An aircraft composite material defect simulation method based on finite element analysis

Abstract

The application provides a simulation method for composite material defects of an airplane based on finite element analysis, and relates to the technical field of material defect simulation. The method comprises the following steps: S1, setting an analysis system, wherein the analysis system 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's modulus; S5, performing mesh division on the geometric model, and decomposing the structure into finite elements; S6, setting boundary conditions of the simulation geometric model, including constraints and loads; S7, running simulation, and obtaining the response of the geometric model under the given loading condition; and S8, analyzing the running result, and checking the influence of defects in the structure. Compared with the prior art, the method can accurately simulate the influence of the composite material defects in the airplane structure, and is helpful to improve the performance of the airplane structure.

Description

TECHNICAL FIELD

[0001] The present application relates to the technical field of material defect simulation, in particular to a simulation method for defects of composite materials of an airplane based on finite element analysis. BACKGROUND

[0002] Composite materials are widely used in aerospace industry, automobile manufacturing and wind power generation due to their high heat resistance, strength, light weight, low thermal expansion coefficient, small heat capacity and specific gravity. The performance of composite materials can reduce the weight of the aircraft structure, reduce air resistance, save fuel and is widely used in general aviation.

[0003] However, during the preparation and service of composite materials, internal defects such as pores, debonding, delamination and cracks may occur, which greatly reduces the safety performance of the composite materials and poses a major safety hazard to the application field. The existing artificial, penetration and eddy current detection methods are difficult to quickly, accurately and non-destructively find and locate the internal defects of the aircraft skin.

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

[0005] The purpose of the present application is to provide a simulation method for defects of composite materials of an airplane based on finite element analysis, which can accurately simulate the influence of composite material defects in the aircraft structure and help improve the performance of the aircraft structure.

[0006] The embodiments of the present application are implemented as follows:

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

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

[0009] Step S2: establishing a geometric model;

[0010] Step S3: introducing defects of composite materials in the geometric model;

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

[0012] Step S5: meshing the geometric model to decompose the structure into finite elements;

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

[0014] Step S7: running simulation, obtaining the response of the geometric model under the given loading condition;

[0015] Step S8: analyzing the running result, checking the defect influence in the structure.

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

[0017] In some embodiments of the present application, the step S2 specifically comprises the following steps:

[0018] Step S2.1: simulating the plate in the DM interface and performing composite material assignment.

[0019] Step S2.2: respectively simulating the laminated plate structure and the sandwich structure, the laminated plate structure is simulated for three layers, and the thickness of each layer is 5-10 mm; the sandwich structure is also simulated for three layers, the thickness of the upper layer and the lower layer is 5-10 mm, and the thickness of the middle layer structure is 10-20 mm.

[0020] In some embodiments of the present application, in the step S3, simulation defects are introduced in the simulation of the laminated plate structure and the sandwich structure, and the simulation defects are added to the middle layer of the two simulation models.

[0021] In some embodiments of the present application, in the step 5, the laminated plate structure is automatically divided, the whole sandwich structure is automatically divided, and the middle layer is modified by encryption for 1-5 times.

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

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

[0024] The simulation method for aircraft composite material defects based on finite element analysis provided by the application comprises the following steps: step S1: setting an analysis system, the analysis system being statics, dynamics or thermal analysis; step S2: establishing a geometric model; step S3: introducing defects of the composite material in the geometric model; step S4: defining material parameters, including elastic modulus, density and Young's modulus; step S5: performing meshing on 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 simulation, and obtaining responses of the geometric model under given loading conditions; and step S8: analyzing the running results, and checking the influence of defects in the structure.

[0025] The application combines finite element analysis with composite material simulation to provide a system specially used for simulation of aircraft composite material defects. By accurately considering anisotropy and defect characteristics of the composite material, the system can provide more accurate tools for design, evaluation and optimization of aircraft structures. The system is unique in that it can simulate the influence of composite material defects in aircraft structures, which helps engineers better understand and improve the performance of aircraft structures. BRIEF DESCRIPTION OF DRAWINGS

[0026] In order to more clearly illustrate the technical solutions of the embodiments of the application, the following will briefly introduce the drawings needed to be used in the embodiments. It should be understood that the following drawings only show some of the embodiments of the application, and therefore should not be regarded as a limitation on the scope. For those skilled in the art, other related drawings can also be obtained without creative labor on the basis of these drawings.

[0027] Figure 1 The step flow chart of the embodiments of the application;

[0028] Figure 2 The simulation schematic diagram of the sandwich structure of the embodiments of the application;

[0029] Figure 3 The simulation schematic diagram of the sandwich structure of the embodiments of the application;

[0030] Figure 4 The simulation schematic diagram of the sandwich structure of the embodiments of the application;

[0031] Figure 5 The simulation schematic diagram of the sandwich structure of the embodiments of the application; DETAILED DESCRIPTION

[0032] In order to make the objects, technical solutions and advantages of the embodiments of the present application clearer, the technical solutions in the embodiments of the present application will be described clearly and completely below with reference to the drawings in the embodiments of the present application. Obviously, the described embodiments are only some but not all of the embodiments of the present application. The components of the embodiments of the present application described and shown in the drawings herein can be arranged and designed in various different configurations.

