Aviation composite material structure fatigue failure process analysis method, equipment and medium

Abstract

The invention discloses an aviation composite material structure fatigue damage process analysis method and device and a medium, and relates to the technical field of aviation composite material structure fatigue damage tolerance analysis, and the method comprises the steps: employing a Newton-Raphson iteration method to calculate an interface fatigue damage model, obtaining an undamaged numerical flux and a thermodynamic conjugate amount of the undamaged numerical flux, calculating to obtain a fatigue damage threshold and a fatigue fracture threshold according to the undamaged numerical flux and the number of fatigue cycles, and calculating to obtain a fatigue fracture threshold according to the thermodynamic conjugate amount of the undamaged numerical flux, the fatigue damage threshold and the fatigue fracture threshold. And determining a fatigue damage state and a fatigue damage degree, and carrying out iterative updating on model parameters of the interface fatigue damage model according to the fatigue damage state and the fatigue damage degree until the current time step is a termination time step. And the precision and efficiency of aviation composite material structure fatigue failure process analysis can be improved.

Description

Technical Field

[0001] The present application relates to the technical field of fatigue damage tolerance analysis of aviation composite materials structures, and in particular to a method, equipment and medium for analyzing fatigue failure process of aviation composite materials structures. Background Art

[0002] In recent years, the application of composite structures in aircraft has been increasing, and the application location has gradually developed from secondary load-bearing structures to primary load-bearing structures. Multilayer structures are the most widely used composite structures on aircraft. Multilayer structures refer to structures with multiple layers made of composite materials, such as fiber-reinforced composite laminates. However, fatigue damage prediction of multilayer structures, especially the coupling problem involving intra-layer and inter-layer damage modes, is still challenging. Therefore, the damage tolerance design of multilayer structures is limited, and the bearing capacity of multilayer structures is not fully utilized.

[0003] Matrix fracture, fiber failure and delamination damage are the three main damage modes of multilayer structures. The interaction between the three main damage modes leads to the growth of complex cracks inside the laminate. At present, the numerical methods for damage simulation of aviation composites can be roughly divided into two categories: continuous damage modeling methods and discrete damage modeling methods. In the finite element method (FEM), the continuous fatigue damage model (CDM) is usually used to simulate the intra-layer fatigue damage (such as fiber failure and matrix fracture) of single-layer or multi-layer structures, and the fatigue cohesive zone model (CZM) is used to simulate the inter-layer delamination fatigue extension of multi-layer structures. Although the above-mentioned existing numerical methods have been proven to be able to accurately and efficiently predict single fatigue damage modes of aviation composites, the fatigue failure process of multilayer structures is very complex, involving the coupling of multiple damage modes within and between layers. Therefore, a numerical method that can predict the coupling of multiple damage modes within and between layers is needed.

[0004] Moreover, in CDM, damage is approximated by reducing the stiffness of the unit, which cannot clearly describe the cracks and the interaction between cracks within and between layers. CZM has been extended to simulate the layered extension of composite materials under fatigue loads. The formula that combines the SN relationship of damage initiation and the Paris law variant of damage growth is a widely used formula. This method does not require the assumption that there is an intrinsic relationship between the evolution of damage variables. However, the use of CZM requires the pre-setting of the crack extension path. Since the fatigue failure process of multi-layer structures is very complex, involving the coupling of multiple damage modes within and between layers, the location and path of complex fatigue cracks are difficult to accurately predict. When the crack extension path is difficult to predict, it is necessary to use extended finite element method (XFEM) and other calculation methods to include a large number of cohesive units, making the model complex and difficult to simulate the extension of complex cracks. There is also the problem of artificial flexibility. Artificial flexibility requires manual adjustment, which is highly subjective and has low accuracy and efficiency.

[0005] Therefore, research on fatigue failure process analysis technology (i.e. fatigue damage evolution analysis) for aviation composite structures is of great significance for the design of aviation composite structures and the verification of their strength and airworthiness compliance. Summary of the invention

[0006] The purpose of this application is to provide a method, equipment and medium for analyzing the fatigue failure process of aviation composite materials, which can predict the coupling of multiple damage modes within and between layers, and improve the accuracy and efficiency of fatigue failure process analysis of aviation composite materials.

[0007] To achieve the above objectives, this application provides the following solutions.

[0008] In a first aspect, the present application provides a method for analyzing fatigue failure process of an aviation composite material structure, the method comprising: The Newton-Raphson iteration method is used to calculate the interface fatigue damage model to obtain the node displacement vector of the aviation composite structure at the current time step, and the node stress vector of the aviation composite structure at the current time step is calculated based on the node displacement vector. The interface parameters of the aviation composite structure at the current time step are calculated based on the node stress vector; the aviation composite structure is a multilayer structure; the interface parameters include the undamaged numerical flux and the thermodynamic conjugate of the undamaged numerical flux; The fatigue damage threshold and fatigue fracture threshold of the aerospace composite material structure at the current time step are calculated based on the undamaged numerical flux and the number of fatigue cycles; the number of fatigue cycles is the number of fatigue cycles of the aerospace composite material structure at the current time step; According to the thermodynamic conjugate of the undamaged numerical flux, the fatigue damage threshold and the fatigue fracture threshold, the fatigue damage state and fatigue damage degree of the aviation composite structure at the current time step are determined; Iteratively updating the model parameters of the interface fatigue damage model according to the fatigue damage state and fatigue damage degree to obtain an updated model; Determine whether the current time step is the termination time step; If yes, then the calculation ends; If not, the updated model is used as the interface fatigue damage model for the next time step, and the process returns to the step of "calculating the interface fatigue damage model using the Newton-Raphson iteration method to obtain the node displacement vector of the aviation composite material structure at the current time step".

[0009] 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 method for analyzing fatigue failure process of aviation composite material structure.

[0010] 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 method for analyzing fatigue failure process of aviation composite material structures.

