Method for determining ultimate load after patch repair of composite laminate

Abstract

The present invention discloses a method for determining the ultimate load after patch repair of a composite laminate, comprising the following steps: establishing a finite element model of an open-hole composite laminate; establishing a constitutive model for composite material damage; establishing a constitutive model for the composite adhesive layer; based on the ABAQUS-UMAT finite element user dynamic subroutine module, using FORTRAN language to write a user-defined subroutine to implement the proposed constitutive model for damage, and solving for stress, strain and damage; performing calculations on the finite element model to predict the ultimate load after patch repair of the composite laminate. The present invention uses the ABAQUS-UMAT user-defined subroutine to numerically implement the established three-dimensional constitutive model for damage, which simultaneously considers the effects of shear nonlinearity and degradation of material properties caused by damage accumulation, and can accurately predict the ultimate load after patch repair of the composite laminate.

Description

Technical Field

[0001] The present invention relates to a method for determining the ultimate load after patch repair of a composite laminate, belonging to the technical field of mechanical property analysis of composites. Background Art

[0002] In the aviation field, the application of composites in the structures of aircraft has been very common. Fiber-reinforced resin matrix composites have high specific modulus and specific strength, excellent energy absorption performance, especially the characteristics of designability of stiffness and strength in all directions, and are widely used in engineering fields such as aerospace, military, marine, civil, and mechanical. Although composites have various advantages compared with metal materials, they also have their disadvantages. In practical applications, it is found that due to the tissue characteristics of composites, the damage to the material after low-velocity impact often hides inside the structure and is not easily detectable by the naked eye. Due to highly integrated and large-sized components, replacing the damaged parts is not a good solution. Therefore, the demand for repair technologies for the main structural parts of aircraft (such as the fuselage or wing) has increased significantly compared with the replacement scheme of components. However, for those not very serious damages, the cost of replacing the structure is extremely high, the cost performance is low, and this scheme is not practical. Thus, the demand for developing repair technologies and repair processes for composite components is very urgent.

[0003] The most commonly used composite repair methods in the aerospace industry are patch repair and patch excision repair. These two repair technologies are different in processing and application. Due to the drive of new aircraft projects, the use of composites has entered the safety-critical primary structure, and the design and certification of repair methods have become more challenging. On the other hand, the structural performance and quality of structures repaired using glued structures depend not only on the bonding process but also on the experience and skills of composite repair technicians. Therefore, there is an urgent need for repair technologies and online monitoring technologies certified by the Civil Aviation Administration. In summary, it is of great significance to conduct simulation research on various repair models of composites using finite element software. Summary of the Invention

[0004] The purpose of the present invention is to overcome the technical defects existing in the prior art, solve the above technical problems, and propose a method for determining the ultimate load after patch repair of a composite laminate.

[0005] The present invention specifically adopts the following technical solutions: A method for determining the ultimate load after patch repair of a composite laminate, comprising the following steps:

[0006] Step SS1: Establish a finite element model of an open-hole composite laminate;

[0007] Step SS2: Establish a composite damage constitutive model;

[0008] Step SS3: Establish the constitutive model of the composite material adhesive layer;

[0009] Step SS4: Based on the ABAQUS-UMAT finite element user dynamic subroutine module, use FORTRAN language to write a user-defined subroutine to implement the proposed damage constitutive model, and solve for stress, strain, and damage;

[0010] Step SS5: Calculate the finite element model in Step SS1 to predict the ultimate load after patch repair of the composite laminate.

[0011] As a preferred embodiment, the specific steps of Step SS1 include:

[0012] The ply angles of the composite laminate are symmetrically arranged with respect to the mid-plane in the thickness direction, and only one element is divided in the thickness direction of each layer;

[0013] The mesh type is C3D8R, and the area around the hole is refined;

[0014] Establish the displacement consistency constraint condition in the loading direction between the reference point and the free end face: for tensile loads, the displacement loading method is adopted. A fixed support constraint is applied to the left loading surface, a reference point is set outside the right free end face, and then the reference point and the end face are bonded. In the Abaqus / CAE module, the creat constraint method is used to establish the coupling constraint equation. At this time, the displacement load is applied to the reference point, and as long as the displacement and reaction force on the reference point, namely U and RF1, are output, the displacement and reaction force on the loading end face can be obtained.

