Analysis Method, Equipment and Medium for Fatigue Failure Process of Aerospace Composite Material Structure

Abstract

The present application discloses a method, device and medium for analyzing the fatigue failure process of an aviation composite material structure, relating to the technical field of fatigue damage tolerance analysis of aviation composite material structures. The method includes: calculating an interface fatigue damage model by using the Newton-Raphson iteration method to obtain an undamaged numerical flux and the thermodynamic conjugate quantity of the undamaged numerical flux; calculating a fatigue damage threshold and a fatigue fracture threshold according to the undamaged numerical flux and the number of fatigue cycles; determining a fatigue damage state and a fatigue damage degree according to the thermodynamic conjugate quantity of the undamaged numerical flux, the fatigue damage threshold and the fatigue fracture threshold; and iteratively updating the 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 the termination time step. The present application can predict the coupling of multiple damage modes within and between layers, and can improve the accuracy and efficiency of analyzing the fatigue failure process of aviation composite material structures.

Description

Technical Field

[0001] This application relates to the technical field of fatigue damage tolerance analysis of aircraft composite structures, and particularly to a method, device, and medium for analyzing the fatigue failure process of aircraft composite structures. Background Art

[0002] In recent years, the application of composite structures in aircraft has been increasing continuously, and the application parts have gradually developed from secondary load-bearing structures to primary load-bearing structures. The multi-layer structure is the most widely used composite structure in aircraft. The multi-layer structure refers to a structure with multiple layers made of composite materials, such as fiber-reinforced composite laminates. However, for the fatigue damage prediction of multi-layer structures, especially for the coupling problem of intra-layer and inter-layer damage modes, it is still challenging. Therefore, the damage tolerance design of multi-layer structures is restricted, and the load-bearing capacity of multi-layer structures has not been fully exerted.

[0003] Matrix fracture, fiber failure, and delamination damage are the three main damage modes of multi-layer 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 simulating aircraft composite material damage can be roughly divided into two categories: continuous damage modeling methods and discrete damage modeling methods. In the finite element method (FEM), the continuum damage mechanics (CDM) model 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 cohesive zone model (CZM) is used to simulate the inter-layer delamination fatigue propagation of multi-layer structures. Although the above existing numerical methods have been proven to be able to accurately and efficiently predict the single fatigue damage mode of aircraft composite materials, the fatigue failure process of multi-layer structures is very complex, involving the coupling of multiple intra-layer and inter-layer damage modes. Therefore, a numerical method capable of predicting the coupling of multiple intra-layer and inter-layer damage modes is needed.

[0004] Moreover, in CDM, damage is approximately calculated by reducing the stiffness of elements, which cannot clearly describe cracks and the interaction between intra - ply and inter - ply cracks. CZM has been extended to simulate the delamination growth of composite materials under fatigue loads. The formula combining the S - N relationship of damage initiation and a variant of Paris' law for damage growth is a widely used formula. This method does not require assuming an inherent relationship in the evolution of damage variables. However, using CZM requires pre - setting the crack propagation path. Since the fatigue failure process of multi - layer structures is very complex, involving the coupling of multiple damage modes within and between layers, it is difficult to accurately predict the position and path of complex fatigue cracks. When the crack propagation path is difficult to predict, computational methods such as the extended finite element method (XFEM) need to incorporate a large number of cohesive elements, making the model complex, difficult to simulate the propagation of complex cracks, and there are artificial flexibility problems. Artificial flexibility requires manual adjustment, which is highly subjective and has low accuracy and efficiency.

[0005] Therefore, the research on technologies applicable to the analysis of the fatigue failure process of aviation composite material structures (i.e., fatigue damage evolution analysis) is of great significance for both the design of aviation composite material structures and the verification of strength airworthiness compliance. Summary of the Invention

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

[0007] To achieve the above - mentioned purpose, the following solutions are provided in this application.

[0008] In the first aspect, this application provides a method for analyzing the fatigue failure process of aviation composite material structures. The method for analyzing the fatigue failure process of aviation composite material structures includes:

[0009] 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. Calculate the node stress vector of the aviation composite material structure at the current time step based on the node displacement vector, and calculate the interface parameters of the aviation composite material structure at the current time step based on the node stress vector. The aviation composite material structure is a multi - layer structure. The interface parameters include undamaged numerical fluxes and the thermodynamic conjugate quantities of undamaged numerical fluxes;

[0010] Calculate the fatigue damage threshold and fatigue fracture threshold of the aviation composite material structure at the current time step according to the undamaged numerical fluxes and the number of fatigue cycles. The number of fatigue cycles is the number of fatigue cycles of the aviation composite material structure at the current time step;

[0011] Determine the fatigue damage state and fatigue damage degree of the aviation composite material structure at the current time step according to the thermodynamic conjugate quantity of the undamaged numerical flux, the fatigue damage threshold, and the fatigue fracture threshold;

[0012] Iteratively update the model parameters of the interface fatigue damage model according to the fatigue damage state and fatigue damage degree to obtain the updated model;

[0013] Judge whether the current time step is the termination time step;

[0014] If so, end the calculation;

[0015] 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 nodal displacement vector of the aviation composite material structure at the current time step".

