A strength analysis method, equipment, medium and product for aviation composite laminated plate and shell structure

Abstract

The present application discloses a strength analysis method, equipment, medium and product for aviation composite laminated plate and shell structures, and relates to the field of data processing. The method comprises: constructing an overall structural plate and shell model based on the structural parameters of aviation composite laminated plate and shell; applying service loads to the overall structural plate and shell model, and performing finite element simulation analysis to obtain a generalized stress state; obtaining a proxy model; inputting the generalized stress state into the proxy model to obtain a strength margin; and performing structural strength verification based on the strength margin. The present application can directly verify the strength of composite shell structures, greatly reducing the calculation cost.

Description

Technical Field

[0001] The present application relates to the field of data processing, and in particular to a strength analysis method, equipment, medium and product for aviation composite laminated plate and shell structures. Background Art

[0002] Composite structures are increasingly used in advanced aircraft structures, so in the field of aircraft design, high requirements are placed on reliable numerical tools for predicting the mechanical properties of composite structures. Composite laminates are the most typical composite structures used on aircraft. In order to reduce computational costs, shell elements are usually used instead of solid elements for modeling. Classical plate and shell theory was developed by Kirchhoff and Mindlin in the last century, and these theories were later extended to various structures. In recent years, because advanced engineering thin-walled structures composed of complex internal structures are difficult to handle with classical theory, a structural homogenization method has been developed to simplify complex three-dimensional structures into two-dimensional shell models. This method can establish an equivalent shell element model of arbitrary periodic plate and shell structures based on unit cell analysis, thereby effectively modeling the structural response.

[0003] Although structural homogenization methods have been widely studied, complex strength models for composite shells are still lacking. Currently widely used composite strength models, such as the Tsai-Wu criterion, the Hashin criterion, the Pinho criterion, etc., are all defined based on the Cauchy stress at the material level. However, the stress measure obtained by the laminate shell model is the cross-sectional internal force, and the strength model at the material level cannot be used directly. The strength model based on the cross-sectional force cannot be used for composite laminates. There is currently no mature composite laminate shell strength model, which brings difficulties to the strength prediction of aircraft composite structures.

[0004] Since there is no shell element-based laminate strength model, the only way is to calculate the Cauchy stress in each layer based on the shell section force, and then apply the stress-based strength model for strength verification. Under complex loads, the stress field in each layer of composite laminates is complex, and it is not easy to calculate the stress field using classical shell theory. In the structural homogenization method, the complex internal stress field can be restored through the dehomogenization process, but since the critical area cannot be predicted based on the section force, the stress field of all shell elements must be restored, which consumes too much computational cost in engineering practice. Summary of the invention

[0005] The purpose of this application is to provide a strength analysis method, equipment, medium and product for aviation composite laminate shell structures, which can directly perform strength verification of composite shell structures and greatly reduce calculation costs.

[0006] To achieve the above objectives, this application provides the following solutions:

[0007] In a first aspect, the present application provides a strength analysis method for an aviation composite laminate shell structure, comprising:

[0008] Based on the structural parameters of the aviation composite laminated plate and shell, an overall structural plate and shell model is constructed; the structural parameters of the aviation composite laminated plate and shell include: the number of layers of the aviation composite laminated plate and shell and the thickness of the aviation composite laminated plate and shell;

[0009] Applying service loads to the integral structural plate-shell model and performing finite element simulation analysis to obtain a generalized stress state; the generalized stress is the stress that causes the integral structural plate-shell model to fail;

[0010] Obtaining a proxy model; the proxy model is a fully connected neural network model trained using a simulation data set;

[0011] Inputting the generalized stress state into the proxy model to obtain a strength margin;

[0012] A structural strength check is performed based on the strength margin.

[0013] Optionally, the process of constructing the simulation data set includes:

[0014] Based on the structural parameters of aviation composite laminated plates and shells, a three-dimensional unit cell model of aviation composite laminated plates and shells is constructed;

[0015] Apply generalized stress to various unit sections of a three-dimensional unit cell model of aviation composite laminated plates and shells, perform unit cell multi-condition deformation simulation, and obtain a multi-dimensional generalized stress space;

[0016] Performing random sampling of generalized stress states in the multidimensional generalized stress space to obtain sampling results;

[0017] Determine the structural response under different generalized stress states in each sampling result based on the linear superposition of the field;

[0018] The structural failure state is judged based on the strength judgment criterion of Cauchy stress and the structural response under different generalized stress states, and the judgment result of the structural failure state is obtained;

[0019] The interlayer interface strength judgment criterion is applied to obtain the delamination failure state judgment result; the interlayer interface strength judgment criterion is used to judge the delamination failure state of the structural response under different generalized stress states;

[0020] Determine the strength margin of the unit cell failure state based on the structural failure state judgment result and the delamination failure state judgment result;

[0021] Different generalized stress states and corresponding strength margins are used as data sample pairs to obtain the simulation data set.

[0022] Optionally, based on the structural parameters of the aviation composite laminated plate and shell, a three-dimensional unit cell model of the aviation composite laminated plate and shell is constructed, including:

[0023] The geometric model of the unit cell is constructed based on the structural parameters of the aviation composite laminated plate and shell, and the finite element mesh is divided to obtain the finite element model;

[0024] In the finite element model, the stiffness properties of each layer of the unit cell and the contact properties between the layers are defined, the degrees of freedom of the virtual nodes representing the generalized stresses are set, and periodic boundary conditions containing disturbance field corrections are applied between the corresponding virtual nodes in the in-plane direction of the unit cell to obtain an initial three-dimensional unit cell model;

[0025] The rigid body displacement of the initial three-dimensional unit cell model is controlled by applying distributed coupling constraints to obtain the three-dimensional unit cell model of the aviation composite laminate shell.

