Method, equipment, medium and system for predicting mechanical properties of aircraft heat-proof composite material

Abstract

The invention discloses a method, a device, a medium and a system for predicting mechanical properties of a heat-proof composite material of an aircraft, and belongs to the field of numerical computation.The method comprises the steps that non-uniform grid division is conducted on grids of the heat-proof composite material of the aircraft, and divided areas comprise areas needing local encryption and areas without encryption; aiming at grid nodes at the grid division interface, sequentially connecting grid unit centers and edge midpoints to form a control body; according to the formed control body discrete balance equation, calculating and solving the balance equation based on midpoint integration; based on a balance equation solving result, the displacements on all the grid nodes are used as variables to be solved, a system matrix is assembled, and solving is carried out; and substituting the solved displacement into a stress, strain and displacement relational expression to obtain the strain and stress at the corresponding position so as to complete material mechanical property prediction. According to the method, the number of grids can be effectively controlled while the calculation precision is greatly improved.

Description

TECHNICAL FIELD

[0001] The present application relates to the field of numerical calculation, more particularly, to a method, device, medium and system for predicting mechanical properties of a thermal protection composite material of an aircraft. BACKGROUND

[0002] During long-time cruise flight, the surface of an aircraft often faces severe force and thermal load environment, which puts high requirements on the structure and material properties of the aircraft. Composite materials have many excellent properties such as high specific modulus, high specific strength, low thermal expansion coefficient, high temperature resistance, corrosion resistance and wear resistance, which provide good material support for the lightweight structure, thermal protection system and internal structure of the aircraft.

[0003] The structural properties of the thermal protection composite material are greatly affected by its microstructure characteristics (such as fiber layup method, particles, air inclusions, etc.). The uneven distribution of mechanical properties and insufficient strength of the material caused by the above microstructure characteristics are important factors affecting its large-scale application. Numerical simulation technology is an effective technical means for the characterization of composite material structural properties. Due to the influence of the microstructure characteristics of the composite material, a very fine structure-grid topology relationship needs to be used in the traditional numerical simulation process to effectively describe the influence law between the structural mechanical properties and the microstructure of the material. The numerical calculation amount is very large, therefore, a relatively coarse grid division size is often used to reduce the number of grids, thus sacrificing the calculation accuracy. In engineering use, the thickness / weight of the composite material is increased by designing a safety factor to achieve the ability to use safely, which is contrary to the starting point of using composite materials to reduce the weight of the missile.

[0004] Therefore, how to effectively and accurately describe the mechanical properties of the thermal protection composite material without significantly increasing the calculation amount, and then realize the optimal selection of the composite material structure, is the key content of the current application of the thermal protection composite material of the aircraft. SUMMARY

[0005] The present application aims to overcome the shortcomings of the prior art and provide a method, device, medium and system for predicting the mechanical properties of a thermal protection composite material of an aircraft, which greatly improves the calculation accuracy while effectively controlling the number of grids.

[0006] The purpose of the present application is achieved by the following scheme: A method for predicting the mechanical properties of a thermal protection composite material of an aircraft, comprising the following steps: The aircraft heat-resistant composite material grid is subjected to non-uniform grid division, and the divided regions include a local encryption region and a non-encryption region; the local encryption region is a fine grid and corresponds to a composite material microstructure feature position; and the non-encryption region is a coarse grid and corresponds to a non-microstructure feature position of the composite material. For the grid nodes at the grid division interface, the grid cell centers and the edge midpoints are sequentially connected to form control volumes. The balance equation is discretized according to the formed control volumes, and calculation is performed based on midpoint integration. The displacements of all the grid nodes are taken as variables to be solved, the system matrix is assembled, and the solution is obtained; the obtained displacement is brought into the stress-strain-displacement relationship to obtain the strain and stress at the corresponding position, thereby completing the prediction of the mechanical properties of the composite material.

[0007] Further, the aircraft heat-resistant composite material grid includes a grid composed of polygonal elements.

[0008] Further, the composite material microstructure features include fibers, particles and air inclusions.

[0009] Further, the grid cell centers and the edge midpoints are sequentially connected to form control volumes, and the CV-FVM control volume construction method is adopted to sequentially connect the grid cell centers and the edge midpoints to form control volumes.