[0011] According to the specific embodiments provided in this application, this application has the following technical effects: The present application provides a fatigue failure process analysis method, device and medium for aviation composite material structure, which uses the Newton-Raphson iteration method to calculate the interface fatigue damage model to obtain the interface parameters of the aviation composite material structure at the current time step, wherein the interface parameters include the undamaged numerical flux and the thermodynamic conjugate of the undamaged numerical flux, and the fatigue damage threshold and fatigue fracture threshold of the aviation composite material structure at the current time step are calculated based on the undamaged numerical flux and the number of fatigue cycles, and the fatigue damage state and fatigue damage degree of the aviation composite material structure at the current time step are determined based on the thermodynamic conjugate of the undamaged numerical flux, the fatigue damage threshold and the fatigue fracture threshold, and the fatigue damage state and fatigue damage degree of the aviation composite material structure are determined based on the fatigue damage state and fatigue damage degree. The model parameters of the fatigue damage model are iteratively updated to obtain the updated model, and the iteration is continued until the current time step is the termination time step. The present application introduces the fatigue damage state and fatigue damage degree into the interface fatigue damage model to update the interface fatigue damage model, and uses the interface fatigue damage model to calculate the node strain vectors and node stress vectors of the aviation composite structure at multiple time steps. A model can be used to simultaneously simulate the intra-layer and inter-layer fatigue damage of the aviation composite structure without inserting a large number of cohesive units and artificial flexibility, which can reduce the complexity of the model and avoid artificial influences, thereby being able to predict the coupling of multiple damage modes within and between layers, and can improve the accuracy and efficiency of the fatigue failure process analysis of the aviation composite structure. BRIEF DESCRIPTION OF THE DRAWINGS

[0012] 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.

[0013] Figure 1 This is an application environment diagram of a method for analyzing fatigue failure process of aviation composite material structures provided in Example 1 of the present application.

[0014] Figure 2 A schematic flow chart of a method for analyzing fatigue failure process of aviation composite material structures provided in Example 1 of the present application.

[0015] Figure 3 A schematic diagram of the sub-domain division method provided in Example 1 of the present application.

[0016] Figure 4 A schematic diagram of an interface fatigue damage model provided in Example 1 of the present application.

[0017] Figure 5 A schematic diagram of another interface fatigue damage model provided in Example 1 of the present application.

[0018] Figure 6 Schematic diagram of the DCB (Double Cantilever Beam) specimen and fatigue simulation results provided in Example 1 of the present application.

[0019] Figure 7 This is a schematic diagram comparing the experimental data and numerical results of the crack growth rate of the DCB sample provided in Example 1 of the present application.

[0020] Figure 8 Schematic diagram of the fatigue delamination behavior of the DCB sample provided in Example 1 of the present application.

[0021] Fig. 9 Schematic diagram of the ENF (End Notched Flexure) specimen and fatigue simulation results provided in Example 1 of the present application.

[0022] Fig.10 This is a schematic diagram comparing the experimental data and numerical results of the crack growth rate of the ENF specimen provided in Example 1 of the present application.

[0023] Fig.11 Schematic diagram of the MMB (Mixed Mode Bending) specimen and fatigue simulation results provided in Example 1 of the present application.

[0024] Fig.12 Schematic diagram of the fatigue delamination behavior of the MMB specimen provided in Example 1 of the present application.

[0025] Fig.13 Schematic diagram of the CTB (Clamped Tapered Beam) specimen provided in Example 1 of the present application.

[0026] Fig.14 This is a schematic diagram of the fatigue simulation results of the CTB specimen provided in Example 1 of the present application; wherein, Fig.14 (a) is a schematic diagram of the injury starting position; Fig.14 (b) is a schematic diagram of the crack transition position.

[0027] Fig.15 Schematic diagram of fatigue delamination initiation and extension behavior of the CTB specimen provided in Example 1 of the present application.

[0028] Fig.16 A schematic diagram of the structure of a computer device provided in Example 2 of the present application. DETAILED DESCRIPTION

[0029] 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.

[0030] Example 1.

[0031] The fatigue failure process analysis method of aviation composite material structure provided in the embodiment of the present application can be applied to Figure 1 In the application environment shown. Among them, the terminal communicates with the server through the network. The data storage system can store the data that the server needs to process. The data storage system can be set up separately, integrated on the server, or placed on the cloud or other servers. The terminal can send the analysis request to be processed to the server. After the server receives the analysis request to be processed, for the analysis request to be processed, the server uses the Newton-Raphson iteration method to calculate the interface fatigue damage model to obtain the node displacement vector of the aviation composite structure at the current time step, and calculate the node stress vector of the aviation composite structure at the current time step according to the node displacement vector. According to the node stress vector, the interface parameters of the aviation composite structure at the current time step are calculated, and the interface parameters include the undamaged numerical flux and the thermodynamic conjugate of the undamaged numerical flux; according to the undamaged numerical flux and the number of fatigue cycles, the aviation composite structure at the current time step is calculated. The fatigue damage threshold and fatigue fracture threshold of the step are determined; the fatigue damage state and fatigue damage degree of the aviation composite structure at the current time step are determined according to the thermodynamic conjugate of the undamaged numerical flux, the fatigue damage threshold and the fatigue fracture threshold; the model parameters of the interface fatigue damage model are iteratively updated according to the fatigue damage state and fatigue damage degree to obtain the updated model; it is determined whether the current time step is the termination time step; if so, the calculation is terminated; if not, the updated model is used as the interface fatigue damage model of the next time step, and the step of "using the Newton-Raphson iteration method to calculate the interface fatigue damage model and obtain the node displacement vector of the aviation composite structure at the current time step" is returned. The server can feed back the analysis results obtained for the analysis request to the terminal.

[0032] In addition, in some embodiments, the fatigue failure process analysis method of aviation composite structures can also be implemented independently by a server or a terminal. For example, the terminal can directly process the analysis request to be processed, or the server can obtain the analysis request to be processed from the data storage system and process the analysis request to be processed.