[0016] In a second aspect, the present application provides a computer device, including: a memory, a processor, and a computer program stored on the memory and executable on the processor, where the processor executes the computer program to implement the above-mentioned method for analyzing the fatigue failure process of an aviation composite material structure.

[0017] In a third aspect, the present application provides a computer-readable storage medium, on which a computer program is stored, and when the computer program is executed by a processor, it implements the above-mentioned method for analyzing the fatigue failure process of an aviation composite material structure.

[0018] According to the specific embodiments provided by the present application, the present application has the following technical effects:

[0019] The present application provides a method, device, and medium for analyzing the fatigue failure process of an aviation composite material structure. The Newton-Raphson iteration method is used to calculate the interface fatigue damage model to obtain the interface parameters of the aviation composite material structure at the current time step. The interface parameters include the undamaged numerical flux and the thermodynamic conjugate quantity of the undamaged numerical flux. According to the undamaged numerical flux and the number of fatigue cycles, the fatigue damage threshold and the fatigue fracture threshold of the aviation composite material structure at the current time step are calculated. According to the thermodynamic conjugate quantity of the undamaged numerical flux, the fatigue damage threshold, and the fatigue fracture threshold, the fatigue damage state and the fatigue damage degree of the aviation composite material structure at the current time step are determined. According to the fatigue damage state and the fatigue damage degree, the model parameters of the interface fatigue damage model are iteratively updated to obtain an updated model. Iteration is continuously performed until the current time step is the termination time step. The present application introduces the fatigue damage state and the 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 nodal strain vector and the nodal stress vector of the aviation composite material structure at multiple time steps. It can use one model to simultaneously simulate the in-layer and inter-layer fatigue damage of the aviation composite material structure, and does not require inserting a large number of cohesive elements and artificial flexibility, which can reduce the model complexity and avoid artificial influence. Therefore, it can predict the coupling of various in-layer and inter-layer damage modes, and can improve the accuracy and efficiency of the fatigue failure process analysis of the aviation composite material structure. BRIEF DESCRIPTION OF THE DRAWINGS

[0020] In order to more clearly illustrate the technical solutions in the embodiments of the present application or the prior art, the following will briefly introduce the drawings required for use in the embodiments. Obviously, the drawings described below are only some embodiments of the present application. For those of ordinary skill in the art, other drawings can be obtained based on these drawings without creative efforts.

[0021] Figure 1 It is an application environment diagram of a method for analyzing the fatigue failure process of an aviation composite material structure provided in Embodiment 1 of the present application.

[0022] Figure 2 It is a flowchart of a method for analyzing the fatigue failure process of an aviation composite material structure provided in Embodiment 1 of the present application.

[0023] Figure 3 It is a schematic diagram of the subdomain division method provided in Embodiment 1 of the present application.

[0024] Figure 4 It is a schematic diagram of an interface fatigue damage model provided in Embodiment 1 of the present application.

[0025] Figure 5 It is another schematic diagram of an interface fatigue damage model provided in Embodiment 1 of the present application.

[0026] Figure 6 Schematic diagram of the DCB (Double Cantilever Beam) specimen and fatigue simulation results provided in Embodiment 1 of the present application.

[0027] Figure 7 Schematic diagram for comparing experimental data and numerical results of the crack propagation rate of the DCB specimen provided in Embodiment 1 of the present application.

[0028] Figure 8 Schematic diagram of the fatigue delamination behavior of the DCB specimen provided in Embodiment 1 of the present application.

[0029] Figure 9 Schematic diagram of the ENF (End Notched Flexure) specimen and fatigue simulation results provided in Embodiment 1 of the present application.

[0030] Figure 10 Schematic diagram for comparing experimental data and numerical results of the crack propagation rate of the ENF specimen provided in Embodiment 1 of the present application.

[0031] Figure 11 Schematic diagram of the MMB (Mixed Mode Bending) specimen and fatigue simulation results provided in Embodiment 1 of the present application.

[0032] Figure 12 Schematic diagram of the fatigue delamination behavior of the MMB specimen provided in Embodiment 1 of the present application.