[0026] Optionally, generalized stress is applied to various unit sections of a three-dimensional unit cell model of an aviation composite laminate shell, and a unit cell multi-condition deformation simulation is performed to obtain a multi-dimensional generalized stress space, including:

[0027] Apply unit force to each virtual node in turn, and simulate and determine the stress and interface force response of the aviation composite laminate shell unit cell under multiple unit generalized stress action conditions; the multiple unit generalized stress action conditions include the generalized stress action as tension, the generalized stress action as bending, the generalized stress action as torsion, and the generalized stress action as in-plane shear;

[0028] Simulate and determine the displacement disturbance field and surface force field of aerospace composite laminated plate and shell unit cells under multiple unit generalized bending moments;

[0029] Applying the displacement disturbance field and the surface force field on the periodic boundary surface, and simulating and determining the stress and interface force response of the aviation composite laminated plate and shell unit cell under multiple unit generalized out-of-plane shear force conditions;

[0030] A multidimensional generalized stress space is generated based on the stress and interface force response of aerospace composite laminated plate and shell unit cells under multiple unit generalized stress conditions, and the stress and interface force response of aerospace composite laminated plate and shell unit cells under multiple unit generalized out-of-plane shear force conditions.

[0031] Optionally, a geometric model of a unit cell is constructed based on the structural parameters of the aviation composite laminate shell, including:

[0032] Based on the structural parameters of the aviation composite laminated plate and shell, a geometric model of the unit cell is constructed with the center of the unit cell as the coordinate origin and the thickness direction as a coordinate axis in a manner that the end planes are parallel to each other.

[0033] Optionally, a distributed coupling constraint is applied to control the rigid body displacement of the initial three-dimensional unit cell model to obtain the three-dimensional unit cell model of the aviation composite laminate shell, including:

[0034] A reference point is set in the initial three-dimensional unit cell model, and a distributed coupling constraint is applied between the reference point and a line passing through the coordinate origin in the unit cell and perpendicular to the neutral surface of the aviation composite laminate shell, while the translation and rotation of the reference point in three directions are set to 0.

[0035] Optionally, virtual nodes with strength margins within a set range are used as valid information, and the failure value of a structural unit cell under a single load multiplied by a set amplification factor is used as a sampling area boundary, and random sampling of generalized stress is performed in the multidimensional generalized stress space to obtain a sampling result.

[0036] In a second aspect, the present application provides a computer device, including: a memory, a processor, and a computer program stored in the memory and executable on the processor, wherein the processor executes the computer program to implement the above-mentioned aviation composite laminate shell structure strength analysis method.

[0037] In a third aspect, the present application provides a computer-readable storage medium having a computer program stored thereon, which, when executed by a processor, implements the above-mentioned strength analysis method for aviation composite laminated plate and shell structures.

[0038] In a fourth aspect, the present application provides a computer program product, including a computer program, which, when executed by a processor, implements the above-mentioned aviation composite laminate shell structure strength analysis method.

[0039] According to the specific embodiments provided in this application, this application discloses the following technical effects:

[0040] The present application provides a strength analysis method, equipment, medium and product for aviation composite laminated plate and shell structures. By adopting a proxy model, failure judgment of the structure can be made directly based on the cross-sectional stress of the aviation composite laminated plate and shell unit without restoring the Cauchy stress inside the laminate. This can greatly reduce the calculation cost, effectively improve the analysis efficiency while ensuring the analysis accuracy, and thus help realize real-time rapid failure prediction and strength verification of large-scale aviation composite plate and shell structures under complex external load conditions. BRIEF DESCRIPTION OF THE DRAWINGS

[0041] In order to more clearly illustrate the embodiments of the present application or the technical solutions in the prior art, the drawings required for use in the embodiments will be briefly introduced below. Obviously, the drawings described below are only some embodiments of the present application. For ordinary technicians in this field, other drawings can be obtained based on these drawings without paying creative work.

[0042] Figure 1 This is an application environment diagram of a strength analysis method for aviation composite laminated plate shell structure in one embodiment of the present application;

[0043] Figure 2 A schematic flow chart of a strength analysis method for an aviation composite laminate shell structure provided in one embodiment of the present application;

[0044] Figure 3 A schematic diagram of the overall implementation process of a strength analysis method for an aviation composite laminate shell structure provided in one embodiment of the present application;

[0045] Figure 4 A schematic diagram of a three-dimensional unit cell model of an aviation composite laminate shell is provided for one embodiment of the present application;

[0046] Figure 5 A schematic diagram of simulation results of structural response under various unit generalized stresses provided in an embodiment of the present application;

[0047] Figure 6 A schematic diagram of a data processing flow of a strength judgment neural network model provided in an embodiment of the present application;

[0048] Figure 7 A schematic diagram of a strength judgment neural network model provided by another embodiment of the present application;

[0049] Figure 8 A schematic diagram of a composite aviation stiffened wall panel structure data-driven strength model predicting ultimate load and displacement response provided by another embodiment of the present application;

[0050] Fig. 9 A schematic diagram of a composite aviation stiffened wall panel structure data-driven strength model predicting a failure location and a generalized force response provided by another embodiment of the present application;

[0051] Fig.10 A schematic diagram of the reference prediction limit load and displacement response of the Composite Layup module of ABAQUS provided in another embodiment of the present application;

[0052] Fig.11 A schematic diagram of the Composite Layup module of ABAQUS for reference prediction of the failure position and generalized force response provided in another embodiment of the present application;

[0053] Fig.12 A schematic diagram of the structure of a computer device provided in one embodiment of the present application. DETAILED DESCRIPTION

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

[0055] This application aims to establish a laminate strength model based on the horizontal section force of the structure, so that the strength check of the composite shell structure can be carried out directly, greatly reducing the calculation cost.