[0033] The terminals may be, but are not limited to, various desktop computers, laptops, smart phones, tablet computers, IoT devices and portable wearable devices. IoT devices may be smart speakers, smart TVs, smart air conditioners, smart car-mounted devices, etc. Portable wearable devices may be smart watches, smart bracelets, head-mounted devices, etc. The server may be implemented as an independent server or a server cluster consisting of multiple servers, or a cloud server.

[0034] In an exemplary embodiment, Figure 2 As shown, a method for analyzing fatigue failure process of aviation composite material structure is provided. The method is executed by a computer device, specifically, it can be executed by a computer device such as a terminal or a server alone, or it can be executed by a terminal and a server together. In the embodiment of the present application, the method is applied to Figure 1 The following steps are used to illustrate the server in the example.

[0035] Step S1, using the Newton-Raphson iteration method to calculate the interface fatigue damage model, obtain the node displacement vector of the aviation composite material structure at the current time step, calculate the node stress vector of the aviation composite material structure at the current time step according to the node displacement vector, and calculate the interface parameters of the aviation composite material structure at the current time step according to the node stress vector; the aviation composite material structure is a multilayer structure; the interface parameters include the undamaged numerical flux and the thermodynamic conjugate of the undamaged numerical flux.

[0036] Step S2, calculating the fatigue damage threshold and fatigue fracture threshold of the aviation composite material structure at the current time step according to the undamaged numerical flux and the fatigue cycle number; the fatigue cycle number is the fatigue cycle number of the aviation composite material structure at the current time step.

[0037] Step S3, determining the fatigue damage state and fatigue damage degree of the aviation composite material structure at the current time step according to the thermodynamic conjugate of the undamaged numerical flux, the fatigue damage threshold and the fatigue fracture threshold.

[0038] Step S4, iteratively updating the model parameters of the interface fatigue damage model according to the fatigue damage state and the fatigue damage degree to obtain an updated model.

[0039] Step S5, determining whether the current time step is a termination time step.

[0040] Step S6: If yes, then end the calculation.

[0041] Step S7, if not, use the updated model as the interface fatigue damage model for the next time step, and return to the step of "using the Newton-Raphson iteration method to calculate the interface fatigue damage model to obtain the node displacement vector of the aviation composite material structure at the current time step".

[0042] By implementing the above-mentioned steps S1 to S7, the present embodiment designs an interface fatigue damage model, calculates the fatigue damage threshold, fatigue fracture threshold, fatigue damage state and fatigue damage degree in sequence, and introduces the fatigue damage state and fatigue damage degree into the interface fatigue damage model to update the interface fatigue damage model. The interface fatigue damage model is used to calculate the node strain vectors and node stress vectors of the aviation composite structure at multiple time steps. One model can be used to simulate the intra-layer and inter-layer fatigue damage of the aviation composite structure at the same time, avoiding the problem of low precision caused by the existing numerical method that needs to use two models to simulate the intra-layer and inter-layer fatigue damage of the aviation composite structure respectively, and there is no need to insert a large number of cohesive units and artificial flexibility, which can reduce the complexity of the model, reduce human influence, and improve efficiency and precision.

[0043] In this embodiment, the aviation composite material structure is a multi-layer structure, such as a fiber-reinforced composite laminate.

[0044] In this embodiment, the interface fatigue damage model is: ; in, is the tangential stiffness matrix; is the set of displacement increment vectors of nodes of all subdomains, where the subdomain is the area obtained by dividing the analysis area of ​​the aerospace composite structure, the analysis area is part of or all of the aerospace composite structure, and the node is the center point of the subdomain; is the global external load vector; is the global internal force vector.

[0045] Tangential stiffness matrix It is expressed as follows: ; in, is a subdomain in the analysis area; It is the analysis area of ​​aerospace composite structures; is the strain matrix; is the second-order elastic tensor; is the boundary of the subdomain; is the set of inner boundaries between all adjacent subdomains in the analysis area; [ ] is the jump operator; is the shape function; is the transformation matrix between the global coordinate system and the local coordinate system; is the local tangent stiffness of the inner interface; is the interface stiffness tensor in the global coordinate system; {} is the average operator; is the projective transformation matrix between the stress tensor and the traction tensor at the interface.

[0046] Global internal force vector It is expressed as follows: ; in, is the Cauchy stress tensor at the inner boundary; is the second-order unit tensor, is the global fatigue damage tensor; is the undamaged numerical flux in the global coordinate system, which is the undamaged numerical flux in the global coordinate system calculated in the previous time step.

[0047] Shape Function It is expressed as follows: ; in, is the shape function; and are the coordinates of any point in the subdomain and the coordinates of the nodes of the subdomain respectively; is the gradient operator; is the support domain matrix.

[0048] ; in, is the number of subdomains contained in the support domain corresponding to the subdomain, and the support domain is the set of other subdomains adjacent to the subdomain.

[0049] After establishing the above-mentioned interface fatigue damage model, this embodiment uses the Newton-Raphson iteration method to calculate the interface fatigue damage model to obtain the node displacement vector of the aviation composite structure at the current time step, and the node displacement vector includes the displacement of the node of each subdomain of the aviation composite structure. The node stress vector of the aviation composite structure at the current time step is calculated based on the node displacement vector, and the node stress vector includes the stress of the node of each subdomain of the aviation composite structure. The process of calculating the node stress vector based on the node displacement vector is an existing conventional process and will not be repeated here. The interface parameters of the aviation composite structure at the current time step are calculated based on the node stress vector, and the interface parameters include the undamaged numerical flux and the thermodynamic conjugate of the undamaged numerical flux.

[0050] In this embodiment, the calculation formula of the undamaged numerical flux is: ; in, is the undamaged numerical flux in the global coordinate system; {} is the average operator; is the Cauchy stress tensor at the inner boundary, which is determined according to the node stress vector, and specifically the node stress vector is used as the Cauchy stress tensor at the inner boundary; is the Cauchy stress tensor at the inner boundary and the undamaged numerical flux in the global coordinate system The projection tensor between ; is the interface stiffness tensor in the global coordinate system; [ ] is the jump operator; is the displacement test function.

