A computational method for pyrolysis and ablation simulation in the analysis of uncontrolled falling and reentry of spacecraft

Abstract

The present invention discloses a method for simulating and calculating pyrolysis and ablation in the analysis of the uncontrolled fall and re-entry disintegration of a spacecraft, comprising: step one, in the analysis of the uncontrolled fall and re-entry disintegration of a spacecraft, establishing a fast algorithm for re-entry aerodynamic thermal ablation and internal temperature distribution that matches carbon-based composite materials; step two, solving the heat conduction inside the spacecraft and its disintegration shape; step three, judging the failure order of the material based on the heat conduction solution structure in step two, and then obtaining a calculation method for the pyrolysis / ablation of the carbon-based composite material.

Description

Technical Field

[0001] The present invention relates to the field of satellites and application products. More specifically, the present invention relates to a pyrolysis and ablation simulation calculation method used in the analysis of uncontrolled fall and re-entry disintegration of spacecraft used in the prediction and simulation of the uncontrolled spacecraft re-entry and disintegration process by a spacecraft measurement and control ground station. Background Art

[0002] The analysis of the re-entry disintegration of complex structure spacecraft requires first clarifying the criteria for material damage, that is, under what circumstances the material is destroyed. For metal materials, this patent adopts a melting point temperature control model, assuming that the surface temperature of the metal material reaches the melting point, the material surface melts, and flows under the action of aerodynamic shear force; for carbon-based composite materials, a pyrolysis / ablation control model is adopted, assuming that the structure is completely pyrolyzed or ablated and damaged. Aerodynamic heat / softening / melting or aerodynamic heat / pyrolysis / ablation calculations and analysis are performed on the spacecraft along an uncontrolled trajectory with a small re-entry angle, and the ablation shape is solved by applying the spring sheet stretching method deformation non-structural dynamic grid technology. The aerodynamic heat calculation uses the outer boundary nodes of the tetrahedral grid. The grid unit and node number information are kept unchanged in the ablation shape solution, and only the node coordinates are changed to realize the data information transmission of different calculation modules of the non-structured grid. The dynamic grid of the ablation shape, three-dimensional finite element heat transfer and aerodynamic heat are coupled and analyzed to give a theoretical prediction of the disintegration of the whole device and its components.

[0003] The existing simulation method steps are mainly divided into three steps:

[0004] Step 1: In the analysis of uncontrolled fall and re-entry disintegration of spacecraft, a fast algorithm for re-entry aerodynamic thermal ablation and internal temperature distribution is established in conjunction with carbon-based composite materials;

[0005] Step 2, solve the heat conduction inside the spacecraft and its disintegrated shape;

[0006] Step three, based on the heat conduction solution structure in step three, determine the failure order of the material, and then obtain the calculation method for pyrolysis / ablation of carbon-based composite materials. The defect is that the calculation amount is too large and the calculation time is too long when solving the three-dimensional temperature field process. Summary of the invention

[0007] An object of the present invention is to solve at least the above problems and / or disadvantages and to provide at least the advantages which will be described hereinafter.

[0008] In order to achieve these purposes and other advantages according to the present invention, a pyrolysis and ablation simulation calculation method in the analysis of uncontrolled fall and re-entry disintegration of a spacecraft is provided, comprising:

[0009] Step 1: In the analysis of uncontrolled fall and re-entry disintegration of spacecraft, a fast algorithm for re-entry aerodynamic thermal ablation and internal temperature distribution is established in conjunction with carbon-based composite materials;

[0010] Step 2, solve the heat conduction inside the spacecraft and its disintegrated shape;

[0011] Step three, based on the heat conduction solution structure in step two, determine the failure order of the material, and then obtain the calculation method of pyrolysis / ablation of carbon-based composite materials.

[0012] Preferably, in step 1, the reentry aerodynamic thermal ablation and internal temperature distribution fast algorithm is configured to include:

[0013] S10, establish a finite element algorithm suitable for the transient temperature field of the aerodynamic thermal environment structure response, and then form a preliminary spacecraft disintegration criterion model;

[0014] S11, for carbon-based composite materials, a pyrolysis-ablation control model is adopted. It is assumed that the structure is completely pyrolyzed or ablated. The established spacecraft disintegration criterion model is used to couple the aerodynamic heat and structural thermal responses to solve them.