[0056] In order to make the above-mentioned objects, features and advantages of the present application more obvious and easy to understand, the present application is further described in detail below in conjunction with the accompanying drawings and specific implementation methods.

[0057] The strength analysis method of aviation composite laminate shell structure provided in the embodiment of the present application can be applied to Figure 1 In the application environment shown. Among them, the terminal 102 communicates with the server 104 through the network. The data storage system can store the data that the server 104 needs to process. The data storage system can be set up separately, integrated on the server 104, or placed on the cloud or other servers. The terminal 102 can send the structural parameters of the aviation composite laminated plate and shell to the server 104. After the server 104 receives the structural parameters of the aviation composite laminated plate and shell, the server 104 constructs an overall structural plate and shell model for the structural parameters of the aviation composite laminated plate and shell. Apply service loads to the overall structural plate and shell model, and perform finite element simulation analysis to obtain a generalized stress state. Obtain a proxy model. Input the generalized stress state into the proxy model to obtain a strength margin. Perform a structural strength check based on the strength margin. The server 104 can feedback the structural strength check results to the terminal 102. In addition, in some embodiments, the aviation composite laminate shell structure strength analysis method can also be implemented independently by the server 104 or the terminal 102. For example, the terminal 102 can directly perform an aviation composite laminate shell structure strength analysis on the structural parameters of the aviation composite laminate shell, or the server 104 can obtain the structural parameters of the aviation composite laminate shell from the data storage system and perform an aviation composite laminate shell structure strength analysis on the structural parameters of the aviation composite laminate shell.

[0058] The terminal 102 may be, but is not limited to, various desktop computers, laptop computers, smart phones, tablet computers, IoT devices, and portable wearable devices. The IoT devices may be smart speakers, smart TVs, smart air conditioners, smart vehicle-mounted devices, etc. The portable wearable devices may be smart watches, smart bracelets, head-mounted devices, etc. The server 104 may be implemented as an independent server or a server cluster consisting of multiple servers, or may be a cloud server.

[0059] In an exemplary embodiment, Figure 2 As shown, a strength analysis method for aviation composite laminated plate shell structure is provided. The method is executed by a computer device, and can be executed by a computer device such as a terminal or a server alone, or by a terminal and a server together. In the embodiment of the present application, the method is applied to Figure 1 The server 104 in the example is used as an example to illustrate the method, which includes the following steps 200 to 204. Among them:

[0060] Step 200: constructing an overall structural plate shell model based on the structural parameters of the aviation composite laminated plate shell. The structural parameters of the aviation composite laminated plate shell include: the number of layers of the aviation composite laminated plate shell and the thickness of the aviation composite laminated plate shell.

[0061] Step 201: Apply service loads to the overall structural plate and shell model, and perform finite element simulation analysis to obtain a generalized stress state. The generalized stress is the stress that causes the overall structural plate and shell model to fail.

[0062] Step 202: Obtain a proxy model. The proxy model is a fully connected neural network model trained using a simulation data set.

[0063] Step 203: Input the generalized stress state into the proxy model to obtain the strength margin.

[0064] Step 204: Perform structural strength verification based on the strength margin.

[0065] Implementing the above steps 200 to 204 can greatly reduce the computational cost, effectively improve the analysis efficiency while ensuring the analysis accuracy, and thus can help realize the real-time rapid failure prediction and strength verification of large-scale composite plate and shell structures of aviation under complex external load conditions. In addition, since composite structures with arbitrary internal topological configurations or material compositions can be analyzed by the plate and shell homogenization method, it is convenient and universal in acquiring data. This application can well solve the problem that the traditional phenomenological strength model constructed by explicit mathematical expressions is often difficult to apply to different composite structures.

[0066] In another exemplary embodiment of the present application, in order to directly judge the failure of the structure based on the cross-sectional internal force of the shell unit without restoring the Cauchy stress inside the laminate, in this embodiment, a strength verification proxy model suitable for real-time, fast and efficient simulation is established based on the data set (i.e., simulation data set) obtained by the unit cell analysis. Based on this, the construction process of the above simulation data set includes:

[0067] Step 1. Based on the structural parameters of the aviation composite laminated plate and shell, a three-dimensional unit cell model of the aviation composite laminated plate and shell is constructed.

[0068] Step 2. Apply generalized stresses to various unit cross sections of the three-dimensional unit cell model of aviation composite laminated plates and shells, perform unit cell multi-condition deformation simulation, and obtain a multi-dimensional generalized stress space.

[0069] Step 3. Perform random sampling of generalized stress states in the multidimensional generalized stress space to obtain sampling results.

[0070] Step 4. Based on the linear superposition of the field, determine the structural response under different generalized stress states in each sampling result. For example, considering that the mechanical behavior of the structure before failure is approximately linear, based on the response results of the unit cell under each unit generalized stress, the structural response under complex loads is solved by linear superposition.

[0071] Specifically: When a unit load is applied to the virtual node degrees of freedom, the stress field of the unit cell of the structure obtained by finite element simulation is: , the generalized stress state at the sampling point is , then the unit cell stress field corresponding to the generalized load state can be calculated , expressed as:

[0072] .

[0073] in, is the generalized stress vector The 8 generalized stress components of .

[0074] In this process, methods for calculating the Cauchy stress field of a unit cell include but are not limited to the linear superposition method. The generalized unit stress can also be applied to the unit cell in the form of boundary conditions to re-solve the finite element calculation.

