to low-speed impact, the specific failure criteria are as follows: Fiber tensile fracture, : ; Fiber compression failure, : ; Matrix cracking, : ; Matrix extrusion, : ； In the formula (i,j=1,2,3) is the effective stress tensor, 1 is the fiber direction, 2 is the matrix direction, 3 is the out-of-plane normal direction, 12 is the shear direction, , , , and They are respectively fiber tensile strength, fiber compressive strength, matrix tensile strength, matrix compressive strength, shear strength in 12 directions, X 13 , X 23 They are the shear strength in the 13 and 23 directions respectively. As long as the stress state of the unit satisfies the equation, the corresponding failure mode will appear, and then the material will begin to lose. Step 6: Get and integrate simulation results; Step 7: Compare and analyze all impact responses and failure models; Step 8: Refer to the impact-resistant structure, comprehensively consider the in-plane low-speed impact resistance performance and the out-of-plane low-speed impact resistance performance, and carry out structural reinforcement design; Step 9: Carry out steps 5 to 7 again for the designed new structure; Step 10: Analyze the strengthening mechanism of the structure with strengthening effect; Step 11: Based on the reinforcement mechanism analysis, comprehensively consider the in-plane low-speed impact resistance performance and the out-of-plane low-speed impact resistance performance, optimize the structural parameters, repeat steps 8 to 10 for optimization iteration, and finally obtain the optimal design.

2. The method for designing aircraft composite material structures considering low-speed impact performance according to claim 1, characterized in that: In the step eight, the impact-resistant structure includes a biological impact-resistant structure.

3. The method for designing aircraft composite material structures considering low-speed impact performance according to claim 1, characterized in that: The in-plane impact damage model is as follows. The failure modes of the fiber and matrix are mainly divided into five failure modes, namely, fiber tensile fracture, fiber compression buckling and kinking, matrix tensile cracking, matrix compression crushing, and in-plane shear failure. The corresponding damage criteria are as follows: Fiber tensile fracture, : ; Fiber compression buckling and kinking, : ; Matrix tensile cracking, : ; Matrix compression and crushing, : ; In-plane shear failure, : ; In the formula, F ft is the longitudinal tensile failure factor, F fc is the longitudinal compression failure factor, F mt is the transverse tensile failure factor, F mc is the lateral compression failure factor, F s is the in-plane shear failure factor.

4. The method for designing aircraft composite material structures considering low-speed impact performance according to claim 1, characterized in that: The composite material is a laminate structure formed by stacking and bonding multiple single-layer boards, and adjacent single-layer boards are bonded via a bonding interface, and the bonding interface is made by the following steps: S1. Laying the fiber cloth in layers according to a preset laying angle, applying a layer of a mixture of resin and curing agent after each layer is laid; S2. Apply release agent to the upper and lower mold plates, and place the laid fiber cloth between the molds after the release agent is dry; S3, hot pressing and curing the fiber cloth in the mold; S4. Open the mold.

5. The method for designing aircraft composite material structures considering low-speed impact performance according to claim 1, characterized in that: The velocity range of the low-speed impact is 1-10 m / s, and the specific value depends on the stiffness and component properties of the material being impacted, as well as the mass and stiffness of the punch.

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