[0015] Preferably, in S10, the establishment of the finite element algorithm is configured to include:

[0016] S101, based on the three-dimensional transient heat conduction equation, the tetrahedron four-node finite element method is used to solve the temperature field of the spacecraft surface structure, and the finite element method structure thermal response equilibrium equation is derived using functional variation;

[0017] S102, based on the overall synthesis of the finite element method, the stiffness matrix has symmetric positive definiteness, high sparseness, and regular distribution of non-zero elements, develops one-dimensional variable bandwidth storage technology, and constructs a numerical solution strategy for the three-dimensional finite element heat transfer calculation model;

[0018] S103, based on the method of concentrated variable temperature matrix coefficients, a two-point backward difference format, Crank-Nicolson format and Galerkin format are constructed to solve the three-dimensional transient temperature field, and then a finite element algorithm suitable for the transient temperature field of the structure response to the aerodynamic thermal environment is obtained.

[0019] Preferably, in step 2, according to the ablation condition obtained in step 1, based on the three-dimensional heat conduction equation, a tetrahedron four-node unit is used to solve the three-dimensional temperature field problem, thereby simulating the heat conduction condition during the spacecraft reentry process;

[0020] In the process of solving the three-dimensional temperature field, the Kriging spatial interpolation algorithm is used to obtain global three-dimensional data using partial three-dimensional data.

[0021] Preferably, in step three, the failure order of the materials is:

[0022] For metal materials, the melting point temperature control model is adopted, assuming that the surface temperature of the metal material reaches the melting point and is lost under the action of aerodynamic shear force;

[0023] For carbon-based composites, a pyrolysis / ablation control model is used, assuming that the structure is completely pyrolyzed or ablated.

[0024] The present invention includes at least the following beneficial effects: In the simulation solution process of the present invention, this step introduces the Kriging spatial interpolation algorithm, which, as a statistical method for spatial interpolation, can use partial three-dimensional data to obtain global three-dimensional data, thereby reducing the amount of calculation in the process of solving the three-dimensional temperature field and reducing the calculation time.

[0025] Other advantages, objectives and features of the present invention will be embodied in part through the following description, and in part will be understood by those skilled in the art through study and practice of the present invention. BRIEF DESCRIPTION OF THE DRAWINGS

[0026] Figure 1 It is a curve diagram of the gas cylinder reentry trajectory;

[0027] Figure 2 This is the ablation and temperature distribution diagram of the gas cylinder at an altitude of 85km;

[0028] Figure 3 This is the ablation and temperature distribution diagram of the gas cylinder at an altitude of 83.3km. DETAILED DESCRIPTION

[0029] The present invention is described in further detail below so that those skilled in the art can implement it according to the description.

[0030] According to the present invention, a method for simulating the pyrolysis and ablation of a spacecraft during uncontrolled fall and reentry is implemented, which includes:

[0031] Step 1: In the analysis of uncontrolled fall and re-entry disintegration of spacecraft, a fast algorithm for re-entry aerodynamic thermal ablation and internal temperature distribution is established in conjunction with carbon-based composite materials;

[0032] Step 2, solve the heat conduction inside the spacecraft and its disintegrated shape;

[0033] Step three, based on the heat conduction solution structure in step two, determine the failure order of the material, and then obtain the calculation method of pyrolysis / ablation of carbon-based composite materials.

[0034] In the carbon-based composite material reentry aerodynamic thermal ablation and internal temperature distribution fast algorithm of step 1, the reentry aerodynamic thermal ablation and internal temperature distribution fast algorithm is configured to include:

[0035] This step mainly simulates the internal temperature changes of carbon-based composite materials due to aerodynamic effects during the re-entry of the spacecraft, and then simulates the thermal ablation of the materials.