[0015] As a preferred embodiment, the specific steps of Step SS2 include:

[0016] Step SS21: Establish the constitutive relationship of the damaged composite laminate;

[0017] Step SS22: Establish the three-dimensional Hashin strength failure criterion to judge fiber and matrix damage, and establish the Ye delamination failure criterion to judge delamination damage;

[0018] Step SS23: Establish a shear nonlinear model;

[0019] Step SS24: Establish a continuous damage degradation model.

[0020] As a preferred embodiment, the specific steps of Step SS21 include:

[0021] The composite material stress-strain constitutive equation is as follows: σ = C(d):ε e ,

[0022] Among them, the symbol ":" represents the contraction calculation of two tensor indices; σ is the effective stress tensor; is the nominal stress tensor; ε e is the elastic strain tensor; e represents elasticity; C(d) is the fourth-order stiffness tensor of the damaged unidirectional composite laminate; C is the fourth-order linear elastic stiffness tensor of the undamaged unidirectional composite laminate; d is a one-dimensional vector (d1, d2, d3, d 23 , d 13 , d 12 ), where d1, d2, and d3 are the damage variables of fiber damage in the fiber direction, the damage variables of matrix damage perpendicular to the fiber direction in the plane, and the damage variables of delamination damage in the out-of-plane direction between layers, respectively; d 12 , d 23 , d 13 are the shear damage variables in the 12, 23, and 13 planes, respectively; the coordinate system x1-x2-x3 is defined as the natural coordinate system of the unidirectional plate, and x1-x n -x1 is the local coordinate system of the fracture surface, and the x1 axes in the two coordinate systems coincide; the 12, 23, and 13 planes correspond to the x1x2 plane, x2x3 plane, and x1x3 plane in the coordinate system, respectively;

[0023] Introducing the damage variable into the stiffness matrix makes the stiffness gradually weaken as the damage develops, that is:

[0024] C(d) = M -1 (d):C:M T，-1 (d);

[0025] Among them, M -1 (d) is the inverse matrix of M(d), and M T，-1 (d) is the inverse matrix of the transposed matrix of M(d); M(d) is the damage factor tensor, and its matrix form in the damage principal axis system can be expressed as follows:

[0026]

[0027]

[0028] The three-dimensional orthogonal anisotropic damage constitutive model of the monolayer plate in the composite material principal coordinate system is as follows:

[0029]

[0030] The composite material principal coordinate system is the natural coordinate system x1-x2-x3 of the unidirectional plate;

[0031] Among them:

[0032]

[0033]

[0034]

[0035]

[0036]

[0037]

[0038]

[0039]

[0040]

[0041]

[0042] Among them, σ1, σ2, and σ3 are the nominal normal stresses in the fiber direction, perpendicular to the fiber direction, and out-of-plane direction of the layer, respectively; τ 23 , τ 12 and τ 13 are the shear stresses in the x1x2 plane, x2x3 plane, and x1x3 plane in the x1 - x2 - x3 coordinate system, respectively; ε1, ε2, and ε3 are the engineering normal strains in the fiber direction, perpendicular to the fiber direction, and out-of-plane direction of the layer, respectively; γ 23 , γ 13 and γ 12 are the engineering shear strains in the x1x2 plane, x2x3 plane, and x1x3 plane in the x1 - x2 - x3 coordinate system, respectively; E1, E2, E3 are the elastic moduli of the undamaged unidirectional composite material monolayer in the fiber direction, perpendicular to the fiber direction, and out-of-plane direction of the layer, respectively, G 23 , G 13 , G 12 are the shear moduli of the undamaged unidirectional composite material monolayer in the x1x2 plane, x2x3 plane, and x1x3 plane, respectively, v 12 , v 13 , v 23 are the Poisson's ratios between the fiber direction and the direction perpendicular to the fiber direction, between the fiber direction and the out-of-plane direction of the layer, and between the direction perpendicular to the fiber direction and the out-of-plane direction of the layer, respectively, v 21 , v 31 , v 32 are the Poisson's ratios between the direction perpendicular to the fiber direction and the fiber direction, between the out-of-plane direction of the layer and the fiber direction, and between the out-of-plane direction of the layer and the direction perpendicular to the fiber direction, respectively, satisfying the relationship:

