Simulation method for aircraft composite material defects based on finite element analysis

Abstract

The invention provides an aircraft composite material defect simulation method based on finite element analysis, and relates to the technical field of material defect simulation. Comprising the following steps: S1, setting an analysis system which is statics, dynamics or thermal analysis; s2, establishing a geometric model; s3, introducing defects of the composite material into the geometric model; s4, defining material parameters including elastic modulus, density and Young modulus; s5, performing mesh generation on the geometric model, and decomposing the structure into finite elements; s6, setting boundary conditions of the simulation geometric model, wherein the boundary conditions comprise constraints and loads; s7, simulation is operated, and the response of the geometric model under the given loading condition is obtained; s8, analyzing an operation result, and checking defect influence in the structure; compared with the prior art, the influence of composite material defects in an aircraft structure can be accurately simulated, and the performance of the aircraft structure can be improved.

Description

Technical Field

[0001] The invention relates to the technical field of material defect simulation, and in particular to a method for simulating defects of aircraft composite materials based on finite element analysis. Background Art

[0002] Composite materials are widely used in aerospace, automotive, wind power and other manufacturing industries due to their high heat resistance and strength, light weight, low thermal expansion coefficient, small heat capacity and specific gravity. Composite materials can reduce the weight of aircraft structures, reduce air resistance, and save fuel, and are widely used in general aviation.

[0003] However, during the preparation and service process of composite materials, internal defects such as pores, debonding, delamination, and cracks will occur, which greatly reduce the safety performance of composite materials and pose a major safety hazard to the application field. Existing manual, penetration, eddy current and other detection methods are difficult to quickly, accurately and non-destructively detect and locate internal defects in aircraft skins.

[0004] Therefore, there is an urgent need for a simulation method of aircraft composite material defects based on finite element analysis to solve the above-mentioned problems. Summary of the invention

[0005] The object of the present invention is to provide a simulation method of aircraft composite material defects based on finite element analysis, which can accurately simulate the influence of composite material defects in aircraft structures and help improve the performance of aircraft structures.

[0006] The embodiment of the present invention is achieved as follows:

[0007] The embodiment of the present application provides a simulation method for aircraft composite material defects based on finite element analysis, which includes the following steps:

[0008] Step S1: Setting an analysis system, wherein the analysis system is statics, dynamics or thermal analysis;

[0009] Step S2: Establishing a geometric model;

[0010] Step S3: introducing defects of the composite material into the geometric model;

[0011] Step S4: defining material parameters, including elastic modulus, density and Young's modulus;

[0012] Step S5: meshing the geometric model and decomposing the structure into finite elements;

[0013] Step S6: setting boundary conditions of the simulation geometric model, including constraints and loads;

[0014] Step S7: running the simulation to obtain the response of the geometric model under the given loading conditions;

[0015] Step S8: Analyze the operation results and check the impact of defects in the structure.

[0016] In some embodiments of the present invention, in the above step S1, the analysis system selects transient heat to perform thermal analysis.

[0017] In some embodiments of the present invention, the above step S2 specifically includes the following steps:

[0018] Step S2.1: Simulate the plate at the DM interface and assign the composite material.

[0019] Step S2.2: The laminate structure and the sandwich structure are simulated respectively. The laminate structure is simulated with three layers, and the thickness of the three layers is 5-10 mm. The sandwich structure is also simulated with three layers, and the thickness of the upper and lower layers is 5-10 mm, and the thickness of the middle layer is 10-20 mm.

[0020] In some embodiments of the present invention, in the above step S3, simulation defects are introduced into the simulation of the laminate structure and the sandwich structure, and simulation defects are added to the middle layers of the two simulation models.

[0021] In some embodiments of the present invention, in the above step 5, the laminate structure is automatically divided, the sandwich structure as a whole is automatically divided, and the middle layer is encrypted and corrected 1 to 5 times respectively.

[0022] In some embodiments of the present invention, in the above step S6, the boundary conditions of the simulation geometric model are set to apply 1*10 6 -2*10 6 pa external force; 1*10 6 -2*10 6 The external force of pa is applied to the upper layer, lower layer and middle layer respectively.

[0023] Compared with the prior art, the embodiments of the present invention have at least the following advantages or beneficial effects:

[0024] The present invention provides a simulation method for aircraft composite material defects based on finite element analysis, comprising the following steps: step S1: setting an analysis system, wherein the analysis system is statics, dynamics or thermal analysis; step S2: establishing a geometric model; step S3: introducing defects of the composite material into the geometric model; step S4: defining material parameters, including elastic modulus, density and Young's modulus; step S5: meshing the geometric model and decomposing the structure into finite elements; step S6: setting boundary conditions of the simulated geometric model, including constraints and loads; step S7: running the simulation to obtain the response of the geometric model under given loading conditions; and step S8: analyzing the running results to check the influence of defects in the structure.