[0010] Further, the grid composed of polygonal elements includes a mixed grid of quadrilateral elements and triangular elements.

[0011] Further, the grid composed of polygonal elements includes a mixed grid of three-dimensional tetrahedral elements and hexahedral elements.

[0012] An aircraft heat-resistant composite material mechanical property prediction device includes a processor and a memory, and the memory stores a computer program, which, when loaded by the processor, executes the method of any one of the above.

[0013] A computer readable storage medium stores a computer program, which, when loaded by a processor, executes the method of any one of the above.

[0014] An aircraft heat-resistant composite material mechanical property prediction system includes an aircraft heat-resistant composite material mechanical property prediction device as described above.

[0015] The beneficial effects of the present application include: (1) Firstly, the local grid encryption concept is introduced for mesh division, then the local encryption is carried out on the microstructure characteristics (fibers, particles, air inclusions, etc.) of the composite material, and the coarse mesh division is maintained at the position of non-microstructure characteristics, finally the combination of coarse and fine meshes is realized to greatly improve the calculation accuracy and effectively control the number of meshes, so as to achieve the fine description of the relationship between the structural mechanical properties and the microstructure characteristics, and further realize the low redundancy design based on the composite structure.

[0016] (2) The present application can realize the mechanical non-corresponding grid topology process between different structures / materials, and does not require that different structure grids must be one-to-one corresponding, thus reducing the difficulty of grid division, in addition, the method of the present application can use mixed grid of triangular element and quadrilateral element for division, further reducing the difficulty of grid division.

[0017] (3) The present application can realize the fine prediction ability of displacement, stress / strain response with local obvious microstructure characteristics, and at the same time does not need to encrypt transition treatment for the structure grid around the microstructure characteristics, thus ensuring the prediction accuracy without greatly increasing the calculation solving time. BRIEF DESCRIPTION OF DRAWINGS

[0018] In order to more clearly illustrate the technical solutions of the embodiments of the present application or the prior art, the following will briefly introduce the drawings needed to be used in the embodiment or prior art description, obviously, the drawings in the following description are only some embodiments of the present application, and for those skilled in the art, other drawings can also be obtained without creative labor.

[0019] Figure 1 A schematic diagram is established for the local encrypted grid point type finite volume method control volume; Figure 2 A control volume schematic diagram is established; Figure 3 A triangular element integration point schematic diagram under different coordinate systems is established; wherein (a) is a triangular element integration point schematic diagram under global coordinate system; (b) is a triangular element integration point schematic diagram under local coordinate system; Figure 4 A quadrilateral element integration point schematic diagram under different coordinate systems is established; wherein (a) is a quadrilateral element integration point schematic diagram under global coordinate system; (b) is a quadrilateral element integration point schematic diagram under local coordinate system. DETAILED DESCRIPTION

[0020] All the features disclosed by all the embodiments in the present specification, or all the steps of the methods or processes disclosed by the embodiments, can be combined and / or extended, replaced or substituted, except for those mutually exclusive features and / or steps, in any way, unless otherwise specified.

[0021] The specific implementation process of the present application is as follows: The embodiment scheme of the present application is particularly related to a fine prediction technology scheme of the mechanical properties of aircraft composite materials in a normal temperature environment. In a preferred embodiment, a method for predicting the mechanical properties of aircraft heat-resistant composite materials is specifically provided, comprising the following steps: The grid of the aircraft heat-resistant composite material is subjected to non-uniform grid division, and is divided into a local encryption area and a non-encryption area. The local encryption area is a fine grid corresponding to the microstructure feature position of the composite material; and the non-encryption area is a coarse grid corresponding to the non-microstructure feature position of the composite material. For the grid nodes at the interface of the grid division, the cell centers and edge midpoints are sequentially connected to form control volumes. The balance equation is discretized according to the formed control volumes, and midpoint integration is performed to calculate and solve the balance equation. Based on the solution of the balance equation, the displacements of all grid nodes are taken as variables to be solved, the system matrix is assembled, and the solution is obtained. Then the obtained displacement is brought into the stress-strain and displacement relationship to obtain the strain and stress at the corresponding position, thereby completing the prediction of the mechanical properties of the material.