[0033] Figure 13 Schematic diagram of the CTB (Clamped Tapered Beam) specimen provided in Embodiment 1 of the present application.

[0034] Figure 14 Schematic diagram of the fatigue simulation results of the CTB specimen provided in Embodiment 1 of the present application; wherein, Figure 14 (a) in is a schematic diagram of the damage initiation position; Figure 14 (b) in is a schematic diagram of the crack transition position.

[0035] Figure 15 Schematic diagram of the fatigue delamination initiation and propagation behavior of the CTB specimen provided in Embodiment 1 of the present application.

[0036] Figure 16 Schematic diagram of the structure of a computer device provided in Embodiment 2 of the present application. Detailed implementation manners

[0037] Next, the technical solutions in the embodiments of the present application will be clearly and completely described in conjunction with the accompanying drawings in the embodiments of the present application. Obviously, the described embodiments are only a part of the embodiments of the present application, rather than all the embodiments. Based on the embodiments in the present application, all other embodiments obtained by those of ordinary skill in the art without creative efforts shall fall within the protection scope of the present application.

[0038] Embodiment 1.

[0039] The method for analyzing the fatigue failure process of an aviation composite material structure provided by the embodiment of the present application can be applied to an application environment as Figure 1 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 separately, integrated on the server, placed in the cloud or on other servers. The terminal can send an analysis request to be processed to the server. After receiving 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, obtains the node displacement vector of the aviation composite material structure at the current time step, calculates the node stress vector of the aviation composite material structure at the current time step according to the node displacement vector, calculates the interface parameter of the aviation composite material structure at the current time step according to the node stress vector, and the interface parameter includes the undamaged numerical flux and the thermodynamic conjugate quantity of the undamaged numerical flux; calculates the fatigue damage threshold and the fatigue fracture threshold of the aviation composite material structure at the current time step according to the undamaged numerical flux and the number of fatigue cycles; determines the fatigue damage state and the fatigue damage degree of the aviation composite material structure at the current time step according to the thermodynamic conjugate quantity of the undamaged numerical flux, the fatigue damage threshold and the fatigue fracture threshold; 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; determines whether the current time step is the termination time step; if so, ends the calculation; if not, uses the updated model as the interface fatigue damage model for the next time step, and returns 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". The server can feedback the analysis result obtained for the analysis request to the terminal.

[0040] In addition, in some embodiments, the method for analyzing the fatigue failure process of an aviation composite material structure can also be implemented separately by the server or the 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.

[0041] Among them, the terminal can be, but is not limited to, various desktop computers, laptop computers, smart phones, tablet computers, Internet of Things devices, and portable wearable devices. The Internet of Things devices can be smart speakers, smart TVs, smart air conditioners, smart in-vehicle devices, etc., and the portable wearable devices can be smart watches, smart bracelets, head-mounted devices, etc. The server can be implemented by an independent server or a server cluster composed of multiple servers, and can also be a cloud server.

[0042] In an exemplary embodiment, as Figure 2 shown, a method for analyzing the fatigue failure process of an aviation composite material structure is provided. This method is executed by a computer device, and specifically can be executed alone by a computer device such as a terminal or a server, or can be jointly executed by a terminal and a server. In the embodiments of the present application, taking this method applied to Figure 1 the server in

[0043] as an example for illustration, the method includes the following steps.

[0044] Step S1: Use 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. 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 multi-layer structure. The interface parameters include the undamaged numerical flux and the thermodynamic conjugate quantity of the undamaged numerical flux.

[0045] Step S2: Calculate 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 number of fatigue cycles. The number of fatigue cycles is the number of fatigue cycles of the aviation composite material structure at the current time step.

[0046] Step S3: Determine the fatigue damage state and fatigue damage degree of the aviation composite material structure at the current time step according to the thermodynamic conjugate quantity of the undamaged numerical flux, the fatigue damage threshold, and the fatigue fracture threshold.

[0047] Step S4: Iteratively update the model parameters of the interface fatigue damage model according to the fatigue damage state and fatigue damage degree to obtain an updated model.

[0047] Step S5: Determine whether the current time step is the termination time step.

[0048] Step S6: If so, end the calculation.

[0049] Step S7, if the answer is no, then 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".

[0050] Implementing the above steps S1 to S7, in this embodiment, an interface fatigue damage model is designed to calculate the fatigue damage threshold, fatigue fracture threshold, fatigue damage state, and fatigue damage degree in sequence, and introduce the fatigue damage state and fatigue damage degree into the interface fatigue damage model to update the interface fatigue damage model. Using the interface fatigue damage model to calculate the node strain vector and node stress vector of the aviation composite material structure at multiple time steps can simulate the intra-layer and inter-layer fatigue damage of the aviation composite material structure using one model, avoiding the problem of low accuracy caused by the need to use two models to separately simulate the intra-layer and inter-layer fatigue damage of the aviation composite material structure in the existing numerical methods, and without the need to insert a large number of cohesive units and artificial flexibility, which can reduce the model complexity, reduce the human influence, and improve the efficiency and accuracy.