[0075] Step 5. The structural failure state is judged based on the strength judgment criterion of the Cauchy stress and the structural response under different generalized stress states to obtain the structural failure state judgment result. Among them, the strength judgment criterion of the material scale includes but is not limited to the Tsai-Wu criterion, the Hashin criterion, the Pinho criterion, etc. The strength judgment criterion is mainly used to judge the damage initiation of aviation composite laminated plates and shells, and to judge whether a single-layer aviation composite laminated plate and shell has failed based on the Cauchy stress, strain or strain energy.

[0076] Step 6. Apply the interlayer interface strength judgment criteria, that is, the structural response under different generalized stress states to judge the delamination failure state and obtain the delamination failure state judgment result. Among them, this step is mainly based on the interlayer interface stress, separation displacement, and whether interlayer cracking occurs. The interlayer interface strength judgment criteria include but are not limited to the maximum stress / separation criteria and the secondary stress / separation criteria. As long as the judgment criteria for the failure of a single layer plate based on the unit cell detail stress, strain or strain energy response are given, they can be applied.

[0077] Step 7. Determine the strength margin of the failure state of the unit cell based on the structural failure state judgment results and the layered failure state judgment results. For example, for each sampling point, take its generalized stress state as input, obtain the corresponding detailed stress and deformation state, and then apply the criteria in the above steps 5 and 6 to judge the failure state of each virtual node of the unit cell one by one, and take the failure state corresponding to the most dangerous node as the output. Among them, the stress state of each node in the unit cell is different, and its corresponding strength margin is different. The most dangerous node used in this step is the point with the largest strength margin among all the nodes of the unit cell.

[0078] Step 8. Take different generalized stress states and corresponding strength margins as data sample pairs to obtain simulation data sets.

[0079] In another exemplary embodiment of the present application, the process of constructing a three-dimensional unit cell model of an aviation composite laminated plate shell based on the structural parameters of the aviation composite laminated plate shell in step 1 includes:

[0080] Step 1-1. Construct a unit cell geometric model based on the structural parameters of the aviation composite laminated plate and shell, and divide the finite element mesh to obtain a finite element model. The structural parameters of the aviation composite laminated plate and shell include the number of layers of the aviation composite laminated plate and shell and the thickness of the aviation composite laminated plate and shell, but are not limited thereto. In actual application, the structural parameters can be selected according to actual needs.

[0081] In practical applications, based on the structural parameters of aviation composite laminated plates and shells, the geometric model of the unit cell is constructed with the center of the unit cell as the coordinate origin, the thickness direction as a coordinate axis, and the end planes parallel to each other. The resulting geometric model of the unit cell is a cubic model divided by the number of layers and thickness. Representing three coordinate directions to construct a cube model, the periodic directions are , the thickness direction of the ply is , the origin of the coordinates is located at the center of the unit cell, and the end planes are parallel to each other. The mesh is divided symmetrically and each layer is independent of each other and does not share nodes.

[0082] Step 1-2. In the finite element model, define the stiffness properties of each layer of the unit cell and the contact properties between layers, set the degrees of freedom of the virtual nodes representing each generalized stress, and apply periodic boundary conditions containing disturbance field corrections between the corresponding virtual nodes in the plane direction of the unit cell to obtain the initial three-dimensional unit cell model. Among them, the defined stiffness properties of each layer of the unit cell can be the material stiffness properties of each layer of the single-layer plate, which is generally transversely isotropic, and the local material coordinate system of the finite element model of each layer is defined according to the ply direction and ply sequence. The stiffness properties of the bonding between the layers are defined by the cohesive contact method, which can include a tensile stiffness in the out-of-plane direction and a shear stiffness in two in-plane directions.

[0083] In practical applications, equation constraints can be used to impose periodic boundary conditions containing disturbance field corrections between corresponding nodes in the unit cell surface. The equation constraints constrain the displacement difference relationship in each direction of the corresponding nodes on the positive and negative boundary surfaces in the unit cell surface. For example, The finite element mesh nodes are The coordinate value of the direction. The mesh nodes on the positive end face of the direction The negative end face has the same The nodes of the coordinates correspond to, accordingly, The mesh nodes on the positive end face of the direction The negative end face has the same The nodes of the coordinates correspond to each other. Based on this, the equation constraints imposed are expressed as:

[0084] .

[0085] .

[0086] .

[0087] .

[0088] .

[0089] .

[0090] in, Respectively The points on the positive and negative end faces are Displacement in direction. The unit cell is subjected to unit bending moment When in effect, The nodes on the positive end face are The disturbance displacement field in the direction, They are The magnitude of the load applied to the corresponding node degree of freedom, For the The displacement of the virtual node degrees of freedom, Respectively represent the unit cell in Length in direction.

[0091] Set up 8 virtual node degrees of freedom corresponding to 8 generalized strains of plates and shells. Displacement of virtual node degrees of freedom The values ​​are:

[0092] .

[0093] .

[0094] .

[0095] .

[0096] .

[0097] .

[0098] .

[0099] .

[0100] in, for Directional stretching deformation, for Directional stretching deformation, is the deformation due to in-plane shear, for Direction bending deformation, for Direction bending deformation, For torsional deformation, for Out-of-plane shear deformation, for Out-of-plane shear deformation.

[0101] Step 1-3. Apply distributed coupling constraints to control the rigid body displacement of the initial three-dimensional unit cell model to obtain a three-dimensional unit cell model of the aviation composite laminated plate and shell. For example, a reference point is set in the initial three-dimensional unit cell model, and a distributed coupling constraint is applied between the reference point and a line in the unit cell that passes through the coordinate origin and is perpendicular to the neutral surface of the aviation composite laminated plate and shell, and the translation and rotation of the reference point in three directions are set to 0.