[0043] As a preferred embodiment, the specific establishment method for judging delamination damage by using the three-dimensional Hashin strength failure criterion and the Ye delamination failure criterion in the step SS22 is as follows:

[0044] (a) For fiber tension and compression, the initial damage criterion is:

[0045] Fiber tensile failure (ε 11 ≥0):

[0046]

[0047] Fiber compressive failure (ε 11 <0):

[0048]

[0049] (b) For matrix tension and compression, the initial damage criterion is:

[0050] Matrix tensile failure (ε 22 +ε 33 ≥0):

[0051]

[0052] Matrix compressive failure (ε 22 +ε 33 ≥0):

[0053]

[0054] (c) Initial delamination damage criterion:

[0055] Delamination failure caused by tension (ε 33 ≥0):

[0056]

[0057] Delamination failure caused by compression (ε 33 <0):

[0058]

[0059]

[0060] In the formula: f i (i = 1, 2, 3) respectively represent the damage states of the fiber, matrix, and delamination; C ii represents the stiffness coefficient of the material; respectively represent the normal strains corresponding to the tensile strength and compressive strength of the fiber in the i direction; γ 12 、γ 13 、γ 23Respectively represent the shear strains corresponding to the plane shear strengths; X T 、X C are the tensile and compressive strengths of the unidirectional plate along the fiber direction respectively; Y T 、Y C are the transverse tensile and compressive strengths of the unidirectional plate respectively; Z T is the normal tensile strength; S 12 、S 13 、S 23 are the shear strengths of the corresponding planes.

[0061] As a preferred embodiment, the specific establishment method of the shear nonlinear model in the step SS23 is as follows:

[0062]

[0063] The expression of the shear modulus G considering shear nonlinearity is:

[0064]

[0065] In the formula, τ and γ are the shear strain and shear stress respectively, G0 is the initial shear modulus, τ0 is the ultimate shear strength, and n is a parameter defining the shape of the shear nonlinear relationship curve.

[0066] As a preferred embodiment, the specific establishment method of the continuous damage degradation model in the step SS24 is as follows:

[0067]

[0068]

[0069]

[0070] Among them: L C is the element characteristic length, determined by mesh division, are the fracture energy dissipation rates in the three material principal directions respectively.

[0071] As a preferred embodiment, the step SS3 specifically includes:

[0072] Step SS31: Establish the constitutive equation of the adhesive layer, and the specific establishment method is as follows:

[0073] Taking the COH3D8 element as an example, the upper top surface and the lower bottom surface can be divided into four groups of node groups that can be separated from each other: 1-5, 2-6, 3-7, 4-8; each node has three degrees of freedom in three directions, so each group of nodes will generate a normal relative displacement δ n and two in-plane tangential displacement components δ s and δ t; Similarly, the cohesive force of the cohesive unit also has three components t that are the same as the displacement n 、t s and t t ;

[0074] Then the constitutive relation of the adhesive layer is obtained as follows:

[0075]

[0076] Where: K ii (i = n, s, t) is the stiffness coefficient;

[0077] Step SS32: Establish the strength failure criterion of the adhesive layer. The specific establishment method is as follows:

[0078] Suppose the interfacial strengths of the adhesive layer in the three failure modes I, II, and III are Under the action of the load in a single mode, the external load needs to reach the interfacial strength of the corresponding mode to cause failure; The quadratic strength criterion based on the relative separation displacement is used as the failure criterion of the adhesive layer:

[0079]

[0080]

[0081] Where are all adhesive layer strength coefficients;

[0082] Step SS33: Establish the performance degradation criterion of the adhesive layer. The specific establishment method is as follows:

[0083] The degradation model of the adhesive layer defines the degradation method of the material properties at the integration point when the integration point of the adhesive layer satisfies the failure criterion; In the separate crack propagation modes of I, II, and III, the strain energy release rate at the integration point needs to satisfy the critical strain energy release rate to carry out the corresponding crack propagation; The quadratic energy criterion is used as the damage degradation model of the adhesive layer:

[0084]

[0085] In the formula, G 1C 、G 2C 、G 3C are the critical strain energy release rates of the three modes respectively.

[0086] As a preferred embodiment, the step SS4 specifically includes:

[0087] Step SS41: Start the current increment step, read the convergence state variables at the previous moment and the strain increment in the current increment step, and update the strain and effective stress;

[0088] Step SS42: Substitute the effective stress into steps SS22 and SS23 of step SS2 to determine whether damage occurs. If damage occurs, update the damage variable through step SS24 of step SS2, and then calculate the nominal stress based on the effective stress and the damage variable.

[0089] As a preferred embodiment, step SS5 specifically includes: combining the finite element model file of the composite laminate established in step SS1 and the ABAQUS-UMAT user subroutine established in step SS4 to complete the prediction of the failure strength of the composite laminate; first establishing a finite element model of the composite laminate in the ABAQUS software, then calling the written subroutine for stress-strain analysis, and finally the load-displacement curve obtained is the mechanical behavior response of the model, and the maximum value obtained is the ultimate load.

[0090] The beneficial effects achieved by the present invention are as follows: First, for the two-dimensional elastic damage constitutive model based on the two-dimensional Hashin failure criterion embedded in ABAQUS, the present invention is based on a more accurate three-dimensional Hashin failure criterion and Ye delamination failure criterion continuous degradation model, adhesive layer failure criterion and degradation model, as well as a three-dimensional damage constitutive model including shear nonlinear effects, which is more in line with engineering practice. Second, the present invention uses the ABAQUS-UMAT user-defined subroutine to numerically implement the established three-dimensional damage constitutive model, which has higher calculation efficiency and calculation accuracy. Description of the Drawings

[0091] Figure 1 is a flowchart of a method for determining the ultimate load of a composite laminate after patch repair according to the present invention;

[0092] Figure 2 is a flowchart of the subroutine of the ABAQUS user material;

[0093] Figure 3 Finite element geometry of the damaged test piece. Detailed Embodiments

[0094] The present invention will be further described below with reference to the drawings. The following embodiments are only used to more clearly illustrate the technical solutions of the present invention and cannot be used to limit the protection scope of the present invention.

[0095] The present invention will be further described in detail below with reference to the embodiments and the drawings, but the embodiments of the present invention are not limited thereto.

[0096] Establish a finite element model in ABAQUS. Establish a geometric model as Figure 3As shown in the figure, where: W = 20 mm is the width of the laminate, L = 120 mm is the length of the laminate, T = 2 mm is the thickness of the laminate, t = 0.1 mm is the thickness of the adhesive layer, a = 20 mm is the length of the reinforcement patch, D is the diameter of the damage hole, and B is the diameter of the patch. The parameter values of D and B are shown in Table 1, and the material properties are shown in Tables 2, 3, and 4.

[0097] Table 1 Geometric Dimension Parameter Table of T700SC / EC240A Material Test Specimens

[0098]

[0099] Table 2 Elastic Coefficient Table of T700SC / EC240A Composite Materials

[0100]

[0101]

[0102] Table 3 Strength Parameter Table of T700SC / EC240A Composite Materials

[0103]

[0104] Table 4 Material Parameter Table of J-159 Adhesive Layer

[0105]

[0106] In this subsection, for the composite laminate part, the UMAT subroutine with continuous damage degradation based on the three-dimensional Hashin-Ye failure criterion established in Chapter 3 is used. For the adhesive layer part, the quadratic strength failure criterion and the quadratic energy degradation model of ABAQUS are used.