[0051] In this embodiment, the aviation composite material structure is a multi-layer structure, such as a fiber-reinforced composite laminate.

[0052] In this embodiment, the interface fatigue damage model is:

[0053] ;

[0054] Among them, is the tangential stiffness matrix; is the set of displacement increment vectors of the nodes of all subdomains. A subdomain is a region obtained by dividing the analysis region of the aviation composite material structure. The analysis region is a part or all of the aviation composite material structure, and the node is the center point of the subdomain; is the global external load vector; is the global internal force vector.

[0055] The tangential stiffness matrix is expressed as follows:

[0056] ;

[0057] Among them, is the subdomain in the analysis region; is the analysis region of the aviation composite material structure; 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 region; [ ] 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 internal interface; is the interface stiffness tensor in the global coordinate system; {} is the averaging operator; is the projection transformation matrix between the stress tensor and the surface force tensor at the interface.

[0058] Global internal force vector is expressed as follows:

[0059] ;

[0060] where is the Cauchy stress tensor at the internal 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 at the previous time step.

[0061] Shape function is expressed as follows:

[0062] ;

[0063] where is the shape function; and are the coordinates of an arbitrary point within the subdomain and the coordinates of the nodes of that subdomain, respectively; is the gradient operator; is the support domain matrix.

[0064] ;

[0065] where is the number of subdomains contained in the support domain corresponding to this subdomain, and the support domain is the set of other subdomains adjacent to this subdomain.

[0066] After establishing the above interface fatigue damage model, in this embodiment, the Newton-Raphson iteration method is used to calculate the interface fatigue damage model to obtain the nodal displacement vector of the aerospace composite structure at the current time step. The nodal displacement vector includes the displacements of the nodes in each sub-domain of the aerospace composite structure. The nodal stress vector of the aerospace composite structure at the current time step is calculated based on the nodal displacement vector. The nodal stress vector includes the stresses of the nodes in each sub-domain of the aerospace composite structure. The process of calculating the nodal stress vector based on the nodal displacement vector is a conventional process and will not be elaborated here. The interface parameters of the aerospace composite structure at the current time step are calculated based on the nodal stress vector. The interface parameters include the undamaged numerical flux and the thermodynamic conjugate quantity of the undamaged numerical flux.

[0067] In this embodiment, the calculation formula for the undamaged numerical flux is:

[0068] ;

[0069] Where, 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 based on the nodal stress vector. Specifically, the nodal stress vector is used as the Cauchy stress tensor at the inner boundary; is the Cauchy stress tensor at the inner boundary and the projection tensor between the undamaged numerical flux in the global coordinate system; is the interface stiffness tensor in the global coordinate system; [ ] is the jump operator; is the displacement trial function.

[0070] In this embodiment, the calculation formula for the thermodynamic conjugate quantity of the undamaged numerical flux is:

[0071] ;

[0072] Where, is the thermodynamic conjugate quantity 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.

[0073] It should be noted that through the transformation matrix between the global coordinate system and the local coordinate system, a certain parameter can be conveniently transformed from the global coordinate system to the local coordinate system, and a certain parameter can also be conveniently transformed from the local coordinate system to the global coordinate system.

[0074] After obtaining the interface parameters, in this embodiment, according to the undamaged numerical flux and the number of fatigue cycles, the fatigue damage threshold and the fatigue fracture threshold of the aviation composite material structure at the current time step are calculated, where the number of fatigue cycles is the number of fatigue cycles of the aviation composite material structure at the current time step.

[0075] When in the crack initiation stage, the calculation formula for the fatigue damage threshold is:

[0076] ;

[0077] Where, is the fatigue damage threshold; is the normal component of the interface stiffness tensor in the local coordinate system; is the interface tensile strength 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.

[0078] The interface tensile strength under fatigue load can be expressed as a function of the original interface tensile strength and the number of fatigue cycles as follows:

[0079] ;

[0080] Where, is the original interface tensile strength; is the first material parameter; is the number of fatigue cycles.

[0081] The interface shear strength under fatigue load can be expressed as a function of the original interface shear strength and the number of fatigue cycles as follows:

[0082] ;

[0083] Where, is the original interface shear strength; is the second material parameter.

[0084] ;

[0085] Where, 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.