[0102] In another exemplary embodiment of the present application, in step 2, the structural unit cell response under eight unit cross-section internal force single action conditions can be simulated and calculated, laying the foundation for rapid simulation of structural stress under complex load conditions. Based on this, in step 2, generalized stress is applied to multiple unit cross sections of the three-dimensional unit cell model of the aviation composite laminate shell, and the unit cell multi-condition deformation simulation is performed to obtain the multi-dimensional generalized stress space. The process may include:

[0103] Step 2-1. Apply unit force to each virtual node in turn, and simulate and determine the stress and interface force response of the aviation composite laminate shell unit cell under multiple unit generalized stress conditions. For example, apply unit force to each virtual node in turn, and simulate and calculate the stress and interface force response of the aviation composite laminate shell unit cell under 6 unit generalized stress conditions including tension, bending, torsion, and in-plane shear. The implementation method of applying unit force is: A unit concentrated force is applied to the corresponding virtual node degrees of freedom.

[0104] Step 2-2. Simulate and determine the displacement disturbance field and surface force field of the aviation composite laminated plate and shell unit cell under multiple unit generalized bending moments. For example, the displacement disturbance field and surface force field of the plate and shell unit cell under 2 unit generalized bending moments are simulated and calculated as follows:

[0105] exist Apply a unit concentrated force on the corresponding virtual node degree of freedom, and let . and in Apply a unit concentrated force on the corresponding virtual node degree of freedom, and let The displacement disturbance field and surface force field of the unit cell boundary surface are obtained through finite element calculation.

[0106] Step 2-3. Apply displacement disturbance field and surface force field on the periodic boundary surface, and simulate and determine the stress and interface force response of the aviation composite laminated plate and shell unit cell under multiple unit generalized surface shear force conditions. For example, apply surface force disturbance field on the periodic boundary surface, and simulate and calculate the stress and interface force response of the plate and shell unit cell under 2 unit generalized surface shear force conditions. Among them:

[0107] (1) The implementation method of applying the surface force disturbance field on the periodic boundary surface is:

[0108] According to the simulation calculation results of step 2-2, the surface force field is applied to the boundary surface of the unit cell by applying node concentrated force.

[0109] The simulation calculation method of the stress and interface force response of the plate and shell unit cell under the working condition of two unit generalized external shear forces is as follows:

[0110] exist A unit concentrated force is applied to the corresponding virtual node degrees of freedom. This response is the response of the unit cell under two unit generalized stresses, recorded as . That is Out-of-plane shear force in the direction.

[0111] Step 2-4. Generate a multidimensional generalized stress space based on the stress and interface force response of the aviation composite laminated plate and shell unit cell under multiple unit generalized stress conditions, and the stress and interface force response of the aviation composite laminated plate and shell unit cell under multiple unit generalized out-of-plane shear force conditions. Among them, the stress and interface force response of the aviation composite laminated plate and shell unit cell under multiple unit generalized stress conditions, and the stress and interface force response of the aviation composite laminated plate and shell unit cell under multiple unit generalized out-of-plane shear force conditions, the unit cell stress field is a three-dimensional Cauchy stress, denoted as , are the Cauchy stress components, respectively The normal stress in the direction and The shear stress in the plane. The interface force response is , represent the interface forces in the normal direction and the two shear directions respectively.

[0112] In another exemplary embodiment of the present application, in order to improve the validity of the data and the efficiency of model training, the scope of the sampling space should be limited to a certain extent. Based on this, the above step 3 can use the virtual nodes with strength margins within the set range as valid information, and the failure value of the structural unit cell under a single load multiplied by the set magnification factor as the sampling area boundary, and perform random sampling of generalized stress in the multidimensional generalized stress space to obtain the sampling result. For example:

[0113] Considering the data points close to the strength envelope in the 8-dimensional generalized stress space as valid information, the failure value of the structural unit cell under a single load multiplied by a certain magnification factor is used as the boundary of the sampling area. The strength envelope here is actually a surface in the 8-dimensional generalized stress space, that is, the surface formed by the strength margin of 1. Based on this, considering the data points close to the strength envelope in the 8-dimensional generalized stress space is actually the data points near the output strength margin of 1.

[0114] In another exemplary embodiment of the present application, the construction process of the proxy model obtained above in the present application can be described as:

[0115] Based on the simulation data set, a fully connected neural network is constructed and trained. The input of the fully connected neural network is the generalized stress state, and the output is defined as the strength margin , when its value is greater than or equal to 1, it indicates structural damage. Among them, the strength margin The strength judgment criterion based on step 5 above is solved. The generalized stress state is expressed as , for Directional pull, for Directional pull, is the in-plane shear force, for Direction bending moment, for Direction bending moment, For torque.

[0116] Assume that the neural network contains hidden layer, The width of the layer is . Use and represents the first Neurons ( ), the nonlinear activation function is the linear rectification (Rectified Linear Unit, ReLU) function, and the neural network satisfies the following formula:

[0117] .

[0118] in, Indicates that the input is Time Tier The value after the nonlinear activation of neurons, Indicates that the input is Time Tier The value after the nonlinear activation of neurons, and They represent the weight and bias parameters trained based on the mean squared error (MSE) loss criterion, respectively:

[0119] .

[0120] in, are vectors consisting of all weight and bias parameters respectively, is the training data, Represents the neural network. Indicates the size of the data.