[0051] In this embodiment, the calculation formula of the thermodynamic conjugate of the undamaged numerical flux is: ; in, is the thermodynamic conjugate of the undamaged numerical flux in the local coordinate system; is the interface stiffness tensor in the local coordinate system; is the undamaged numerical flux in the local coordinate system.

[0052] It should be noted that a parameter can be easily converted from the global coordinate system to the local coordinate system through the conversion matrix between the global coordinate system and the local coordinate system, and a parameter can also be easily converted from the local coordinate system to the global coordinate system.

[0053] After obtaining the interface parameters, this embodiment calculates the fatigue damage threshold and fatigue fracture threshold of the aviation composite material structure at the current time step based on the undamaged numerical flux and the number of fatigue cycles, wherein the number of fatigue cycles is the number of fatigue cycles of the aviation composite material structure at the current time step.

[0054] When in the crack initiation stage, the calculation formula of fatigue damage threshold is: ; in, is the fatigue damage threshold; is the normal component of the interface stiffness tensor in the local coordinate system; is the tensile strength of the interface under fatigue load, determined according to the number of fatigue cycles; is the interface shear strength under fatigue load, determined according to the number of fatigue cycles; is the component ratio.

[0055] Interfacial tensile strength under fatigue loading It can be expressed as the original tensile strength of the interface and fatigue cycles The function is as follows: ; in, is the original tensile strength of the interface; is the first material parameter; is the number of fatigue cycles.

[0056] Interface shear strength under fatigue loading It can be expressed as the original shear strength of the interface and fatigue cycles The function is as follows: ; in, is the original shear strength of the interface; is the second material parameter.

[0057] ; in, is the tangential component of the undamaged numerical flux in the local coordinate system; is the normal component of the undamaged numerical flux in the local coordinate system.

[0058] The calculation formula of fatigue fracture threshold is: ; in, is the fatigue fracture threshold; is the interface fracture energy; is the interface stiffness tensor in the global coordinate system.

[0059] The critical energy release rate of the mixed mode is determined according to the BK law. At this time, the calculation formula of the interface fracture energy is: ; in, is the interfacial tensile fracture energy; is the interface shear fracture energy; is the mixed mode ratio, , is the shear strain energy release rate, is the tensile strain energy release rate; Parameters determined for the experiment.

[0060] After obtaining the fatigue damage threshold and the fatigue fracture threshold, this embodiment determines the fatigue damage state and fatigue damage degree of the aviation composite material structure at the current time step according to the thermodynamic conjugate of the undamaged numerical flux, the fatigue damage threshold and the fatigue fracture threshold. Specifically, the thermodynamic conjugate of the undamaged numerical flux and the fatigue damage threshold are used as inputs, and the fatigue damage state of the aviation composite material structure at the current time step is determined based on the interface fatigue damage criterion. The fatigue fracture threshold is further combined to determine the fatigue damage degree based on the fatigue damage state, and the fatigue damage condition of the aviation composite material structure is updated.

[0061] In this embodiment, the fatigue damage state and fatigue damage degree of the aviation composite material structure at the current time step are determined according to the thermodynamic conjugate of the undamaged numerical flux, the fatigue damage threshold and the fatigue fracture threshold, which specifically includes the following steps.

[0062] (1) Determine whether the thermodynamic conjugate of the undamaged numerical flux of the interface of the aerospace composite structure in the global coordinate system is greater than the fatigue damage threshold; if so, the fatigue damage state of the interface is interface damage; if not, the fatigue damage state of the interface is interface undamaged.

[0063] The interface fatigue damage criterion is: if the thermodynamic conjugate of the undamaged numerical flux in the global coordinate system of the interface Fatigue damage threshold , then the fatigue damage state of the interface is interface damage. If the thermodynamic conjugate of the undamaged numerical flux in the global coordinate system of the interface Fatigue damage threshold , then the fatigue damage state of the interface is that the interface is not damaged.

[0064] (2) If the fatigue damage state of the interface is interface damage, the thermodynamic conjugate of the undamaged numerical flux in the global coordinate system of the interface, the fatigue damage threshold, and the fatigue fracture threshold are used as inputs, and the fatigue damage degree calculation formula is used to calculate the fatigue damage degree of the interface at the current time step.

[0065] The fatigue damage calculation formula is: ; in, is the fatigue damage degree; is the fatigue damage factor; is the fatigue damage threshold; is the fatigue fracture threshold.

[0066] ; in, is the intermediate parameter; The minimum number of fatigue cycles required to cause any interface in the analysis area of ​​an aerospace composite structure to break; The number of fatigue cycles required to cause the current interface to break.

[0067] ; in, It is the thermodynamic conjugate of the undamaged numerical flux in the global coordinate system of the current interface.

[0068] ; in, is the effective characteristic length of the interface; is the third material parameter; is the maximum strain energy release rate of the interface in the current state; is the fourth material parameter.

[0069] (3) If the fatigue damage state of the interface is that the interface is not damaged, no additional operation is required, that is, the fatigue damage degree will not be updated, and the fatigue damage degree of the interface in the current time step is determined to be equal to the fatigue damage degree of the interface in the previous time step.

[0070] After obtaining the fatigue damage state and the fatigue damage degree, this embodiment iteratively updates the model parameters of the interface fatigue damage model according to the fatigue damage state and the fatigue damage degree to obtain an updated model. Specifically, based on the updated fatigue damage situation (i.e., the fatigue damage state and the fatigue damage degree), the model parameters of the interface fatigue damage model are iteratively updated by Newton-Raphson to obtain the updated model.

[0071] The model parameters of the interface fatigue damage model include the local tangent stiffness of the inner interface and shape function .

[0072] Since some undamaged interfaces may be affected and damaged due to the occurrence of fatigue damage, in this embodiment, the model parameters of the interface fatigue damage model are iteratively updated according to the fatigue damage state and fatigue damage degree to obtain an updated model, which specifically includes the following steps.