[0086] The calculation formula for the fatigue fracture threshold is:

[0087] ;

[0088] Among them, is the fatigue fracture threshold; is the interface fracture energy; is the interface stiffness tensor in the global coordinate system.

[0089] The critical energy release rate of the mixed mode is determined according to the B-K law. At this time, the calculation formula for the interface fracture energy is:

[0090] ;

[0091] Among them, is the interface 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; is a parameter determined by experiments.

[0092] After obtaining the fatigue damage threshold and the fatigue fracture threshold, in this embodiment, based on the thermodynamic conjugate quantity of the undamaged numerical flux, the fatigue damage threshold, and the fatigue fracture threshold, the fatigue damage state and the fatigue damage degree of the aviation composite material structure at the current time step are determined. Specifically, taking the thermodynamic conjugate quantity of the undamaged numerical flux and the fatigue damage threshold as inputs, the fatigue damage state of the aviation composite material structure at the current time step is determined based on the interface fatigue damage criterion. Further, in combination with the fatigue fracture threshold, the fatigue damage degree is determined based on the fatigue damage state, and the fatigue damage condition of the aviation composite material structure is updated.

[0093] In this embodiment, based on the thermodynamic conjugate quantity of the undamaged numerical flux, the fatigue damage threshold, and the fatigue fracture threshold, the fatigue damage state and the fatigue damage degree of the aviation composite material structure at the current time step are determined, which specifically includes the following steps.

[0094] (1) Determine whether the thermodynamic conjugate quantity of the undamaged numerical flux of the interface of the aviation composite material 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.

[0095] The interface fatigue damage criterion is: if the thermodynamic conjugate quantity of the undamaged numerical flux of the interface in the global coordinate system fatigue damage threshold , then the fatigue damage state of the interface is interface damage, if the thermodynamic conjugate quantity of the undamaged numerical flux of the interface in the global coordinate system Fatigue damage threshold , the fatigue damage state of the interface is that the interface is undamaged.

[0096] (2) If the fatigue damage state of the interface is that the interface is damaged, then using the thermodynamic conjugate quantity of the undamaged numerical flux in the global coordinate system of the interface, the fatigue damage threshold, and the fatigue fracture threshold as inputs, the fatigue damage degree of the interface at the current time step is calculated using the fatigue damage degree calculation formula.

[0097] The fatigue damage degree calculation formula is:

[0098] ;

[0099] Among them, is the fatigue damage degree; is the fatigue damage factor; is the fatigue damage threshold; is the fatigue fracture threshold.

[0100] ;

[0101] Among them, is the intermediate parameter; is the minimum number of fatigue cycles required for any interface in the analysis area of the aviation composite material structure to fracture; is the number of fatigue cycles required for the current interface to fracture.

[0102] ;

[0103] Among them, is the thermodynamic conjugate quantity of the undamaged numerical flux in the global coordinate system of the current interface.

[0104] ;

[0105] Among them, 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.

[0106] (3) If the fatigue damage state of the interface is that the interface is undamaged, then no additional operation is required, that is, the fatigue damage degree will not be updated, and it is determined that the fatigue damage degree of the interface at the current time step is equal to the fatigue damage degree of the interface at the previous time step.

[0107] After obtaining the fatigue damage state and fatigue damage degree, 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. Specifically, based on the updated fatigue damage condition (i.e., the fatigue damage state and fatigue damage degree), the model parameters of the interface fatigue damage model are updated by Newton-Raphson iteration to obtain an updated model.

[0108] The model parameters of the interface fatigue damage model include the local tangent stiffness of the inner interface and the shape function .

[0109] Since some undamaged interfaces may be affected by the appearance of fatigue damage and thus become damaged, 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.

[0110] (1) If the fatigue damage state of the interface is that the interface is undamaged, the local tangent stiffness of the inner interface is updated using the first update formula; if the fatigue damage state of the interface is that the interface is damaged, the local tangent stiffness of the inner interface is updated using the second update formula; if the fatigue damage degree of the interface is 1, the interface breaks, the support domain of the subdomain is adjusted, specifically, a 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.

[0111] (2) The node displacement vector, node stress vector, and interface parameters of the current iteration step are calculated using the first updated 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 according to the quasi-static damage threshold and quasi-static fracture threshold of the current iteration step; the local tangent stiffness of the inner interface of the first updated model is updated based on the quasi-static damage state and quasi-static damage degree of the current iteration step to obtain the second updated model.

[0112] (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 ends, and the second updated model is used as the updated model; if not, the iteration continues, the second updated model is used as the first updated model of the next iteration step, and the step of "calculating the node displacement vector, node stress vector, and interface parameters of the current iteration step using the first updated model" is returned.