[0025] The present invention combines finite element analysis with composite material simulation to provide a system specifically for aircraft composite material defect simulation. By accurately accounting for the anisotropy and defect characteristics of composite materials, the system can provide more accurate tools for the design, evaluation and optimization of aircraft structures. The uniqueness of the system lies in its ability to simulate the effects of composite material defects in aircraft structures, helping engineers better understand and improve the performance of aircraft structures. BRIEF DESCRIPTION OF THE DRAWINGS

[0026] In order to more clearly illustrate the technical solutions of the embodiments of the present invention, the drawings required for use in the embodiments are briefly introduced below. It should be understood that the following drawings only show certain embodiments of the present invention and therefore should not be regarded as limiting the scope. For ordinary technicians in this field, other related drawings can be obtained based on these drawings without creative work.

[0027] Figure 1 is a flowchart of steps of an embodiment of the present invention;

[0028] Figure 2 It is a simulation schematic diagram of the laminate structure of an embodiment of the present invention;

[0029] Figure 3 A simulation schematic diagram of a sandwich structure according to an embodiment of the present invention;

[0030] Figure 4 It is a simulation schematic diagram of the structural defects of the laminated plate according to an embodiment of the present invention;

[0031] Figure 5 It is a simulation schematic diagram of the sandwich structure defect according to an embodiment of the present invention. DETAILED DESCRIPTION

[0032] In order to make the purpose, technical solutions and advantages of the embodiments of the present invention clearer, the technical solutions in the embodiments of the present invention will be clearly and completely described below in conjunction with the drawings in the embodiments of the present invention. Obviously, the described embodiments are part of the embodiments of the present invention, not all of the embodiments. Generally, the components of the embodiments of the present invention described and shown in the drawings here can be arranged and designed in various different configurations.

[0033] Therefore, the following detailed description of the embodiments of the present invention provided in the accompanying drawings is not intended to limit the scope of the invention claimed for protection, but merely represents selected embodiments of the present invention. Based on the embodiments of the present invention, all other embodiments obtained by ordinary technicians in this field without creative work are within the scope of protection of the present invention.

[0034] Example:

[0035] Please refer to Figure 1-5 This embodiment provides a simulation method for aircraft composite material defects based on finite element analysis, which includes the following steps:

[0036] Step S1: Setting an analysis system, wherein the analysis system is statics, dynamics or thermal analysis;

[0037] Step S2: Establishing a geometric model;

[0038] Step S3: introducing defects of the composite material into the geometric model;

[0039] Step S4: defining material parameters, including elastic modulus, density and Young's modulus;

[0040] Step S5: meshing the geometric model and decomposing the structure into finite elements;

[0041] Step S6: setting boundary conditions of the simulation geometric model, including constraints and loads;

[0042] Step S7: running the simulation to obtain the response of the geometric model under the given loading conditions;

[0043] Step S8: Analyze the operation results and check the impact of defects in the structure.

[0044] In step S1, the thermal impact is studied, so thermal analysis is selected. Selecting this analysis is conducive to analyzing the heat distribution of the aircraft in flight and improving flight safety and stability. In step S2, first start the ANSYS Workbench 2022R1 software, select "Transient Thermal" in "Analysis System", select DesignModeler to enter the simulation interface, and unify the unit to mm during simulation. First simulate a board in the DM interface, click the xy plane, click to face it, click "Sketch", click the rectangle, the surface area is 90,000 square millimeters, and then simulate the laminate structure and the sandwich structure respectively. The simulation steps are similar to the above. The laminate structure simulates three layers, and the thickness of the three layers is 10 mm. The sandwich structure also simulates three layers, and the thickness of the upper and lower layers is 10 mm, and the middle layer structure is 20 mm.

[0045] In step S3, a cylinder is simulated as a defect in the middle layer of the laminate structure. First, a new coordinate plane needs to be established in the middle layer, and then a circular sketch is simulated on the new plane. Finally, the extrusion depth is set to 2mm, and the model is generated to obtain the internal defects. In step S4, the material parameters are defined, including elastic modulus, density, Young's modulus, etc. After the DM interface simulation is completed, return to the operation interface and click "Engineering Data" to assign material parameters. Since this simulation is a composite material simulation, the parameters are all set with reference to the specific data of the composite material.

[0046] In step S4, a cylinder is also simulated as a defect in the middle layer of the sandwich structure. It is also necessary to establish a new coordinate plane in the middle layer. The establishment of the coordinate plane needs to take the center of mass of the middle structure as the origin. A circular sketch is also drawn on the new plane. Finally, the extrusion depth is set to 5 mm. The internal defect can be obtained by generating a model.