[0073] (1) If the fatigue damage state of the interface is that the interface is not damaged, the first update formula is used to update the local tangent stiffness of the inner interface; if the fatigue damage state of the interface is that the interface is damaged, the second update formula is used to update the local tangent stiffness of the inner interface; if the fatigue damage degree of the interface is 1, the interface is fractured, the support domain of the subdomain is adjusted, and the crack is explicitly introduced. The subdomains on both sides of the inner interface where the crack is located are removed from each other's support domains, and the shape function is reconstructed to obtain the first updated model.

[0074] (2) The node displacement vector, node stress vector and interface parameters of the current iteration step are calculated using the first update model; the quasi-static damage threshold and quasi-static fracture threshold of the current iteration step are calculated based on the interface parameters of the current iteration step; the quasi-static damage state and quasi-static damage degree of the current iteration step are calculated based on the quasi-static damage threshold and quasi-static fracture threshold of the current iteration step; based on the quasi-static damage state and quasi-static damage degree of the current iteration step, the local tangent stiffness of the inner interface of the first update model is updated to obtain the second update model.

[0075] (3) The node displacement vector of the next iteration step is calculated using the second updated model, and it is determined whether the ratio of the node displacement vector of the next iteration step to the node displacement vector of the current iteration step is less than a preset ratio. If so, the iteration is terminated and the second updated model is used as the updated model; if not, the iteration is continued and the second updated model is used as the first updated model of the next iteration step, and the process returns to the step of "calculating the node displacement vector, node stress vector and interface parameters of the current iteration step using the first updated model".

[0076] In this embodiment, the first update formula is: ; in, is the local tangent stiffness of the inner interface; is the thermodynamic conjugate of the undamaged numerical flux in the local coordinate system; is the quasi-static damage degree; is the first matrix; is the damage-free interface free energy in the local coordinate system.

[0077] ; in, is the damage-free interface free energy in the global coordinate system; is the interface stiffness tensor in the global coordinate system.

[0078] The second update formula is: ; in, is the second matrix; is the quasi-static damage threshold; is the quasi-static fracture threshold; is the interface stiffness tensor in the global coordinate system; It is the thermodynamic conjugate of the undamaged numerical flux in the global coordinate system of the current interface.

[0079] ; in, is the normal component of the undamaged numerical flux in the local coordinate system.

[0080] Among them, the calculation formula of quasi-static damage degree is: ; in, is the quasi-static damage degree; is the quasi-static damage factor; is the quasi-static damage threshold; is the quasi-static fracture threshold.

[0081] ; in, It is the thermodynamic conjugate of the undamaged numerical flux in the global coordinate system of the current interface.

[0082] When in the crack initiation stage, the calculation formula of the quasi-static damage threshold is: ; in, is the quasi-static damage threshold; is the normal component of the interface stiffness tensor in the local coordinate system; is the interfacial tensile strength; is the interface shear strength; is the component ratio.

[0083] The calculation formula of the quasi-static fracture threshold is: ; in, is the quasi-static fracture threshold; is the interface fracture energy; is the interface stiffness tensor in the global coordinate system.

[0084] At this time, if the fatigue damage state of the interface is that the interface is not damaged, the quasi-static damage threshold, quasi-static fracture threshold and quasi-static damage degree are calculated, and then the local tangent stiffness of the inner interface is updated using the first update formula; if the fatigue damage state of the interface is interface damage, the quasi-static damage threshold, quasi-static fracture threshold and quasi-static damage degree are calculated, and then the local tangent stiffness of the inner interface is updated using the second update formula. At the same time, the shape function is updated to obtain the first update model.

[0085] Among them, according to the quasi-static damage threshold and the quasi-static fracture threshold of the current iteration step, the quasi-static damage state and the quasi-static damage degree of the current iteration step are calculated, which specifically includes the following steps.

[0086] (1) Determine whether the thermodynamic conjugate of the undamaged numerical flux of the interface in the global coordinate system is greater than the quasi-static damage threshold; if so, the quasi-static damage state of the interface is interface damage; if not, the quasi-static damage state of the interface is interface undamaged.

[0087] (2) If the quasi-static damage state of the interface is interface damage, the quasi-static damage degree of the interface is calculated using the thermodynamic conjugate of the undamaged numerical flux in the global coordinate system of the interface, the quasi-static damage threshold, and the quasi-static fracture threshold as inputs using the quasi-static damage degree calculation formula.

[0088] (3) If the quasi-static damage state of the interface is that the interface is not damaged, then the quasi-static damage degree of the interface is determined to be equal to the quasi-static damage degree of the interface in the previous iteration step.

[0089] Among them, based on the quasi-static damage state and quasi-static damage degree of the current iteration step, the local tangent stiffness of the inner interface of the first updated model is updated to obtain the second updated model, specifically including: if the quasi-static damage state of the interface is that the interface is not damaged, the local tangent stiffness of the inner interface is updated by the first update formula; if the quasi-static damage state of the interface is interface damage, the local tangent stiffness of the inner interface is updated by the second update formula to obtain the second updated model.

[0090] After obtaining the updated model, this embodiment determines whether the current time step is the termination time step. If so, the calculation is terminated. If not, the updated model is used as the interface fatigue damage model for the next time step, and the process returns to the step of "calculating the interface fatigue damage model using the Newton-Raphson iteration method to obtain the node displacement vector of the aviation composite material structure at the current time step", and enters the next time step until all time steps are calculated.