[0113] In this embodiment, the first update formula is:

[0114] ;

[0115] Among them, is the local tangent stiffness of the internal interface; is the thermodynamic conjugate quantity of the undamaged numerical flux in the local coordinate system; is the quasi-static damage degree; is the first matrix; is the undamaged interface free energy in the local coordinate system.

[0116] ;

[0117] Among them, is the undamaged interface free energy in the global coordinate system; is the interface stiffness tensor in the global coordinate system.

[0118] The second update formula is:

[0119] ;

[0120] Among them, 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 quantity of the undamaged numerical flux of the current interface in the global coordinate system.

[0121] ;

[0122] Among them, is the normal component of the undamaged numerical flux in the local coordinate system.

[0123] Among them, the calculation formula of the quasi-static damage degree is:

[0124] ;

[0125] Among them, is the quasi-static damage degree; is the quasi-static damage factor; is the quasi-static damage threshold; is the quasi-static fracture threshold.

[0126] ;

[0127] Among them, is the thermodynamic conjugate quantity of the undamaged numerical flux of the current interface in the global coordinate system.

[0128] When in the crack initiation stage, the calculation formula of the quasi-static damage threshold is:

[0129] ;

[0130] wherein, is the quasi-static damage threshold; is the normal component of the interface stiffness tensor in the local coordinate system; is the interface tensile strength; is the interface shear strength; is the component ratio.

[0131] The calculation formula for the quasi-static fracture threshold is:

[0132] ;

[0133] wherein, is the quasi-static fracture threshold; is the interface fracture energy; is the interface stiffness tensor in the global coordinate system.

[0134] At this time, if the fatigue damage state of the interface is that the interface is undamaged, the quasi-static damage threshold, the quasi-static fracture threshold and the quasi-static damage degree are calculated, and then the local tangent stiffness of the inner interface is updated by using the first update formula; if the fatigue damage state of the interface is that the interface is damaged, the quasi-static damage threshold, the quasi-static fracture threshold and the quasi-static damage degree are calculated, and then the local tangent stiffness of the inner interface is updated by using the second update formula. At the same time, the shape function is updated to obtain the first update model.

[0135] 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 include the following steps.

[0136] (1) Judge whether the thermodynamic conjugate quantity 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 that the interface is damaged; if not, the quasi-static damage state of the interface is that the interface is undamaged.

[0137] (2) If the quasi-static damage state of the interface is that the interface is damaged, taking the thermodynamic conjugate quantity of the undamaged numerical flux of the interface in the global coordinate system, the quasi-static damage threshold and the quasi-static fracture threshold as inputs, the quasi-static damage degree of the interface is calculated by using the quasi-static damage degree calculation formula.

[0138] (3) If the quasi-static damage state of the interface is that the interface is undamaged, it is determined that the quasi-static damage degree of the interface is equal to the quasi-static damage degree of the interface in the previous iteration step.

[0139] 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, which specifically includes: if the quasi-static damage state of the interface is that the interface is undamaged, the local tangent stiffness of the inner interface is updated using the first update formula; if the quasi-static damage state of the interface is that the interface is damaged, the local tangent stiffness of the inner interface is updated using the second update formula to obtain the second updated model.

[0140] After obtaining the updated model, this embodiment determines whether the current time step is the termination time step. If so, the calculation ends; if not, the updated model is used as the interface fatigue damage model for the next time step, and the step of "using the Newton-Raphson iteration method to calculate the interface fatigue damage model to obtain the nodal displacement vector of the aerospace composite structure at the current time step" is returned, entering the next time step until all time steps are calculated.

[0141] This embodiment proposes a finite element numerical method based on a numerical flux interface fatigue damage model for simulating the complex damage of fiber-reinforced composite laminates in the fatigue state. It is a new numerical method for predicting the interlaminar coupling fatigue damage within the laminate. The Newton-Raphson iteration method is used to calculate the interface fatigue damage model to obtain the interface parameters of the aerospace composite structure; based on the fatigue load and interface parameters, the fatigue damage threshold and fatigue fracture threshold of the aerospace composite structure are obtained; taking the fatigue damage threshold and fatigue fracture threshold as inputs, the fatigue damage state and fatigue damage degree of the aerospace 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 iteration is performed. The above interface fatigue damage criteria for the matrix and interlayer are derived based on the S-N curve and are used to describe the influence of fatigue damage accumulation on the remaining strength of the interface. The above interface fatigue damage evolution law for the matrix and interlayer is derived using the Paris formula and is used to simulate the crack propagation process under interface fatigue damage. This embodiment analyzes the fatigue damage process of the aerospace composite structure by adopting the "broken point method" and the supporting interface fatigue damage model, and can solve the problem of difficult simulation of the failure process of the mixed fatigue damage mode.