[0121] In this embodiment, the nonlinear activation function includes but is not limited to the linear rectification function, and any nonlinear activation equation obtained in the network is applicable. The loss criteria used include but are not limited to the mean square error loss criteria, and any loss criteria for parameter optimization based on model prediction and data deviation are applicable.

[0122] In another exemplary embodiment of the present application, a laminate made of ZT7H / 5429 material is taken as an example to illustrate the implementation process of the strength analysis method of the aviation composite laminate shell structure provided by the present application. Figure 3 As shown, including:

[0123] Step 1. Three-dimensional unit cell modeling of composite materials.

[0124] First, according to the thickness and number of layers of the laminate, the unit cell is modeled by finite element and the finite element mesh is divided. The end face mesh nodes in the two periodic directions of the unit cell model are paired. Figure 4 As shown in part (a) of the diagram, a coordinate system is constructed with the center point of the unit cell as the origin. The mesh nodes on the positive end face of the direction The negative end face has the same The node pairing of the coordinates, accordingly, The mesh nodes on the positive end face of the direction The negative end face has the same Node pairing of coordinates.

[0125] The unit cell model is established for a laminate made of ZT7H / 5429 material. The ply sequence is [45 / -45 / 0 / -45 / 45 / 0 / -45 / 45 / 90 / 45 / -45 / 0]s. The thickness of each layer is about 0.125mm and the total thickness is about 3mm. The ply sequence is as follows: Figure 4 As shown in part (b).

[0126] Secondly, define the stiffness properties of each layer of the unit cell and the contact properties between layers. The specific parameters are shown in Table 1.

[0127] Table 1 Material rigidity and strength performance parameters

[0128]

[0129] Again, set 8 virtual node degrees of freedom and define the The displacement magnitude of the virtual node degree of freedom is , whose numerical values ​​correspond to the eight generalized strains of plates and shells: , representing Directional stretching, in-plane shear, Directional bending, twisting, and Out-of-plane shear direction.

[0130] Finally, an additional reference point is set for rigid body displacement constraint, and a distributed coupling constraint is applied between the reference point and a line in the plate and shell unit cell that passes through the coordinate origin and is perpendicular to the neutral plane of the plate and shell. At the same time, the translation and rotation of the reference point in the three directions are set to 0.

[0131] Step 2. Single cell multi-condition deformation simulation.

[0132] The structural unit cell response under eight unit section internal force conditions is simulated. Unit force is applied to each reference node in turn, and the stress and interface force response of the unit cell under six unit generalized stress conditions including tension, bending, torsion, and in-plane shear are simulated. The displacement disturbance field and surface force field of the plate and shell unit cell under two unit generalized bending moments are simulated. The surface force disturbance field is applied on the periodic boundary surface, and the stress and interface force response of the plate and shell unit cell under two unit generalized out-of-plane shear conditions are simulated. The simulation results are as follows Figure 5 shown. Figure 5 Part (a) indicates the unit Unit cell Directional normal stress response; Figure 5 Part (b) indicates the unit Unit cell Directional normal stress response; Figure 5 Part (c) indicates the unit Unit cell Directional normal stress response; Figure 5 Part (d) indicates the unit Unit cell Directional normal stress response; Figure 5 The (e) part indicates the unit Unit cell Directional normal stress response; Figure 5 The (f) part indicates the unit Unit cell Directional normal stress response; Figure 5 The (g) part indicates the unit Unit cell Directional normal stress response; Figure 5 The (h) part indicates the unit The interface between the monolayers Directional separation displacement; Figure 5 The (i) part indicates the unit Unit cell Directional normal stress response; Figure 5 The (j) part indicates the unit The interface between the unit cells Directional separation displacement.

[0133] Step 3. Construction of failure judgment dataset.

[0134] The process of this step is as follows Figure 6 As shown in Figure 1, random sampling is performed in the 8-dimensional generalized stress space. Based on the linear superposition of the field, the structural response under different generalized stress states in each sample is calculated.

[0135] Five failure modes of laminates are considered here, including fiber fracture under tension. , fiber buckling and kinking under compression , Matrix failure under tension and compression and inter-layer delamination failure The Hashin criterion is used to judge the damage of single-layer plate materials, and the secondary stress model is used to judge the delamination failure state. The above five quantities used to characterize the failure state are:

[0136] .

[0137] In the formula, represents the tensile strength of the fiber, represents the shear strength, It represents the fiber compression strength, represents the matrix tensile strength, represents the matrix compressive strength, Indicates the interlaminar tensile strength.

[0138] Then the strength margin of the entire unit cell failure state is It is expressed as:

[0139] .

[0140] That is, a simulation data set can be obtained with the generalized stress state as input and the strength margin of the unit cell failure state as output. Based on this method, 40,000 data points can be generated in about 10 minutes.

[0141] Step 4. Proxy model construction.

[0142] Construct a fully connected neural network, such as Figure 7As shown. The model is trained based on the simulation data set. The simulation data is divided into 36,000 training data and 4,000 test data. The neural network is trained using PyTorch, and a neural network with 4 hidden layers of 64-132-64-8 units is obtained. The MSE value on the training validation set reaches 1.073, and the MSE value on the test set reaches 1.291. The model is used to ) and not expired ( ) has a prediction accuracy of over 99.8%.

[0143] Step 5. Plate and shell structure response simulation and strength verification.

[0144] The plate and shell unit in the finite element simulation software is used to establish the plate and shell model of the original three-dimensional overall structure, and the macro load under the actual service condition to be analyzed is applied to perform finite element simulation to obtain the overall generalized stress field of the structure. The generalized stress response is read, and the failure of each plate and shell unit is judged through the trained proxy model to perform structural strength verification.