[0107] The elements of the composite laminate part are C3D8R solid elements, and the elements of the adhesive layer part are COH3D8R cohesive elements. The force is transferred between the adhesive layer and the mother plate, and between the adhesive layer and the patch by the way of co-nodal coupling. The boundary conditions of the model are set as follows: one end is fixed, and the other end uses the coupling method of the reference point and the surface. The displacement load is applied to the reference point, and finally the load-displacement curve of the reference point is output.

[0108] The load is applied in the form of displacement load. The displacement load should be appropriate. If the load is too small, the failure strength limit of the material cannot be reached, and the failure starting point and the final failure load of the material cannot be determined. If the load is too large, it may lead to non-convergence of the calculation. In this example, a displacement load of 2 mm is planned to be applied to the model.

[0109] The tensile failure process of the composite laminate is calculated and simulated using ABAQUS / STANDRAD. The user subroutine UMAT is used to read the current strain increment, update the strain and effective stress, determine whether the element enters the damaged state based on the effective stress, and calculate the damage variable according to the damage evolution model when entering the damage stage, so as to obtain the nominal stress. Finally, the load-displacement curve of the model can be obtained.

[0110] Table 5 Analysis Table of Simulation and Test Errors of T700SC / EC240A Material Laminate

[0111]

[0112] Table 5 is the analysis table of simulation and test errors of the ultimate load of the T700SC / EC240A material laminate after patch repair. It can be seen from Table 5 that the error between the simulation results and the test results in this paper is small.

[0113] The present invention is the development of a user subroutine based on the ABAQUS software, proposing a three-dimensional damage constitutive model and an adhesive layer constitutive model. Considering both the shear nonlinear effect and the influence of material property degradation caused by damage accumulation, it can ideally predict the failure strength of the composite laminate, providing technical support for deeply clarifying the damage failure characteristics of composite structures and improving the lightweight strength design level.

[0114] The above are only the preferred embodiments of the present invention. It should be noted that for those of ordinary skill in the art, without departing from the technical principle of the present invention, several improvements and deformations can be made, and these improvements and deformations should also be regarded as the protection scope of the present invention.