[0091] This embodiment proposes a finite element numerical method based on a numerical flux interface fatigue damage model, which is used to simulate the complex damage of fiber reinforced composite laminates under fatigue conditions. It is a new numerical method for predicting the coupled fatigue damage between layers in laminates. The interface fatigue damage model is calculated using the Newton-Raphson iteration method to obtain the interface parameters of the aviation composite structure; based on the fatigue load and interface parameters, the fatigue damage threshold and fatigue fracture threshold of the aviation composite structure are obtained; the fatigue damage threshold and fatigue fracture threshold are used as inputs, and the fatigue damage state and fatigue damage degree of the aviation composite structure are calculated based on the interface fatigue damage criterion and the interface fatigue damage evolution law (i.e., the fatigue damage degree calculation formula), and it is iterated. The above-mentioned interface fatigue damage criterion between the matrix and the layer is derived based on the SN curve, which is used to describe the influence of fatigue damage accumulation on the residual strength of the interface. The above-mentioned interface fatigue damage evolution law between the matrix and the layer is derived using the Paris formula, which is used to simulate the crack propagation process under interface fatigue damage. This embodiment analyzes the fatigue damage process of the aviation composite structure by using the "broken point method" and the corresponding interface fatigue damage model, which can solve the problem of difficulty in simulating the destruction process of the mixed fatigue damage mode.

[0092] In the following, the fatigue failure process analysis method of aviation composite structures is verified through multiple tests. Three damage benchmark tests, namely DCB test, ENF test and MMB test, are carried out. They are type I, type II and mixed modal delamination tests respectively. The test is regarded as a plane strain problem in a two-dimensional case. The material parameters used in the test are shown in Table 1.

[0093] Table 1 Material parameters used in the test

[0094] In Table 1, , is the elastic modulus of the material, is the material shear modulus, is the Poisson's ratio, is the original tensile strength of the interface, is the original shear strength of the interface, is the interfacial tensile fracture energy, is the interface shear fracture energy, is the third material parameter, is the fourth material parameter, and These are all parameters in the Paris formula of the material. , , The units are , , The units are , , The unit is N / mm.

[0095] like Figure 3 As shown in the figure, it is a schematic diagram of subdomain division, and the support domain corresponding to the subdomain is clearly defined. The blue one is the subdomain, and the green one is the support domain corresponding to the subdomain. For example, for the example of irregular subdomain, For subdomains, , , , , is the supporting domain corresponding to the subdomain, Figure 4 and Figure 5 All of them are interface fatigue damage models. Figure 4 and Figure 5 middle, is the damage-containing numerical flux, is the interface strength, is the interfacial fracture energy, is the displacement jump, is the quasi-static interface strength, is the fatigue interface strength, is the fatigue damage threshold, is the quasi-static damage threshold, is the quasi-static fracture threshold, is the fatigue fracture threshold, is the thermodynamic conjugate of the undamaged numerical flux in the local coordinate system, is the numerical flux of damage on the interface at the current time step, is the thermodynamic conjugate damage threshold, is the thermodynamic conjugate quantity of the interface at the current time step, is the thermodynamic conjugation breaking threshold, Figure 4 The left side is a schematic diagram of the fatigue cohesion zone model. Figure 4 The dark blue triangle on the right is the interface strength reduction (i.e., the introduction of and Interface fatigue damage model (after), Figure 5 After considering the Paris formula (i.e. introducing Interface fatigue damage model (after), Figure 4 The dark blue triangle on the right turns into Figure 5 The solid black line plus the dotted red line in Figure 5 The black solid line plus the red dotted line in FIG. 1 is the interface fatigue damage model used in this embodiment.

[0096] like Figure 6As shown in the figure, in the DCB test, the beam used is a composite laminate with a size of length L = 150 mm and width d = 3.1 mm. A DCB specimen is supported at one end and a pair of concentrated lateral fatigue displacement loads of equal amplitude and opposite direction are applied to the other end. , load The maximum and minimum action amplitudes are and The initial delamination starts from the free end of the beam to the neutral plane with a length of a = 30.5 mm, as shown in Figure 7 In FPM (FragilePoint Method), the specimen model is divided into quadrilateral subdomains and initial stratification is applied. As the number of fatigue cycles increases, the crack propagation is as follows: Figure 8 As shown, Figure 8 In this article, Carvalho et al. 2019 refers to A new approach to model delamination growth in fatigue using the Virtual Crack Closure Techniquewithout re-meshing, published by De Carvalho NV et al. in Engineering Fracture Mechanics in 2019.

[0097] Figure 7 The simulated crack growth rate and the maximum energy release rate at the crack tip are shown in Figure 1, which is in good agreement with the experimental results. The crack length-fatigue cycle number relationship results simulated by VCCT (Virtual Crack Closure Technique) are plotted together with the simulation results of this embodiment. Figure 8 It is verified that this embodiment can better predict type I delamination of composite laminates.

[0098] like Fig. 9 As shown in Figure 1, the ENF test can be used to simulate the type II delamination of composite laminates. The selected specimen length L = 150 mm and width D = 3.1 mm. One end of the beam is fixed and the other end is restrained in the longitudinal direction. is the distance from the initial neutral plane to the clamped end. At a distance of C=25mm from the beam end, the same fatigue load P is applied at points A and B. The initial delamination extends from the clamped end to the neutral plane of the specimen with a length of 30.5mm.

[0099] The relationship between the crack growth rate and the normalized maximum energy release rate at the crack tip obtained by simulation in this embodiment is as follows: Fig.10 As shown, it is in good agreement with the experimental results.

[0100] like Fig.11 As shown in Figure 1, in the MMB test, the specimen forms a mixed delamination consisting of type I and type II under load. The beam is 2L = 100.8 mm long and 2d = 4.5 mm wide. One end is fixed and the other end is longitudinally restrained. The loading arm is fixed to the beam at points A and B. The mixed mode ratio is applied at point C, which is c = 41.3 mm away from point B. Downward displacement fatigue load , The maximum value of , Minimum value of ,length The initial neutral surface delamination begins at the beam end at point A.

[0101] like Fig.12 As shown in FIG. 1 , the growth of the delamination length predicted by this embodiment with the increase of fatigue cycle number is consistent with the baseline case. Fig.12 Here, R Krueger et al. 2016 refers to In search of a time efficient approach to crack and delamination growth predictions in composites, published by Krueger R et al. in ASC Technical Conference in 2016.