[0145] In another exemplary embodiment of the present application, the analysis object is a section of composite aviation stiffened wall panel, and the material used is a ZT7H / 5429 material laminate with a ply pattern of [45 / -45 / 0 / -45 / 45 / 0 / -45 / 45 / 90 / 45 / -45 / 0]s. The structural macro model is as follows Figure 8 and Fig. 9 As shown in Figure 1, it consists of 14796 shell elements. One end of the wall panel is fixed, and a uniform distributed load is applied to the bottom. The load size is continuously increased, and the generalized force response of the overall structural plate-shell model is calculated and evaluated using the established proxy model. When the distributed load reaches 0.036MPa, the unit fails. Under this load condition, the wall panel deformation, the location of the damaged unit and its generalized force state are shown in Figure 1. Fig.10 and Fig.11 shown.

[0146] To verify the correctness, the Composite Layup module of ABAQUS was used to establish the overall structural plate-shell model. The strength of the wall panel was calculated. The predicted load when the first unit failed was 0.035MPa. The displacement field, the position of the failed unit, and the stress state of the section are shown in the figure below. Fig. 9 shown.

[0147] It can be seen that the error between the failure load predicted by the proxy model and the failure load predicted by the ABAQUS Layup module is 2.78%, and the predicted failure position is the same. The comparison shows that the accuracy of strength assessment using the proxy model is comparable to that of the Composite Layup module of ABAQUS. In addition, through the constructed proxy model, the damage assessment can be performed directly based on the cross-sectional force of the shell unit, omitting the process of calculating the Cauchy stress of each integral point of each layer in ABAQUS during structural analysis. Therefore, the proxy model constructed in this application has higher computational efficiency in application.

[0148] In an exemplary embodiment, a computer device is provided. The computer device may be a server or a terminal. The internal structure diagram thereof may be as follows: Fig.12 As shown. The computer device includes a processor, a memory, an input / output interface (Input / Output, referred to as I / O) and a communication interface. Among them, the processor, the memory and the input / output interface are connected through a system bus, and the communication interface is connected to the system bus through the input / output interface. Among them, the processor of the computer device is used to provide computing and control capabilities. The memory of the computer device includes a non-volatile storage medium and an internal memory. The non-volatile storage medium stores an operating system, a computer program and a database. The internal memory provides an environment for the operation of the operating system and the computer program in the non-volatile storage medium. The database of the computer device is used to store strength analysis data of aviation composite laminated plate and shell structures. The input / output interface of the computer device is used to exchange information between the processor and an external device. The communication interface of the computer device is used to communicate with an external terminal through a network connection. When the computer program is executed by the processor, a strength analysis method for aviation composite laminated plate and shell structures is implemented.

[0149] Those skilled in the art will understand that Fig.12 The structure shown in the figure is only a block diagram of a part of the structure related to the solution of the present application, and does not constitute a limitation on the computer device to which the solution of the present application is applied. The specific computer device may include more or fewer components than those shown in the figure, or combine certain components, or have a different arrangement of components. In an exemplary embodiment, a computer device is provided, including a memory and a processor, wherein a computer program is stored in the memory, and the processor implements the steps in the above-mentioned method embodiments when executing the computer program.

[0150] In an exemplary embodiment, a computer-readable storage medium is provided, storing a computer program, and when the computer program is executed by a processor, the steps in the above method embodiments are implemented.

[0151] In an exemplary embodiment, a computer program product is provided, including a computer program, and when the computer program is executed by a processor, the steps in the above method embodiments are implemented.

[0152] It should be noted that the user information (including but not limited to user device information, user personal information, etc.) and data (including but not limited to data used for analysis, stored data, displayed data, etc.) involved in this application are all information and data authorized by the user or fully authorized by all parties, and the collection, use and processing of relevant data must comply with relevant regulations.

[0153] Those skilled in the art can understand that all or part of the processes in the above-mentioned embodiment methods can be completed by instructing the relevant hardware through a computer program, and the computer program can be stored in a non-volatile computer-readable storage medium. When the computer program is executed, it can include the processes of the embodiments of the above-mentioned methods. Among them, any reference to the memory, database or other medium used in the embodiments provided in this application can include at least one of non-volatile and volatile memory. Non-volatile memory can include read-only memory (ROM), magnetic tape, floppy disk, flash memory, optical memory, high-density embedded non-volatile memory, resistive random access memory (ReRAM), magnetoresistive random access memory (MRAM), ferroelectric random access memory (FRAM), phase change memory (PCM), graphene memory, etc. Volatile memory can include random access memory (RAM) or external cache memory, etc. By way of illustration and not limitation, RAM may be in various forms, such as Static Random Access Memory (SRAM) or Dynamic Random Access Memory (DRAM).

[0154] The database involved in each embodiment provided in this application may include at least one of a relational database and a non-relational database. The non-relational database may include a distributed database based on blockchain, etc., but is not limited thereto. The processor involved in each embodiment provided in this application may be a general-purpose processor, a central processing unit, a graphics processor, a digital signal processor, a programmable logic device, a data processing logic device based on quantum computing, etc., but is not limited thereto.

[0155] The technical features of the above embodiments may be combined arbitrarily. To make the description concise, not all possible combinations of the technical features in the above embodiments are described. However, as long as there is no contradiction in the combination of these technical features, they should be considered to be within the scope of this specification.

[0156] This article uses specific examples to illustrate the principles and implementation methods of this application. The description of the above embodiments is only used to help understand the method and core ideas of this application. At the same time, for those skilled in the art, according to the ideas of this application, there will be changes in the specific implementation methods and application scope. In summary, the content of this specification should not be understood as limiting this application.