Claims

1. A method for determining the ultimate load after patch repair of a composite laminate, characterized in that, It includes the following steps: Step SS1: Establish a finite element model of the composite laminate with an opening; Step SS2: Establish a composite damage constitutive model, specifically including: Step SS21: Establish the constitutive relationship of the damaged composite laminate, specifically including: The constitutive equation of stress and strain for the composite material is as follows: σ = C(d):ε e , Among them, the symbol ":" represents the contraction calculation of two tensor indices; σ is the effective stress tensor; is the nominal stress tensor; ε e is the elastic strain tensor; e represents elasticity; C(d) is the fourth-order stiffness tensor of the damaged unidirectional composite laminate; C is the fourth-order linear elastic stiffness tensor of the undamaged unidirectional composite laminate; d is a one-dimensional vector (d1, d2, d3, d 23 , d 13 , d 12 ), where d1, d2, and d3 are the damage variables of fiber damage in the fiber direction, the damage variables of matrix damage perpendicular to the fiber direction in the plane, and the damage variables of delamination damage in the out-of-plane direction between layers, respectively; d 12 , d 23 , d 13 are the shear damage variables in the 12, 23, and 13 planes, respectively; the coordinate system x1x2x3 is defined as the natural coordinate system of the unidirectional plate, and x1x n x1 is the local coordinate system of the fracture surface, and the x1 axes in the two coordinate systems coincide; the 12, 23, and 13 planes correspond to the x1x2 plane, the x2x3 plane, and the x1x3 plane in the coordinate system, respectively; Introduce the damage variable into the stiffness matrix to make the stiffness gradually weaken as the damage develops, that is: C(d) = M -1 (d): C: M -,-1 (d); Among them, M -1 (d) is the inverse matrix of M(d), and M T,-1 (d) is the inverse matrix of the transposed matrix of M(d); M(d) is the damage factor tensor, and its matrix form in the damage principal axis system can be expressed as follows: The three-dimensional orthotropic damage constitutive model of a single ply in the principal coordinate system of the composite material is as follows: The principal coordinate system of the composite material is the natural coordinate system x1x2x3 of the unidirectional plate; Where: where σ1, σ2, and σ3 are the nominal normal stresses in the fiber direction, perpendicular to the fiber direction, and out-of-plane of the layer plane, respectively; τ 23 , τ 12 , and τ 13 are the shear stresses in the x1x2 plane, x2x3 plane, and x1x3 plane in the x1x2x3 coordinate system, respectively; ε1, ε2, and ε3 are the engineering normal strains in the fiber direction, perpendicular to the fiber direction, and out-of-plane of the layer plane, respectively; γ 23 , γ 13 , and γ 12 are the engineering shear strains in the x1x2 plane, x2x3 plane, and x1x3 plane in the x1x2x3 coordinate system, respectively; E1, E2, and E3 are the elastic moduli of the undamaged unidirectional composite ply in the fiber direction, perpendicular to the fiber direction, and out-of-plane of the layer plane, respectively, G 23 , G 13 , G 12 are the shear moduli of the undamaged unidirectional composite ply in the x1x2 plane, x2x3 plane, and x1x3 plane, respectively, v 12 , v 13 , v 23 are the Poisson's ratios between the fiber direction and perpendicular to the fiber direction, between the fiber direction and out-of-plane of the layer plane, and between perpendicular to the fiber direction and out-of-plane of the layer plane, respectively, v 21 , v 31 , v 32 are the Poisson's ratios between perpendicular to the fiber direction and the fiber direction, between out-of-plane of the layer plane and the fiber direction, and between out-of-plane of the layer plane and perpendicular to the fiber direction, respectively, satisfying the relationship: Step SS22: Establish a three-dimensional Hashin strength failure criterion to judge fiber and matrix damage, and establish a Ye delamination failure criterion to judge delamination damage; The specific establishment methods of the three-dimensional Hashin strength failure criterion and the Ye delamination failure criterion for judging delamination damage in Step SS22 are: (a) For fiber tension and compression, the initial damage criterion is: Fiber tensile failure (ε 11 ≥ 0): Fiber compression failure (ε 11 <0): (b) For matrix tension and compression, the initial damage criterion is: Matrix tensile failure (ε 22 + ε 33 ≥ 0): Matrix compression failure (ε 22 + ε 33 ≥ 0): (c) Initial delamination damage criterion: Tensile-induced delamination failure (ε 33 ≥ 0): Compression-induced delamination failure (ε 33 <0): where: f i (i = 1, 2, 3) represent the damage states of the fiber, matrix, and delamination respectively; C ii represents the stiffness coefficient of the material; respectively represent the positive strains corresponding to the tensile strength and compressive strength of the fiber in the i direction; γ 12 、γ 13 、γ 23 respectively represent the shear strains corresponding to the shear strengths of each plane; X T 、X C are respectively the tensile and compressive strengths of the unidirectional laminate along the fiber direction; Y T 、Y C are respectively the transverse tensile and compressive strengths of the unidirectional laminate; Z T is the normal tensile strength; S 12 、S 13 、S 23 are the shear strengths of the corresponding planes; Step SS23: Establish a shear nonlinear model: The specific establishment method of the shear nonlinear model is: The expression of the shear modulus G considering shear nonlinearity is: In the formula, τ and γ are the shear strain and shear stress respectively, G0 is the initial shear modulus, τ0 is the ultimate shear strength, and n is a parameter defining the shape of the shear nonlinear relationship curve; Step SS24: Establish a continuous damage degradation model; The specific establishment method of the continuous damage degradation model is: Where: L C is the unit characteristic length, determined by mesh division, are the fracture energy dissipation rates in the three main material directions respectively; Step SS3: Establish a composite adhesive layer constitutive model; Step SS4: Based on the ABAQUS UMAT finite element user-defined dynamic subroutine module, use FORTRAN language to write a user-defined subroutine to implement the proposed damage constitutive model, and solve for stress, strain and damage; Step SS5: Calculate the finite element model in Step SS1 to predict the ultimate load after patch repair of the composite laminate.