[0022] In other embodiments, based on the above embodiment, the following more detailed implementation steps are further provided: Step 1: After grid division, the nodes at the interface position establish discrete control volumes. Specifically, taking the discrete area formed by quadrilateral and triangular grids as an example (see Figure 1 , the black points represent cell nodes, the hollow points represent cell centers and material property storage positions, the red lines represent the interface, the blue line area is the grid encryption area, the black grid is the non-grid encryption area, and the dashed line area is the control volume), for the nodes in the non-encryption area (CV1) or the non-grid transition area (CV3), the control volume construction method is the same as the traditional CV-FVM (Cell-Vertex Finite Volume Method) discrete method, and the construction method is shown in Figure 1 ; for the nodes at the interface, in the inventive concept, the traditional CV-FVM control volume (Cell-Vertex Finite Volume Method) construction method is referenced, and the cell centers and edge midpoints are sequentially connected to form a CV2 type control volume as shown in Figure 1 or Figure 2 .

[0023] Step 2: According to the control volume established in step 1, the balance equation discrete format is established. Specifically, the balance equation is discretized for the control volume formed in step 1, wherein the balance equation (without body force) for isotropic material is as follows (1):

[0024] wherein, p , t denote density, time, u i is displacement u component along i direction, n j is outward normal n component along j direction, μ and λ Lame coefficients, δ ij is Kronicker symbol, V denotes integration volume, S denotes integration surface (integration line for 2D problem).

[0025] The control volume CV2 formed in the present scheme is taken as an example (as shown in Figure 2 , the control equation (1) is discretized by CV-FVM. The triangular element is denoted by character T (Triangle), and the quadrilateral element is denoted by character Q (Quadrilateral). Specifically, the control volume CV2 is composed of integral line segments in the quadrilateral elements Q1 (node 1-node 2-node 4-node 5), Q2 (node 6-node 7-node 8-node 10), Q3 (node 4-node 9-node 8-node 10), and the triangular elements T1 (node 4-node 5-node 6), T2 (node 2-node 3-node 9), T3 (node 3-node 9-node 4) (1-8), and the integral items on the right side of equation (1) are approximately calculated based on the midpoint integral formula, the integral position is shown in l 1- l 8, and for the two-dimensional problem, equation (1) is written as: Figure 2

[0026] For the control volume CV2, equation (2) is converted to:

[0027] wherein, the upper index "midpoint" represents the midpoint of the integral line (1-8), l i , i i l i represents the length of the integral line (1-8), i l i and k 单元i and n x and n y ​​​​are constants and are stored directly before numerical calculation, u and v respectively represent x and y the velocity in the direction of μ i and λ i represent the Lame coefficients at the i th integration point.

[0028] For the calculation of equation (2) and equation (3), the spatial derivatives of the displacement in the global coordinate system at the integration point are needed , . , which are calculated by mapping them to a standard quadrilateral element, a triangular element (see Figures 3-4 ).

[0029] While in the local coordinate system, the spatial derivatives of the shape functions can be calculated using the following equation:

[0030] where J is the Jacobian matrix.

[0031] Further, we have:

[0032] The Jacobian matrix J is expressed based on the numerical basis functions:

[0033] where N / ξ and N / η The expressions of

[0034] In the above equation, the local coordinates (ξ, η) of the integration point for the triangular element are:

[0035] For the quadrilateral element, the local coordinates (ξ, η) of the integration point are:

[0036] Put formula (4) - formula (9) into formula (2) - formula (3), both can obtain the expression of the integral item on the right side of formula (1), which is the first equation expression of the node displacement and constant product.

[0037] The time integral item of formula (2) is approximated by using the midpoint integral formula:

[0038] Wherein, the upper index t Indicates the current time, t -△ t Indicates the last time, V i Indicates the volume (2-dimensional problem is area) of the control body in the first i Unit, t -2△ t Indicates the time before the last time.

[0039] Step 3: based on the displacement on all nodes in step 2 as a to-be-solved variable, the system matrix is assembled and solved, and the displacement obtained by the solution is brought into the stress, strain and displacement relationship formula, and then the strain and stress at the position of interest are obtained. It should be noted that the stress, strain and displacement relationship formula can be completed by using the existing calculation formula, and will not be repeated here.