[0036] Starting from the three-dimensional transient heat conduction equation, the tetrahedron four-node finite element method is proposed to solve the temperature field of the spacecraft surface structure, and the finite element method structure thermal response equilibrium equation is derived using functional variation. The numerical solution strategy of the three-dimensional finite element heat transfer calculation model is constructed, and the overall synthesis of the finite element method is used to make the stiffness matrix symmetrical, positive, highly sparse, and regular with non-zero element distribution. The one-dimensional variable bandwidth storage technology is developed to solve the data storage problem of large sparse matrices; the method of concentrating the variable temperature matrix coefficient is applied to effectively suppress the temperature oscillation phenomenon in the solution process, and the two-point backward difference format, Crank-Nicolson format and Galerkin format are constructed to solve the three-dimensional transient temperature field, and a set of finite element algorithms suitable for the transient temperature field of the aerodynamic thermal environment structure response is established. A preliminary model of the complex structure spacecraft disintegration criterion is formed. For carbon-based composite materials, a pyrolysis / ablation control model is adopted, assuming that the structure is completely pyrolyzed or ablated, that is, damaged. Applying the established model, the aerodynamic heat and structural thermal response are coupled and solved, and the transient aerodynamic thermal environment and material disintegration damage of the Tianzhou-1 cargo spacecraft and the two cabins with thin shell structures that re-enter at a small re-entry angle at a distance of 120km to 90km are calculated and analyzed. The results are in good agreement with the calculation results of the finite element software, and the reliability of the heat transfer model is verified.

[0037] Research shows that the ablation process of carbon-based materials can be described by a layered model. The part of the material close to the heating surface is the carbonized layer, which is a porous material formed after the original material is completely pyrolyzed. The pyrolysis zone where the pyrolysis reaction is taking place is between the carbonized layer and the original material layer. The pyrolysis reaction absorbs heat, and the pyrolysis gas produced by the reaction flows in the carbonized layer, blocking the heat transfer. Below the pyrolysis zone is the original material that has not yet been pyrolyzed.

[0038] In another example, in step 2, according to the ablation situation obtained in step 1, the three-dimensional heat conduction equation is used to solve the three-dimensional temperature field problem using tetrahedral four-node units to simulate the heat conduction situation during the spacecraft re-entry process.

[0039] In the solution process, this step introduces the Kriging spatial interpolation algorithm, which, as a statistical method for spatial interpolation, can use partial three-dimensional data to obtain global three-dimensional data, thereby reducing the amount of calculation in the process of solving the three-dimensional temperature field and reducing the calculation time.

[0040] In order to verify the ablation shape solution technology of complex spacecraft structures and the moving boundary aerodynamic heat / ablation and internal heat conduction coupling solution technology, the gas cylinder component is taken as an example to calculate the ablation and surface temperature distribution of the typical re-entry height of the gas cylinder with multiple disintegration components at an angle of attack of 20°. In order to observe the effect of grid deformation, the material of the gas cylinder component is set to an easily ablative aluminum alloy material. Figure 1 is the reentry trajectory curve of the gas cylinder, the initial trajectory height is 90km, the initial reentry speed is 7.48km / s, and the flight attack angle of the gas cylinder is 20°. Figure 2-3 The calculation results of the ablation and temperature distribution of the cylinder along the trajectory show that the temperature of the cylinder on the windward side is higher and the ablation is more obvious. In the case of flight with an angle of attack, the aircraft will be ablated into an asymmetric shape, and the disintegration will first occur in the windward end area.

[0041] Through the above tests, the patented method with the Kriging spatial interpolation algorithm and the original method obtain results with an error within 1%. However, due to the introduction of the Kriging spatial interpolation algorithm, step two of the patented method reduces the amount of calculation by 80% and shortens the calculation time by nearly 30%. Therefore, compared with the original method, the patented method has obvious advantages in computational efficiency.

[0042] In the calculation method of carbon-based composite material pyrolysis / ablation in step 3, the re-entry damage analysis of complex structure spacecraft needs to first clarify the criterion of material damage, that is, under what circumstances the material fails. This step determines the failure order of the material based on the heat conduction simulation results of the previous step.

[0043] For metal (alloy) materials, the melting point temperature control model is adopted, assuming that the surface temperature of the metal material reaches the melting point, the material surface will melt, and it will flow under the action of aerodynamic shear force; for carbon-based composite materials, the pyrolysis / ablation control model is adopted, assuming that the structure is completely pyrolyzed or ablated and destroyed. Based on the surface aerodynamic thermal data of low-orbit spacecraft at an altitude of 120km to 90km at the end of its service life, the aerodynamic heat / melting or aerodynamic heat / pyrolysis / ablation calculation and analysis of the thin shell structure of the low-orbit spacecraft and its components along the uncontrolled trajectory with a small re-entry angle are performed, and the theoretical prediction of its disintegration is given, providing data support for the analysis of the falling point of the wreckage.