[0102] NASA (National Aeronautics and Space Administration) designed a representative specimen called CTB, which consists of a cross-laminated beam clamped at both ends and a loading section, such as Fig.13 As shown. One end of the short beam is tapered, the distance from the left clamping end is a=40.16mm, b=44.7mm, the maximum displacement is L=0.825mm, and the fatigue load of R=0.2 is applied to the loading part ( Fig.13 The right end of the specimen (blue part in the middle) is c=57.5mm away from the left end of the specimen. The total length of the CTB specimen is l=115mm, excluding the clamped ends, and the thickness is w=5.12mm. The specimen is IM7 / 8552 laminate, the material parameters are shown in Tables 2 and 3, and the stacking order is [0 / 903 / 0 / 90 / 02 / 90 / 0 / 902 / 0 / 90 / 903 / 02 / 90 / 02 / 902].

[0103] Table 2 IM7 / 8552 material parameters

[0104] Table 3 IM7 / 8552 fatigue material parameters

[0105] In Table 3, C I , C II 、m I 、m II are the parameters in the Paris formula, S1 and S2 are the parameters in the SN curve.

[0106] like Fig.14 As shown in (a), the initial crack appears in the conical area of ​​the specimen. Subsequently, the matrix crack propagates along the separation interface between the long and short beams, with different material accumulation directions on both sides. Finally, the crack will migrate upward through the matrix crack to the upper interface and continue to propagate there. Fig.15 The expansion result of the extended delamination (which may be located at the initial interface or at the upper interface of the migration) is a function of the number of cycles. Compared with the experimental results, the prediction of this embodiment is generally consistent with that of the experimental results. Fig.15 In the text, Kenneth L. R. et. al. 2020 refers to the Durability of aerospace material systems published by Reifsnider KL et al. in Durability of Composite Systems. Woodhead Publishing in 2020. In particular, it should be noted that there is a slow growth area before the migration, which is believed to be related to multiple migration attempts. Fig.14 (b) in the figure confirms that in this region, both the interlayer and intralayer interfaces are damaged and competed, and as the fatigue load increases, matrix cracking becomes the main damage mode leading to crack migration.

[0107] The present application also provides an application scenario, which applies the above-mentioned method for analyzing the fatigue failure process of aviation composite materials. Specifically, the method for analyzing the fatigue failure process of aviation composite materials provided in this embodiment can be applied in the design scenario of aviation composite materials structure. The design scenario of aviation composite materials structure includes a simulation link and a design link. The simulation link is used to simulate and analyze the fatigue failure process of the aviation composite materials structure, and the design link is used to design the aviation composite materials structure based on the simulation analysis results. The method for analyzing the fatigue failure process of aviation composite materials structure provided in this embodiment belongs to the simulation link.

[0108] Example 2.

[0109] 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.16As shown. The computer device includes a processor, a memory, an input / output interface (Input / Output, referred to as I / O) and a communication interface. 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. 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 data. 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 method for analyzing the fatigue failure process of an aviation composite material structure is implemented.

[0110] Those skilled in the art will understand that Fig.16 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.

[0111] 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 when the processor executes the computer program, the method for analyzing fatigue failure process of aviation composite material structure in Example 1 is implemented.

[0112] Example 3.

[0113] 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 method for analyzing fatigue failure process of aviation composite material structure in embodiment 1 is implemented.

[0114] 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.

[0115] 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.

[0116] 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 method for analyzing fatigue failure process of aviation composite material structure, characterized in that: The fatigue failure process analysis method of aviation composite material structure comprises: The Newton-Raphson iteration method is used to calculate the interface fatigue damage model to obtain the node displacement vector of the aviation composite structure at the current time step, and the node stress vector of the aviation composite structure at the current time step is calculated based on the node displacement vector. The interface parameters of the aviation composite structure at the current time step are calculated based on the node stress vector; the aviation composite structure is a multilayer structure; the interface parameters include the undamaged numerical flux and the thermodynamic conjugate of the undamaged numerical flux; The fatigue damage threshold and fatigue fracture threshold of the aerospace composite material structure at the current time step are calculated based on the undamaged numerical flux and the number of fatigue cycles; the number of fatigue cycles is the number of fatigue cycles of the aerospace composite material structure at the current time step; According to the thermodynamic conjugate of the undamaged numerical flux, the fatigue damage threshold and the fatigue fracture threshold, the fatigue damage state and fatigue damage degree of the aviation composite structure at the current time step are determined; Iteratively updating the model parameters of the interface fatigue damage model according to the fatigue damage state and fatigue damage degree to obtain an updated model; Determine whether the current time step is the termination time step; If yes, then the calculation ends; If not, the updated model is used as the interface fatigue damage model for the next time step, and the process returns to the step of "calculating the interface fatigue damage model using the Newton-Raphson iteration method to obtain the node displacement vector of the aviation composite material structure at the current time step".

2. The method for analyzing fatigue failure process of aviation composite material structure according to claim 1, characterized in that: The calculation formula of the undamaged numerical flux is: ； in, is the undamaged numerical flux in the global coordinate system; {} is the average operator; is the Cauchy stress tensor at the inner boundary, determined from the nodal stress vectors; is the Cauchy stress tensor at the inner boundary and the undamaged numerical flux in the global coordinate system The projection tensor between ; is the interface stiffness tensor in the global coordinate system; [ ] is the jump operator; is the displacement test function; The thermodynamic conjugate of the undamaged numerical flux is calculated as: ； in, is the thermodynamic conjugate of the undamaged numerical flux in the local coordinate system; is the interface stiffness tensor in the local coordinate system; is the undamaged numerical flux in the local coordinate system.

3. The method for analyzing fatigue failure process of aviation composite material structure according to claim 1, characterized in that: The calculation formula of fatigue damage threshold is: ； in, is the fatigue damage threshold; is the normal component of the interface stiffness tensor in the local coordinate system; is the tensile strength of the interface under fatigue load, determined according to the number of fatigue cycles; is the interface shear strength under fatigue load, determined according to the number of fatigue cycles; is the component ratio; ； in, is the original tensile strength of the interface; is the first material parameter; is the number of fatigue cycles; ； in, is the original shear strength of the interface; is the second material parameter; ； in, is the tangential component of the undamaged numerical flux in the local coordinate system; is the normal component of the undamaged numerical flux in the local coordinate system; The calculation formula of fatigue fracture threshold is: ； in, is the fatigue fracture threshold; is the interface fracture energy; is the interface stiffness tensor in the global coordinate system.