[0040] It should be noted that the above scheme of the present application introduces local grid encryption technology based on the traditional unencrypted grid grid type finite volume method (CV-FVM), which can ensure the realization of non-uniform mesh division of the encryption region and the non-encryption region of the composite material under the basic displacement coordination condition, avoid the use of global refined grid solving method when researching the macroscopic mechanical properties of the microstructure, greatly reduce the number of solving grids, and improve the calculation precision without additional calculation time. And the method of the present application can be easily extended to three-dimensional tetrahedral element and hexahedral mixed grid element mechanical problems, so that it is suitable for the research of three-dimensional composite mechanical properties.

[0041] The elements described in the embodiments of the present application can be realized by software or hardware, and the described elements can also be arranged in a processor. In some cases, the names of these elements do not constitute a limitation on the elements themselves.

[0042] According to an aspect of an embodiment of the present application, a computer program product or computer program is provided, which includes computer instructions stored in a computer readable storage medium. The processor of the computer device reads the computer instructions from the computer readable storage medium, and the processor executes the computer instructions, so that the computer device executes the method provided in the various optional implementation manners described above.

[0043] As another aspect, the embodiments of the present application also provide a computer readable medium, which can be included in the electronic device described in the above embodiments, or can exist separately without being assembled into the electronic device. The computer readable medium carries one or more programs, which, when executed by the electronic device, enable the electronic device to implement the method described in the above embodiments.

Claims

1. A method for predicting the mechanical properties of thermally insulating composite materials for aircraft, characterized in that, Includes the following steps: The heat-resistant composite material mesh for aircraft is divided into non-uniform meshes. The divided regions include regions that require local mesh refinement and regions that do not require mesh refinement. Regions that require local mesh refinement are fine meshes, corresponding to the microstructural features of the composite material. Regions that do not require mesh refinement are coarse meshes, corresponding to the nonstructural features of the composite material. For the grid nodes at the grid interface, connect the center of the grid cell and the midpoint of the edge in sequence to form a control volume; The solution is obtained by calculating the discrete equilibrium equations of the control volume based on the midpoint integral. The displacements on all grid nodes are treated as variables to be solved, the system matrix is ​​assembled, and then solved. The obtained displacement is then substituted into the stress-strain-displacement relationship to obtain the strain and stress at the corresponding position, thereby completing the prediction of the material's mechanical properties.

2. The method for predicting the mechanical properties of heat-resistant composite materials for aircraft according to claim 1, characterized in that, The heat-resistant composite material mesh for the aircraft comprises a mesh of polygonal units.

3. The method for predicting the mechanical properties of heat-resistant composite materials for aircraft according to claim 1, characterized in that, The microstructure of the composite material includes fibers, particles, and air inclusions.

4. The method for predicting the mechanical properties of heat-resistant composite materials for aircraft according to claim 1, characterized in that, The control volume is formed by sequentially connecting the center of the grid cell and the midpoint of the edge. Specifically, the control volume is constructed by sequentially connecting the center of the grid cell and the midpoint of the edge using the CV-FVM control volume construction method.

5. The method for predicting the mechanical properties of heat-resistant composite materials for aircraft according to claim 2, characterized in that, The mesh composed of polygonal units includes a mixed mesh of quadrilateral and triangular units.

6. The method for predicting the mechanical properties of heat-resistant composite materials for aircraft according to claim 2, characterized in that, The mesh composed of polygonal units includes a mixture of three-dimensional tetrahedral units and hexahedral units.

7. A device for predicting the mechanical properties of thermally insulating composite materials for aircraft, characterized in that, It includes a processor and a memory, wherein the memory stores a computer program that, when loaded by the processor, executes the method as described in any one of claims 1 to 6.

8. A computer-readable storage medium, characterized in that, A computer program is stored in a readable storage medium, which, when loaded by a processor, executes the method as described in any one of claims 1 to 6.

9. A system for predicting the mechanical properties of thermally insulating composite materials for aircraft, characterized in that, Includes the device for predicting the mechanical properties of thermally protective composite materials for aircraft as described in claim 7.

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