[0044] For metal materials, this patent adopts a melting point temperature control model, assuming that the surface temperature of the metal material reaches the melting point, the material surface will melt, and will flow under the action of aerodynamic shear force; for carbon-based composite materials, a pyrolysis / ablation control model is adopted, assuming that the structure is completely pyrolyzed or ablated and destroyed. Aerodynamic heat / softening / melting or aerodynamic heat / pyrolysis / ablation calculation and analysis are performed on the spacecraft along an uncontrolled trajectory with a small reentry angle, and the ablation shape is solved by applying the spring sheet stretching method deformation non-structural dynamic grid technology. The aerodynamic heat calculation uses the outer boundary nodes of the tetrahedral grid. The grid unit and node number information are kept unchanged in the ablation shape solution, and only the node coordinates are changed to realize the data information transmission of different calculation modules of the unstructured grid. The dynamic grid of the ablation shape, three-dimensional finite element heat transfer and aerodynamic heat are coupled and analyzed to give a theoretical prediction of the disintegration of the whole device and its components.

[0045] The above solution is only an illustration of a preferred embodiment, but is not limited thereto. When implementing the present invention, appropriate replacement and / or modification can be performed according to user needs.

[0046] The number of devices and processing scales described here are used to simplify the description of the present invention. Applications, modifications and variations of the present invention will be obvious to those skilled in the art.

[0047] Although the embodiments of the present invention have been disclosed above, they are not limited to the applications listed in the description and the embodiments. It can be fully applied to various fields suitable for the present invention. For those familiar with the art, additional modifications can be easily realized. Therefore, without departing from the general concept defined by the claims and equivalent scope, the present invention is not limited to specific details.

Claims

1. A pyrolysis and ablation simulation calculation method for the analysis of uncontrolled fall and reentry of a spacecraft, characterized in that: include: Step 1: In the analysis of uncontrolled fall and re-entry disintegration of spacecraft, a fast algorithm for re-entry aerodynamic thermal ablation and internal temperature distribution is established in conjunction with carbon-based composite materials; Step 2, solve the heat conduction inside the spacecraft and its disintegrated shape; Step 3, according to the heat conduction solution structure in step 2, determine the failure order of the material, and then obtain the calculation method of pyrolysis / ablation of carbon-based composite materials; Wherein, in step 2, according to the ablation situation obtained in step 1, based on the three-dimensional heat conduction equation, a tetrahedron four-node unit is used to solve the three-dimensional temperature field problem, thereby simulating the heat conduction situation during the spacecraft reentry process; In the process of solving the three-dimensional temperature field, the Kriging spatial interpolation algorithm is used to obtain the global three-dimensional data using partial three-dimensional data; In step 1, the reentry aerodynamic thermal ablation and internal temperature distribution fast algorithm is configured to include: S10, establish a finite element algorithm suitable for the transient temperature field of the aerodynamic thermal environment structure response, and then form a preliminary spacecraft disintegration criterion model; S11, for carbon-based composite materials, a pyrolysis-ablation control model is used. It is assumed that the structure is completely pyrolyzed or ablated. The established spacecraft disintegration criterion model is used to couple the aerodynamic heat and structural thermal responses. In S10, the establishment of the finite element algorithm is configured to include: S101, based on the three-dimensional transient heat conduction equation, the tetrahedron four-node finite element method is used to solve the temperature field of the spacecraft surface structure, and the finite element method structure thermal response equilibrium equation is derived using functional variation; S102, based on the overall synthesis of the finite element method, the stiffness matrix has symmetric positive definiteness, high sparseness, and regular distribution of non-zero elements, develops one-dimensional variable bandwidth storage technology, and constructs a numerical solution strategy for the three-dimensional finite element heat transfer calculation model; S103, based on the method of concentrated variable temperature matrix coefficients, constructs two-point backward difference format, Crank-Nicolson format and Galerkin format for solving three-dimensional transient temperature field, and then obtains the finite element algorithm suitable for transient temperature field of structure response in aerodynamic thermal environment; In step three, the failure order of the materials is: For metal materials, the melting point temperature control model is adopted, assuming that the surface temperature of the metal material reaches the melting point and is lost under the action of aerodynamic shear force; For carbon-based composites, a pyrolysis / ablation control model is used, assuming that the structure is completely pyrolyzed or ablated.

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