4. The method for analyzing fatigue failure process of aviation composite material structure according to claim 1, characterized in that: According to the thermodynamic conjugate of the undamaged numerical flux, the fatigue damage threshold and the fatigue fracture threshold, the fatigue damage state and fatigue damage degree of the aviation composite structure at the current time step are determined, including: Determine whether the thermodynamic conjugate of the undamaged numerical flux in the global coordinate system of the interface of the aviation composite structure is greater than the fatigue damage threshold; if so, the fatigue damage state of the interface is interface damage; if not, the fatigue damage state of the interface is interface undamaged; If the fatigue damage state of the interface is interface damage, the thermodynamic conjugate of the undamaged numerical flux in the global coordinate system of the interface, the fatigue damage threshold and the fatigue fracture threshold are used as inputs, and the fatigue damage degree calculation formula is used to calculate the fatigue damage degree of the interface at the current time step; If the fatigue damage state of the interface is that the interface is not damaged, it is determined that the fatigue damage degree of the interface in the current time step is equal to the fatigue damage degree of the interface in the previous time step; The fatigue damage calculation formula is: ； in, is the fatigue damage degree; is the fatigue damage factor; is the fatigue damage threshold; is the fatigue fracture threshold; ； in, is the intermediate parameter; The minimum number of fatigue cycles required to cause any interface in the analysis area of ​​an aerospace composite structure to break; The number of fatigue cycles required to break the current interface; ； in, is the thermodynamic conjugate of the undamaged numerical flux in the global coordinate system of the current interface; ； in, is the effective characteristic length of the interface; is the third material parameter; is the maximum strain energy release rate of the interface in the current state; is the fourth material parameter.

5. The method for analyzing fatigue failure process of aviation composite material structure according to claim 1, characterized in that: The model parameters of the interface fatigue damage model include the local tangent stiffness and shape function of the inner interface. The model parameters of the interface fatigue damage model are iteratively updated according to the fatigue damage state and fatigue damage degree to obtain the updated model, which specifically includes: If the fatigue damage state of the interface is that the interface is not damaged, the first update formula is used to update the local tangent stiffness of the inner interface; if the fatigue damage state of the interface is that the interface is damaged, the second update formula is used to update the local tangent stiffness of the inner interface; if the fatigue damage degree of the interface is 1, the interface is fractured, the support domain of the subdomain is adjusted, and the shape function is reconstructed to obtain the first update model; The node displacement vector, node stress vector and interface parameters of the current iteration step are calculated using the first update model; the quasi-static damage threshold and quasi-static fracture threshold of the current iteration step are calculated based on the interface parameters of the current iteration step; the quasi-static damage state and quasi-static damage degree of the current iteration step are calculated based on the quasi-static damage threshold and quasi-static fracture threshold of the current iteration step; based on the quasi-static damage state and quasi-static damage degree of the current iteration step, the local tangent stiffness of the inner interface of the first update model is updated to obtain the second update model; The node displacement vector of the next iteration step is calculated using the second updated model, and it is determined whether the ratio of the node displacement vector of the next iteration step to the node displacement vector of the current iteration step is less than a preset ratio. If so, the iteration is terminated and the second updated model is used as the updated model; if not, the iteration is continued and the second updated model is used as the first updated model of the next iteration step, and the process returns to the step of "calculating the node displacement vector, node stress vector and interface parameters of the current iteration step using the first updated model".

6. The method for analyzing fatigue failure process of aviation composite material structure according to claim 5, characterized in that: The first update formula is: ； in, is the local tangent stiffness of the inner interface; is the thermodynamic conjugate of the undamaged numerical flux in the local coordinate system; is the quasi-static damage degree; is the first matrix; is the damage-free interface free energy in the local coordinate system; The second update formula is: ； in, is the second matrix; is the quasi-static damage threshold; is the quasi-static fracture threshold; is the interface stiffness tensor in the global coordinate system; is the thermodynamic conjugate of the undamaged numerical flux in the global coordinate system of the current interface; ； in, is the normal component of the undamaged numerical flux in the local coordinate system.

7. The method for analyzing fatigue failure process of aviation composite material structure according to claim 5, characterized in that: The expression of the shape function is: ； in, is the shape function; and are the coordinates of any point in the subdomain and the coordinates of the nodes in the subdomain respectively; is the gradient operator; is the support domain matrix; ； in, is the number of subdomains contained in the support domain corresponding to the subdomain, and the support domain is the set of other subdomains adjacent to the subdomain.

8. The method for analyzing fatigue failure process of aviation composite material structure according to claim 5, characterized in that: The calculation formula of quasi-static damage degree is: ； in, is the quasi-static damage degree; is the quasi-static damage factor; is the quasi-static damage threshold; is the quasi-static fracture threshold; The calculation formula of quasi-static damage threshold is: ； in, is the normal component of the interface stiffness tensor in the local coordinate system; is the interfacial tensile strength; is the interface shear strength; is the component ratio; ； in, is the tangential component of the undamaged numerical flux in the local coordinate system; is the normal component of the undamaged numerical flux in the local coordinate system; The calculation formula of the quasi-static fracture threshold is: ； in, is the quasi-static fracture threshold; is the interface fracture energy; is the interface stiffness tensor in the global coordinate system.

9. 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 method for analyzing fatigue failure process of aviation composite material structure according to any one of claims 1 to 8.

10. A computer-readable storage medium having a computer program stored thereon, characterized in that: When the computer program is executed by a processor, the method for analyzing fatigue failure process of aviation composite material structure according to any one of claims 1 to 8 is implemented.

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