[0142] Next, the fatigue failure process analysis method of the aerospace composite structure is verified through multiple tests. Three damage benchmark tests, namely the DCB test, ENF test, and MMB test, are carried out respectively. They are type I, type II, and mixed-mode delamination tests respectively. The tests are regarded as plane strain problems in two-dimensional cases, and the material parameters used in the tests are shown in Table 1.

[0143] Table 1 Material parameters used in the tests

[0144]

[0145] In Table 1, and are the elastic modulus of the material, is the shear modulus of the material, is the Poisson's ratio, is the original tensile strength of the interface, is the original shear strength of the interface, is the tensile fracture energy of the interface, is the shear fracture energy of the interface, is the third material parameter, is the fourth material parameter, and are both parameters in the Paris formula of the material. and and all have the unit of , and all have the unit of , and all have the unit of N / mm.

[0146] As shown in Figure 3 , it is a schematic diagram of subdomain division, and the corresponding support domain of the subdomain is clarified. The blue part is the subdomain, and the green part is the support domain corresponding to the subdomain. For example, for the example of an irregular subdomain, is the subdomain, and and and and are the support domains corresponding to the subdomain, Figure 4 and Figure 5 are both interface fatigue damage models. Figure 4 and Figure 5 In is the numerical flux containing damage, is the interface strength, is the interface 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 quantity of the undamaged numerical flux in the local coordinate system, is the numerical flux containing damage of this interface at the current time step, is the damage threshold of the thermodynamic conjugate quantity, is the thermodynamic conjugate quantity of the interface at the current time step, is the fracture threshold of the thermodynamic conjugate quantity, Figure 4 The left side is a schematic diagram of the fatigue cohesive zone model, Figure 4 The dark blue triangle on the right side is the interface fatigue damage model after the interface strength is reduced (i.e., after introducing and ), Figure 5 is the interface fatigue damage model after considering the Paris formula (i.e., after introducing ), Figure 4 The dark blue triangle on the right side becomes Figure 5 the black solid line plus the red dashed line in Figure 5 The black solid line plus the red dashed line in

[0147] As Figure 6 shown, in the DCB test, the beam used is a composite laminate, and the size configuration is length L = 150 mm and width d = 3.1 mm. One end of the DCB specimen is supported solidly, and a pair of concentrated transverse fatigue displacement loads in equal amplitude and opposite directions is applied at the other end , the load The maximum and minimum action amplitudes of are and . The initial delamination starts from the free end of the beam towards the neutral plane with a length of a = 30.5 mm, as Figure 7 shown. In the FPM (Fragile Point Method), the specimen model is divided into quadrilateral subdomains and the initial delamination is applied. As the number of fatigue cycles increases, the crack propagation is as Figure 8 shown, Figure 8 In, Carvalho et al. 2019 refers to A new approach to model delamination growth in fatigue using the Virtual Crack Closure Technique without re-meshing published by De Carvalho N V et al. in Engineering Fracture Mechanics in 2019.

[0148] Figure 7 shows the functional relationship between the simulated crack propagation rate and the maximum energy release rate at the crack tip, which is in good agreement with the experimental results. The results of the crack length - number of fatigue cycles relationship simulated by VCCT (Virtual Crack Closure Technique) and the simulation results of this embodiment are plotted together in Figure 8In this case, it is verified that this embodiment can better predict the mode I delamination of composite laminates.

[0149] As Figure 9 shown, the ENF test can be used to simulate the mode 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 longitudinally constrained. is the distance from the initial neutral plane to the fixed end. At a distance C = 25 mm from the beam end, the same fatigue load P is applied at points A and B. The initial delamination extends from the fixed end towards the neutral plane of the specimen with a length of 30.5 mm.

[0150] The relationship between the crack growth rate obtained by simulation in this embodiment and the normalized maximum energy release rate at the crack tip is as Figure 10 shown, which is in good agreement with the experimental results.

[0151] As Figure 11 shown, in the MMB test, a mixed delamination composed of mode I and mode II is formed in the specimen under the action of the load. The beam length 2L = 100.8 mm, width 2d = 4.5 mm, one end is fixed, and the other end is longitudinally constrained. The loading arms are fixed to the beam at points A and B, and a downward displacement fatigue load with a mixed mode ratio is applied at point C, which is c = 41.3 mm away from point B. , the maximum value of , the minimum value of , length The initial neutral surface delamination starts from the beam end where point A is located.