Claims

1. A strength analysis method for aviation composite laminate shell structure, characterized in that: The strength analysis method of the aviation composite material laminated plate and shell structure comprises: Based on the structural parameters of the aviation composite laminated plate and shell, an overall structural plate and shell model is constructed; the structural parameters of the aviation composite laminated plate and shell include: the number of layers of the aviation composite laminated plate and shell and the thickness of the aviation composite laminated plate and shell; Applying service loads to the integral structural plate-shell model and performing finite element simulation analysis to obtain a generalized stress state; the generalized stress is the stress that causes the integral structural plate-shell model to fail; Obtaining a proxy model; the proxy model is a fully connected neural network model trained using a simulation data set; Inputting the generalized stress state into the proxy model to obtain a strength margin; Performing structural strength verification based on the strength margin; The construction process of the simulation data set includes: Based on the structural parameters of aviation composite laminated plates and shells, a three-dimensional unit cell model of aviation composite laminated plates and shells is constructed; Apply generalized stress to various unit sections of a three-dimensional unit cell model of aviation composite laminated plates and shells, perform unit cell multi-condition deformation simulation, and obtain a multi-dimensional generalized stress space; Performing random sampling of generalized stress states in the multidimensional generalized stress space to obtain sampling results; Determine the structural response under different generalized stress states in each sampling result based on the linear superposition of the field; The structural failure state is judged based on the strength judgment criterion of Cauchy stress and the structural response under different generalized stress states, and the judgment result of the structural failure state is obtained; The interlayer interface strength judgment criterion is applied to obtain the delamination failure state judgment result; the interlayer interface strength judgment criterion is used to judge the delamination failure state of the structural response under different generalized stress states; Determine the strength margin of the unit cell failure state based on the structural failure state judgment result and the delamination failure state judgment result; Using different generalized stress states and corresponding strength margins as data sample pairs to obtain the simulation data set; Based on the structural parameters of aviation composite laminated plates and shells, a three-dimensional unit cell model of aviation composite laminated plates and shells is constructed, including: The geometric model of the unit cell is constructed based on the structural parameters of the aviation composite laminated plate and shell, and the finite element mesh is divided to obtain the finite element model; In the finite element model, the stiffness properties of each layer of the unit cell and the contact properties between the layers are defined, the degrees of freedom of the virtual nodes representing the generalized stresses are set, and periodic boundary conditions containing disturbance field corrections are applied between the corresponding virtual nodes in the in-plane direction of the unit cell to obtain an initial three-dimensional unit cell model; Applying distributed coupling constraints to control the rigid body displacement of the initial three-dimensional unit cell model to obtain the three-dimensional unit cell model of the aviation composite laminate shell; Apply generalized stress to various unit sections of the three-dimensional unit cell model of aviation composite laminated plates and shells, perform unit cell multi-condition deformation simulation, and obtain a multi-dimensional generalized stress space, including: Apply unit force to each virtual node in turn, and simulate and determine the stress and interface force response of the aviation composite laminate shell unit cell under multiple unit generalized stress action conditions; the multiple unit generalized stress action conditions include the generalized stress action as tension, the generalized stress action as bending, the generalized stress action as torsion, and the generalized stress action as in-plane shear; Simulate and determine the displacement disturbance field and surface force field of aerospace composite laminated plate and shell unit cells under multiple unit generalized bending moments; Applying the displacement disturbance field and the surface force field on the periodic boundary surface, and simulating and determining the stress and interface force response of the aviation composite laminated plate and shell unit cell under multiple unit generalized out-of-plane shear force conditions; A multidimensional generalized stress space is generated based on the stress and interface force response of aerospace composite laminated plate and shell unit cells under multiple unit generalized stress conditions, and the stress and interface force response of aerospace composite laminated plate and shell unit cells under multiple unit generalized out-of-plane shear force conditions.

2. The strength analysis method of aviation composite laminate shell structure according to claim 1, characterized in that: The geometric model of the unit cell is constructed based on the structural parameters of the aviation composite laminate shell, including: Based on the structural parameters of the aviation composite laminated plate and shell, a geometric model of the unit cell is constructed with the center of the unit cell as the coordinate origin and the thickness direction as a coordinate axis in a manner that the end planes are parallel to each other.

3. The strength analysis method of aviation composite laminate shell structure according to claim 2, characterized in that: Applying distributed coupling constraints to control the rigid body displacement of the initial three-dimensional unit cell model to obtain the three-dimensional unit cell model of the aviation composite laminate shell, including: A reference point is set in the initial three-dimensional unit cell model, and a distributed coupling constraint is applied between the reference point and a line passing through the coordinate origin in the unit cell and perpendicular to the neutral surface of the aviation composite laminate shell, while the translation and rotation of the reference point in three directions are set to 0.

4. The strength analysis method of aviation composite laminate shell structure according to claim 1, characterized in that: The virtual nodes with strength margins within a set range are taken as valid information, the failure value of the structural unit cell under a single load multiplied by the set amplification factor is taken as the sampling area boundary, and random sampling of generalized stress is performed in the multidimensional generalized stress space to obtain the sampling results.

5. A computer device comprising: A memory, a processor, and a computer program stored in the memory and executable on the processor, wherein the processor executes the computer program to implement the strength analysis method for aviation composite laminate shell structures according to any one of claims 1 to 4.

6. A computer-readable storage medium having a computer program stored thereon, characterized in that: When the computer program is executed by a processor, the strength analysis method of aviation composite laminated plate and shell structure described in any one of claims 1 to 4 is implemented.

7. A computer program product, comprising a computer program, characterized in that When the computer program is executed by a processor, the strength analysis method of aviation composite laminated plate and shell structure described in any one of claims 1 to 4 is implemented.

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