[0047] In step S5, the simulation system is first automatically meshed and then encrypted and corrected. Specifically, the laminate structure is automatically meshed, the sandwich structure as a whole is automatically meshed, and the middle layer is encrypted and corrected, and the corrections are performed 1-3 times respectively.

[0048] In step S6, the boundary conditions of the simulation geometric model are set, including constraints and loads; a 1*10 6 The external force of pa is 1*10 6The external force of pa is applied to the upper and lower layers, as well as the middle layer, respectively, to facilitate the comparison of the effects of external forces in different directions on defective materials during flight. In step S7, the simulation is run to obtain the response of the geometric model under given loading conditions. In step S8, the results are analyzed: the running results are analyzed to check the effects of defects in the structure; the heat distribution of different layers in the laminate structure and the sandwich structure are analyzed respectively, and the heat distribution of defects in the three cases is analyzed at the same time. For the composite plate, the total deformation is observed, and the equivalent stress can also be observed.

[0049] The present invention combines finite element analysis with composite material simulation to provide a system specifically for aircraft composite material defect simulation. By accurately accounting for the anisotropy and defect characteristics of composite materials, the system can provide more accurate tools for the design, evaluation and optimization of aircraft structures. The uniqueness of the system lies in its ability to simulate the effects of composite material defects in aircraft structures, helping engineers better understand and improve the performance of aircraft structures.

[0050] In summary, an embodiment of the present invention provides a simulation method for defects in aircraft composite materials based on finite element analysis, comprising the following steps: step S1: setting an analysis system, wherein the analysis system is static, dynamic or thermal analysis; step S2: establishing a geometric model; step S3: introducing defects in composite materials into the geometric model; step S4: defining material parameters, including elastic modulus, density and Young's modulus; step S5: meshing the geometric model and decomposing the structure into finite elements; step S6: setting boundary conditions of the simulated geometric model, including constraints and loads; step S7: running the simulation to obtain the response of the geometric model under given loading conditions; step S8: analyzing the running results to check the effect of defects in the structure. The present invention combines finite element analysis with composite material simulation to provide a system specifically for simulating defects in aircraft composite materials. By accurately considering the anisotropy and defect characteristics of composite materials, the system can provide more accurate tools for the design, evaluation and optimization of aircraft structures. The uniqueness of the system is that it can simulate the effects of composite material defects in aircraft structures, which helps engineers better understand and improve the performance of aircraft structures.

[0051] The above are only preferred embodiments of the present invention and are not intended to limit the present invention. For those skilled in the art, the present invention may have various modifications and variations. Any modification, equivalent replacement, improvement, etc. made within the spirit and principle of the present invention shall be included in the protection scope of the present invention.

Claims

1. A simulation method for aircraft composite material defects based on finite element analysis, characterized in that: The following steps are involved: Step S1: Setting an analysis system, wherein the analysis system is statics, dynamics or thermal analysis; Step S2: Establishing a geometric model; Step S3: introducing defects of the composite material into the geometric model; Step S4: defining material parameters, including elastic modulus, density and Young's modulus; Step S5: meshing the geometric model and decomposing the structure into finite elements; Step S6: setting boundary conditions of the simulation geometric model, including constraints and loads; Step S7: running the simulation to obtain the response of the geometric model under the given loading conditions; Step S8: Analyze the operation results and check the impact of defects in the structure.

2. The method for simulating defects of aircraft composite materials based on finite element analysis according to claim 1, characterized in that: In step S1, the analysis system selects transient heat to perform thermal analysis.

3. The method for simulating defects of aircraft composite materials based on finite element analysis according to claim 2, characterized in that: The step S2 specifically includes the following steps: Step S2.1: Simulate the plate at the DM interface and assign the composite material. Step S2.2: The laminate structure and the sandwich structure are simulated respectively. The laminate structure is simulated with three layers, and the thickness of the three layers is 5-10 mm. The sandwich structure is also simulated with three layers, and the thickness of the upper and lower layers is 5-10 mm, and the thickness of the middle layer is 10-20 mm.

4. The method for simulating defects of aircraft composite materials based on finite element analysis according to claim 3, characterized in that: In the step S3, simulation defects are introduced into the simulation of the laminate structure and the sandwich structure, and simulation defects are added to the middle layers of the two simulation models.

5. The method for simulating defects of composite aircraft materials based on finite element analysis according to claim 4, characterized in that: In the step S5, the laminate structure is automatically divided, the sandwich structure as a whole is automatically divided, and the middle layer is encrypted and corrected 1 to 5 times respectively.

6. The method for simulating defects of aircraft composite materials based on finite element analysis according to claim 5, characterized in that: In step S6, the boundary conditions of the simulation geometric model are set to apply 1*10 6 -2*10 6 pa external force; 1*10 6 -2*10 6 The external force of pa is applied to the upper layer, lower layer and middle layer respectively.

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