[0152] As Figure 12 shown, the growth of the delamination length predicted in this embodiment with the increase of the number of fatigue cycles is in agreement with the reference case. Figure 12 In

[0153] 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 part, as Figure 13As shown. One end of the short beam is tapered, with distances a = 40.16 mm and b = 44.7 mm from the left clamping end. The maximum displacement L = 0.825 mm. A fatigue load with R = 0.2 is applied to the right end of the loading part ( Figure 13 the blue part in), at a distance c = 57.5 mm from the left end of the specimen. The total length of the CTB specimen is l = 115 mm, excluding the two clamped ends, and the thickness is w = 5.12 mm. The specimen is an IM7 / 8552 laminate, and the material parameters are shown in Tables 2 and 3. The stacking sequence is [0 / 903 / 0 / 90 / 02 / 90 / 0 / 902 / 0 / 90 / 903 / 02 / 90 / 02 / 902].

[0154] Table 2 IM7 / 8552 Material Parameters

[0155]

[0156] Table 3 IM7 / 8552 Fatigue Material Parameters

[0157]

[0158] In Table 3, C I 、C II 、m I 、m II are parameters in the Paris formula, and S1, S2 are parameters in the S-N curve.

[0159] As Figure 14 shown in (a) of, the initial crack appears in the tapered region of the specimen. Subsequently, the matrix crack propagates along the separation interface of the long and short beams, with different material stacking directions on both sides. Finally, the crack will migrate upward to the upper interface through the matrix crack and continue to propagate there. Figure 15 is the function of the propagation result of the extended delamination (which may be located at the initial interface or the migrated upper interface) with the number of cycles. Compared with the experimental results, the prediction in this embodiment is generally in good agreement. Figure 15 In, Kenneth L. R. et al. 2020 refers to Durability of aerospace material systems published by Reifsnider K L et al. in 2020 in Durability of Composite Systems. Woodhead Publishing. It should be particularly noted that there is a region with slow growth before migration, which is considered to be related to multiple migration attempts. Figure 14 As confirmed by (b) in, in this region, both the interlaminar and intralaminar interfaces are damaged and competing. With the increase of the fatigue load, matrix cracking becomes the main damage mode leading to crack migration.

[0160] The present application also provides an application scenario, which applies the above method for analyzing the fatigue failure process of an aviation composite material structure. Specifically, the method for analyzing the fatigue failure process of an aviation composite material structure provided in this embodiment can be applied in the aviation composite material structure design scenario. The aviation composite material structure design scenario 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 material structure, and the design link is used to design the aviation composite material structure based on the simulation analysis results. The method for analyzing the fatigue failure process of an aviation composite material structure provided in this embodiment belongs to the simulation link.

[0161] Embodiment 2.

[0162] In an exemplary embodiment, a computer device is provided. The computer device can be a server or a terminal, and its internal structure diagram can be as Figure 16 shown. The computer device includes a processor, a memory, an input / output interface (Input / Output, abbreviated 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 data. The input / output interface of the computer device is used to exchange information between the processor and external devices. The communication interface of the computer device is used to communicate with external terminals through a network connection. When the computer program is executed by the processor, it implements a method for analyzing the fatigue failure process of an aviation composite material structure.

[0163] Those skilled in the art can understand that Figure 16 the structure shown in

[0164] is only a block diagram of some structures 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 some components, or have different component arrangements.

[0165] Embodiment 3.

[0166] In an exemplary embodiment, a computer-readable storage medium is provided, storing a computer program which, when executed by a processor, implements the method for analyzing the fatigue failure process of an aviation composite material structure in Embodiment 1.

[0167] 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 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 need to comply with relevant regulations.

[0168] The technical features of the above embodiments can be combined arbitrarily. For the sake of brevity of description, 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, it should be considered as the scope recorded in this specification.

[0169] In this article, specific examples are used to elaborate on the principle and implementation manner of this application. The description of the above embodiments is only used to help understand the method and its core idea of this application; at the same time, for those of ordinary skill in the art, according to the idea of this application, there will be changes in the specific implementation manner and application scope. In summary, the content of this specification should not be construed as a limitation to 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 material structure at the current time step are determined, specifically including: judging whether the thermodynamic conjugate of the undamaged numerical flux in the global coordinate system of the interface of the aviation composite material 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; 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 "using the Newton-Raphson iteration method to calculate the interface fatigue damage model and obtain the node displacement vector of the aviation composite material structure at the current time step"; 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; The calculation formula for the 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; 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.

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: ； 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.

4. The method for analyzing fatigue failure process of aviation composite material structure according to claim 1, characterized in that: 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; Among them, the calculation formula of the fatigue damage factor in the fatigue damage degree calculation formula is: ； 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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