[0033] Therefore, the following detailed description of the embodiments of the present application provided in the drawings is not intended to limit the scope of the claimed present application, but only represents selected embodiments of the present application. Based on the embodiments in the present application, all other embodiments obtained by those of ordinary skill in the art without creative work under the premise that the scope of protection of the present application belongs to.

[0034] Embodiments:

[0035] Please refer to Figures 1-5 The present embodiment provides a simulation method for defects of composite materials of an airplane based on finite element analysis, which comprises the following steps:

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

[0037] Step S2: establishing a geometric model;

[0038] Step S3: introducing defects of composite materials in the geometric model;

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

[0040] Step S5: meshing the geometric model, 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, obtaining the response of the geometric model under the given loading condition;

[0043] Step S8: analyzing the running results, checking the influence of defects in the structure.

[0044] In step S1, the heat affected zone is studied, so thermal analysis is selected. This analysis is beneficial to analyze the heat distribution of the aircraft in flight and improve the safety and stability of flight. In step S2, first, the ANSYS Workbench 2022R1 software is started, and in the "analysis system", "transient heat" is selected. Select DesignModeler to enter the simulation interface. When simulating, the unit is unified to mm. In the DM interface, first simulate a plate. Click xy plane, click front view, click "sketch drawing", click rectangle, and the surface area is 90000 square millimeters. Then simulate the laminated plate structure and the sandwich structure respectively. The simulation steps are similar to the above. The laminated plate structure is simulated for three layers, and the thickness of each layer is 10mm. The sandwich structure is also simulated for three layers. The thickness of the upper and lower layers is 10mm, and the thickness of the middle layer is 20mm.

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

[0046] In step S4, a cylinder is also simulated as a defect in the middle layer of the sandwich structure. Similarly, a new coordinate plane needs to be established in the middle layer. The origin of the coordinate plane needs to be set to the center of mass of the middle structure. Similarly, a circular sketch is drawn on the new plane. Finally, the depth of the extrusion is set to 5mm to generate the model and obtain the internal defect.

[0047] In step S5, the simulation system is first automatically divided into meshes, and then encryption correction is performed. Specifically, the laminated plate structure is automatically divided, the sandwich structure is automatically divided as a whole, and the middle layer is encrypted and corrected, with 1-3 corrections respectively.

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

[0049] The present application 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 lies in its ability to simulate the influence of composite material defects in aircraft structures, which helps engineers better understand and improve the performance of aircraft structures.

[0050] In summary, the embodiments of the present application provide a simulation method for defects in aircraft composite materials based on finite element analysis, comprising the following steps: step S1: setting an analysis system, the analysis system is statics, dynamics or thermal analysis; step S2: establishing a geometric model; step S3: introducing defects in the composite material in the geometric model; step S4: defining material parameters, including elastic modulus, density and Young's modulus; step S5: meshing the geometric model to decompose the structure into finite elements; step S6: setting the 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 condition; step S8: analyzing the running results to check the influence of the defects in the structure. The present application 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 lies in its ability to simulate the influence of composite material defects in aircraft structures, which helps engineers better understand and improve the performance of aircraft structures.

[0051] The above is only the preferred embodiment of the present application and is not intended to limit the present application. For those skilled in the art, the present application can have various modifications and changes. Any modification, equivalent replacement, improvement, etc. made within the spirit and principles of the present application shall be included in the protection scope of the present application.

Claims

1. A method of simulating defects in composite materials of an aircraft based on finite element analysis, characterized in that, The method comprises the following steps: Step S1: setting an analysis system, which is statics, dynamics or thermal analysis; Step S2: establishing a geometric model; Step S3: introducing defects of the composite material in the geometric model; Step S4: defining material parameters, including elastic modulus, density and Young's modulus; Step S5: meshing the geometric model to decompose 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 condition; Step S8: analyzing the running result to check the influence of the defects in the structure; In the step S1, the analysis system is selected as transient heat for thermal analysis; The step S2 specifically comprises the following steps: Step S2.1: simulating the plate in the DM interface and performing composite material assignment; Step S2.2: respectively simulating the laminated plate structure and the sandwich structure, the laminated plate structure is simulated for three layers, and the thickness of each layer is 5-10 mm, and the sandwich structure is also simulated for three layers, the thickness of the upper layer and the lower layer is 5-10 mm, and the thickness of the middle layer structure is 10-20 mm; In the step S3, simulation defects are introduced in the simulation of the laminated plate structure and the sandwich structure, and the simulation defects are added to the middle layer of the two simulation models; In the step S6, boundary conditions of the simulation geometric model are set, and 1*10 6 -2*10 6 external forces of pa are applied to the upper layer, the lower layer and the middle layer, respectively. 6 -2*10 6 external forces of pa are applied to the upper layer, the lower layer and the middle layer, respectively.

2. The method for simulation of defects in composite materials of an aircraft based on finite element analysis as claimed in claim 1, wherein, In the step S5, the laminated plate structure is automatically divided, the whole sandwich structure is automatically divided, and the middle layer is modified by encryption, and the modification is performed for 1-5 times respectively.

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