2. A method for determining the ultimate load after patch repair of a composite laminate according to claim 1, characterized in that The specific content of Step SS1 includes: The ply angles of the composite laminate are symmetrically arranged with respect to the mid-plane in the thickness direction, and only one element is divided in the thickness direction of each layer; The mesh type is C3D8R, and the area around the hole is refined; Establish a displacement consistency constraint condition between the reference point and the free end face: For tensile loads, the displacement loading method is adopted. A fixed support constraint is applied to the left loading surface, a reference point is set outside the right free end face, and then the reference point and the end face are bonded. In the Abaqus / CAE module, use the creat constraint method to establish a coupling constraint equation. At this time, apply the displacement load to the reference point, and as long as the displacement and reaction force on the reference point, namely U and RF1, are output, the displacement and reaction force on the loading end face can be obtained.

3. The method for determining the ultimate load after patch repair of a composite laminate according to claim 1, wherein The specific content of Step SS3 includes: Step SS31: Establish the adhesive layer constitutive equation, and the specific establishment method is: Taking the COH3D8 element as an example, the upper top surface and the lower bottom surface can be divided into four groups of node groups that can be separated from each other: 15, 26, 37, 48; each node has three degrees of freedom in three directions, so each group of nodes will generate a normal relative displacement δ n and two in-plane tangential displacement components δ s and δ t ; similarly, the cohesive force of the cohesive element also has three components t n 、t s and t t ; Then the constitutive relationship of the adhesive layer is obtained as follows: Where: K ii (i = n, s, t) is the stiffness coefficient; Step SS32: Establish the adhesive layer strength failure criterion, and the specific establishment method is: Let the interfacial strengths of the adhesive layer in three failure modes Ⅰ, Ⅱ, and Ⅲ be Under the action of the load in a single mode, the external load needs to reach the interfacial strength of the corresponding mode to cause failure; the quadratic strength criterion based on the relative separation displacement is used as the failure criterion for the adhesive layer: wherein are all adhesive layer strength coefficients; Step SS33: Establish the adhesive layer performance degradation criterion, and the specific establishment method is: The definition of the adhesive layer degradation model is the degradation method of the material properties at the integration point when the integration point of the adhesive layer meets the failure criterion; in the separate crack propagation modes of I, II, and III, the strain energy release rate at the integration point needs to meet the critical strain energy release rate to carry out the corresponding crack propagation; the quadratic energy criterion is used as the damage degradation model of the adhesive layer: where G 1C , G 2C , and G 3C are the critical strain energy release rates of the three modes, respectively.

4. The method for determining the ultimate load after patch repair of a composite laminate according to claim 1, characterized in that The specific steps of step SS4 include: Step SS41: Start the current increment step, read the convergence state variables at the previous moment and the strain increment in the current increment step, and update the strain and effective stress; Step SS42: Substitute the effective stress into steps SS22 and SS23 of step SS2 to judge whether damage occurs. If damage occurs, update the damage variable through step SS24 of step SS2, and then calculate the nominal stress through the effective stress and the damage variable.

5. The method for determining the ultimate load after patch repair of a composite material laminate according to claim 1, wherein The specific steps of step SS5 include: Combine the finite element model file of the composite laminate established in step SS1 and the ABAQUS UMAT user subroutine established in step SS4 to complete the prediction of the failure strength of the composite laminate; first establish the finite element model of the composite laminate in the ABAQUS software, then call the written subroutine for stress-strain analysis, and finally the load-displacement curve obtained is the mechanical behavior response of the model, and the maximum value obtained is the ultimate load.

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